Citations w/abstracts NASA TP[793]

NASA/TP-1999-206568 Fifty Years of Flight Research: An Annotated Bibliography of Technical Publications of NASA Dryden Flight Research Center, 1946–1996 David F. Fisher NASA Dryden Flight Research Center Edwards, California Cnβ = Clβ /C n δa /C l δ a Cn * = C β nβ – α Iz 0 Ix Clβ Whitcomb area rule F.S. 70 F.S. 85 F.S. 107 Nose ring pressure orifices 180° 270° F.S. 184 F.S. 142 Yawing moment 90° 0° F.S. 142 Vortical flow structure concept May 1999 The NASA STI Program Office… in Profile Since its founding, NASA has been dedicated to the advancement of aeronautics and space science. The NASA Scientific and Technical Information (STI) Program Office plays a key part in helping NASA maintain this important role. The NASA STI Program Office is operated by Langley Research Center, the lead center for NASA’s scientific and technical information. The NASA STI Program Office provides access to the NASA STI Database, the largest collection of aeronautical and space science STI in the world. The Program Office is also NASA’s institutional mechanism for disseminating the results of its research and development activities. These results are published by NASA in the NASA STI Report Series, which includes the following report types: • TECHNICAL PUBLICATION. Reports of completed research or a major significant phase of research that present the results of NASA programs and include extensive data or theoretical analysis. Includes compilations of significant scientific and technical data and information deemed to be of continuing reference value. NASA’s counterpart of peer-reviewed formal professional papers but has less stringent limitations on manuscript length and extent of graphic presentations. TECHNICAL MEMORANDUM. Scientific and technical findings that are preliminary or of specialized interest, e.g., quick release reports, working papers, and bibliographies that contain minimal annotation. Does not contain extensive analysis. CONTRACTOR REPORT. Scientific and technical findings by NASA-sponsored contractors and grantees. • CONFERENCE PUBLICATION. Collected papers from scientific and technical conferences, symposia, seminars, or other meetings sponsored or cosponsored by NASA. SPECIAL PUBLICATION. Scientific, technical, or historical information from NASA programs, projects, and mission, often concerned with subjects having substantial public interest. TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and technical material pertinent to NASA’s mission. • • Specialized services that complement the STI Program Office’s diverse offerings include creating custom thesauri, building customized databases, organizing and publishing research results…even providing videos. For more information about the NASA STI Program Office, see the following: • Access the NASA STI Program Home Page at http://www.sti.nasa.gov E-mail your question via the Internet to help@sti.nasa.gov Fax your question to the NASA Access Help Desk at (301) 621-0134 Telephone the NASA Access Help Desk at (301) 621-0390 • • • • • • Write to: NASA Access Help Desk NASA Center for AeroSpace Information 7121 Standard Drive Hanover, MD 21076-1320 NASA/TP-1999-206568 Fifty Years of Flight Research: An Annotated Bibliography of Technical Publications of NASA Dryden Flight Research Center, 1946–1996 David F. Fisher NASA Dryden Flight Research Center Edwards, California National Aeronautics and Space Administration Dryden Flight Research Center Edwards, California 93523-0273 May 1999 ACKNOWLEDGMENTS The author acknowledges Dryden Librarian Dennis Ragsdale for his help in finding many of the entries in this bibliography and Library Technicians Charito Lopez and Lisa Carbaugh for their help scanning and typing many of the missing abstracts. Also, the author thanks the Dryden Graphics Office; Steve Lighthill, chief, Justine Mack, James Zeitz and Dennis Calaba for the many airplane three-view drawings and cover. The author is especially grateful to the Dryden Technical Publications Office; Camilla McArthur, chief, editorial assistants Lois Williams, Mary Whelan, and Karen Wick, editors Sue Luke and Muriel Khachooni for their patience and persistence in seeing this bibliography through to publication. NOTICE Use of trade names or names of manufacturers in this document does not constitute an official endorsement of such products or manufacturers, either expressed or implied, by the National Aeronautics and Space Administration. Available from the following: NASA Center for AeroSpace Information (CASI) 7121 Standard Drive Hanover, MD 21076-1320 (301) 621-0390 National Technical Information Service (NTIS) 5285 Port Royal Road Springfield, VA 22161-2171 (703) 487-4650 CONTENTS Page FOREWORD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . v ABSTRACT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 ACRONYMS AND ABBREVIATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11 ORDERING INFORMATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 KEY TO CITATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15 BIBLIOGRAPHY. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16 1947 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16 1948 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16 1949 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18 1950 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 1951 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25 1952 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27 1953 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29 1954 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32 1955 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34 1956 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38 1957 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43 1958 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47 1959 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51 1960 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56 1961 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59 1962 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64 1963 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68 1964 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71 1965 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74 1966 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78 1967 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83 1968 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88 1969 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 94 1970 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 1971 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109 iii 1972 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116 1973 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124 1974 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133 1975 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 141 1976 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 148 1977 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 157 1978 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 166 1979 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179 1980 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 186 1981 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194 1982 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 1983 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 1984 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 1985 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 245 1986 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 258 1987 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 276 1988 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 289 1989 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 307 1990 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 322 1991 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 339 1992 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 356 1993 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 377 1994 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 393 1995 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 406 1996 Technical Publications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 418 APPENDIX A. NASA CONTRACTOR REPORTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 430 APPENDIX B. UCLA FLIGHT SYSTEMS RESEARCH CENTER PUBLICATIONS . . . . . . . . . . 473 APPENDIX C. VIDEOTAPES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 494 AUTHOR INDEX . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 496 AIRPLANE INDEX . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 513 iv FOREWORD The legacy of a research organization is the written history of its findings, whether the results were expected or unexpected, whether the predictions were valid or not. The search for understanding and the advancement of ideas to technology is often a non-linear path. Ideas beget ideas, and data changes the reference point. Without written evidence of this progress, the map ends. Future explorers may have to retrace and relearn at great time and cost. I am pleased to have the bibliography of technical publications of the NASA Dryden Flight Research Center published for future researchers to have. The first entries in this document deal with transonic and supersonic flight research at the dawn of the high speed aircraft era at Muroc. The latter entries deal with tailless flight, hypersonic flight, and civil transport safety. The technical publications of the Dryden Flight Research Center are a reflection of the progress made in NASA’s aeronautics research and technology program. They show what teams of people believed could fly or be flown. They document the claims, hopes, and aspirations of designers from across the country. These reports document the real and the imagined, the overlooked, and the unexpected, as Dr. Hugh L. Dryden would say. I hereby dedicate this annotated bibliography of technical publications of the Dryden Flight Research Center to the memory of Dr. Hugh L. Dryden, scientist, engineer, manager, father, and technical author. His own research provided some of the foundation of the transonic flight exploration which spawned what is now the Dryden Flight Research Center. I acknowledge the outstanding work of Mr. David Fisher, a research engineer at Dryden, to produce this bibliography. Kenneth J. Szalai NASA Dryden Flight Research Center Director 1994–1998 v Fifty Years of Flight Research: An Annotated Bibliography of Technical Publications of NASA Dryden Flight Research Center, 1946–1996 David F. Fisher NASA Dryden Flight Research Center Edwards, California 93523-0273 ABSTRACT Titles, authors, report numbers, and abstracts are given for more than 2200 unclassified and unrestricted technical reports and papers published from September 1946 to December 1996 by NASA Dryden Flight Research Center and its predecessor organizations. These technical reports and papers describe and give the results of 50 years of flight research performed by the NACA and NASA, from the X-1 and other early X-airplanes, to the X-15, Space Shuttle, X-29 Forward Swept Wing, and X-31 aircraft. Some of the other research airplanes tested were the D-558, phase 1 and 2; M-2, HL-10 and X-24 lifting bodies; Digital Fly-By-Wire and Supercritical Wing F-8; XB-70; YF-12; AFTI F-111 TACT and MAW; F-15 HiDEC; F-18 High Alpha Research Vehicle, and F-18 Systems Research Aircraft. The citations of reports and papers are listed in chronological order, with author and aircraft indices. In addition, in the appendices, citations of 233 contractor reports, more than 200 UCLA Flight System Research Center reports and 25 video tapes are included. INTRODUCTION In September of 1946, a small band of engineers and technicians came to Muroc Army Air Field in southern California from the National Advisory Committee for Aeronautics (NACA) Langley Memorial Aeronautical Laboratory at Hampton, Virginia. These people were to assist in a supersonic flight research program involving the Bell XS-1 aircraft. The following year, the small group, which became known as the NACA Muroc Flight Test Unit, were key participants in the first known supersonic flight of an airplane on October 14, 1947. In 1949, they became the NACA High-Speed Flight Research Station (HSFRS), a division of the Langley laboratory and during 1950, they published 19 technical reports on various aspects of pioneering flight research. In 1954, the HSFRS became the NACA High-Speed Flight Station and moved from the south base location shared with the Air Force to the present location north of the base flight line. In 1959, after the creation of the National Aeronautic and Space Administration (NASA), the center was designated NASA Flight Research Center (FRC). On March 26, 1976, the center was redesignated Hugh L. Dryden Flight Research Center, in honor of the American aerospace pioneer, former Director of NACA and the first Deputy Director of NASA. In 1981, the center became a facility as part of Ames Research Center. In 1994, the facility became an independent NASA Center again as the Hugh L. Dryden Flight Research Center. This document attempts to capture all the unrestricted reports, papers, and journal articles published by authors while they were employed by the NASA Dryden Flight Research Center and its predecessor organizations from September 1946 to December 1996. Also included are NASA Contractor Reports that were sponsored by DFRC. Reports from the UCLA Flight Systems Research Center under NASA Dryden Grant NCC2-374 are included in Appendix B. Many of the citations are from NASA CASI RECON and RECONplus databases, the NASA Dryden card catalog, as well as information from the authors that was not included in the databases. The author regrets any documents that may have been inadvertently left out. Some of the highlights of the first 50 years of what is now known as the Hugh L. Dryden Flight Research Center are as follows: Sept. 30, 1946 Five NACA engineers, headed by Walter Williams, arrive at Muroc Army Air Base, California (now Edwards AFB) by this date from Langley Memorial Aeronautical Laboratory (now NASA Langley Research Center, Virginia), to prepare for X-1 supersonic research flights in a joint NACA-Army Air Forces program. First NACANASA presence is established at the Mojave Desert site. (Note: Some sources report the arrival of 13 individuals, but an early chronology shows that 13 were not present at Muroc until December.) Bell pilot Chalmers Goodlin flies the first successful rocket-powered flight of the X-1 (then designated XS-1). D-558-1 sets a world speed record of 640.7 mph. The NACA Muroc Flight Test Unit receives permanent status from Hugh L. Dryden, the NACA Director of Research. The staff now includes 27 people with Walt Williams as Head. X-1 exceeds the speed of sound in history’s first supersonic flight flown by then Maj. Charles E. Yeager, attaining maximum Mach number of 1.057. Howard C. Lilly is first NACA pilot to fly the jet-powered D-558-1 Skystreak. Herb Hoover becomes first NACA pilot and first civilian to fly supersonic in the X-1. The NACA Muroc unit, with about 100 people, is designated NACA High-Speed Flight Research Station (HSFRS), with Walt Williams remaining as director. John Griffith is first NACA pilot to fly the X-4 aircraft, studying flying qualities of tailless vehicles. Joe Walker is first to fly the variable-swept-wing X-5 to a full 60-degree angle. This concept is used today on F-14, F-111, and B-1 aircraft. The NACA first flight of XF-92A, a delta-wing aircraft used to study the problem of pitching up during maneuvering caused by the delta configuration. Last flight of XF-92A by the NACA. The flight research with this aircraft, the D-588-2 and the X-15, showed the desirability of low horizontal tail surface. That low horizontal tail configuration was later used on such supersonic swept-wing fighters as the F-100 Super Saber and F8U Crusader. The NACA pilot Scott Crossfield, in a rocket-powered D-558-2 Skyrocket, is the first to fly at twice the speed of sound, Mach 2.005. The NACA personnel move from the old south base site to new facilities that make up the original core of today’s Dryden complex. The cost to build the new facilities was $3.8 million. Personnel on this date number more than 200. 2 Dec. 9, 1946 Aug. 20, 1947 Sept. 7, 1947 Oct. 14, 1947 Nov. 25, 1947 Mar. 10, 1948 Nov. 14, 1949 Sept. 25, 1950 Mar. 4, 1952 Apr. 9, 1953 Oct. 14, 1953 Nov. 20, 1953 Jun. 26, 1954 Jul. 1, 1954 Aug. 23, 1954 Aug. 27, 1956 Sept. 27, 1956 The NACA HSFRS is redesignated the NACA High-Speed Flight Station. Joe Walker makes the first of 20 NACA research flights in the X-3 “Flying Stiletto” supersonic program. The NACA research pilot Joe Walker makes the first flight by NACA of an F-104A aircraft (the seventh F-104 aircraft off the assembly line). Air Force Capt. Milburn G. Apt flies the X-2 to Mach 3.2 in the first flight of an aircraft beyond 3 times the speed of sound. Unfortunately, he subsequently loses control of the airplane due to inertial coupling, and it crashes, killing him and destroying the vehicle. The NACA never flew the X-2 but did assist the program with advice and data analysis. National Advisory Committee for Aeronautics (NACA) becomes National Aeronautics and Space Administration (NASA). First of three X-15 rocket research aircraft arrive at NASA High-Speed Flight Station as preparations move ahead for the highly successful NASA-Air Force-Navy program that lasts 10 years to investigate hypersonic flight. John McKay makes last flight in the X-1E, final model flown of the X-1 series. This aircraft is now displayed in front of Dryden headquarters building. First unpowered glide flight of the X-15 is flown, with Scott Crossfield at the controls, is made from NASA’s “003,” B-52 launch aircraft. Paul F. Bikle succeeds Walt Williams as director of NASA High-Speed Flight Station. NASA High-Speed Flight Station at Edwards is redesignated NASA Flight Research Center. NASA personnel number about 340. First NASA X-15 aircraft flight is made, piloted by Joe Walker. First piloted flight above Mach 4 is made; Mach 4.43 is achieved by X-15 flown by USAF Capt. Robert M. White. First piloted flight above Mach 5 is made; Mach 5.27 is achieved by X-15 flown by USAF Capt. Robert M. White. First piloted flight above Mach 6 is made; Mach 6.04 is achieved by X-15 flown by USAF Capt. Robert M. White. Flight tests begin with the Paraglider Research Vehicle (Parasev). Developed to study ways of returning Gemini and Apollo spacecraft to Earth using a hang-glider-type wing. Pilot is Milt Thompson. M2-F1 lightweight lifting body is towed into the air over Rogers Dry Lake for the first time by a Pontiac convertible tow vehicle with Milt Thompson as pilot. Sets the stage for research with several lifting body designs to study atmospheric reentry of a vehicle like a Space Shuttle. Joe Walker flies X-15 to unofficial world altitude record of 354,200 ft. 3 Oct. 1, 1958 Oct. 15, 1958 Nov. 7, 1958 Jun. 8, 1959 Sept. 15, 1959 Sept. 27, 1959 Mar. 25, 1960 Mar. 7, 1961 Jun. 23, 1961 Nov. 9, 1961 Early 1962 Apr. 5, 1963 Aug. 22, 1963 Oct. 30, 1964 Joe Walker pilots the first flight of the Lunar Landing Research Vehicle (LLRV), “Flying Bedstead.” LLRV used to develop techniques of landing a spacecraft on the moon’s surface. High Temperature Loads Laboratory is formally accepted. With this facility, a complete YF-12 would be heated and loaded to simulate a high-speed flight. First flight of a heavyweight Milton O. Thompson. lifting body, the M2-F2, piloted by Mar. 11, 1966 Jul. 12, 1966 Apr. 25, 1967 First NASA flight of the XB-70A with Air Force Col. Joe Cotton and NASA research pilot Fitz Fulton at the controls. The XB-70 flights investigated the stability and handling qualities of large, delta-wing aircraft flying at high rates of speed. X-15 sets world’s absolute speed record for winged aircraft—4520 mph or Mach 6.72—with Air Force Maj. William Knight as the pilot. Last X-15 flight, 199th mission, is piloted by NASA pilot Bill Dana. World’s first hypersonic aircraft is most successful research aircraft to date. Last research flight of XB-70 is flown by Fitz Fulton and Air Force Lt. Col. Ted Sturmthal, reaching Mach 2.53. Program produced data on sonic booms, flight dynamics, and handling qualities associated with large supersonic aircraft. Flight is on 65th anniversary of Wright Brothers flight at Kitty Hawk, North Carolina. HL-10 becomes first lifting body to fly supersonic. John Manke, later to become Dryden site manager, is pilot. First NASA checkout flight of YF-12A, with Fitz Fulton as pilot. First flight of the M2-F3 lifting body is made by Bill Dana. NASA research pilots Tom McMurtry and Hugh Jackson reach a Dryden single-day record of six missions flown, in an F-104B while deployed to obtain data for the “Big Boom” experiments that sought to focus the energy from a sonic boom over a limited area. First flight of supercritical wing flown by NASA pilot Tom McMurtry. Unusual wing profile, tested on a modified F-8, increases flight efficiency and lowers fuel usage. This concept is now used widely on commercial and military aircraft. A Piper PA-30 Twin Commanche becomes testbed to develop remotely piloted aircraft techniques from a ground-based cockpit. This concept leads to successful projects such as three-eighths-scale F-15 spin research vehicle, HiMAT, and Boeing 720 jetliner purposely flown to a controlled crash landing in FAA test of anti-mist fuel additive. First flight of aircraft with all-electric fly-by-wire flight control system, the NASA F-8 Digital Fly-By-Wire research aircraft, with Gary Krier as pilot. This concept is now used in many aircraft, including Space Shuttles. The Boeing 747 shuttle carrier aircraft is used in wake vortex research program to study ways of reducing clear air turbulence trailing behind large aircraft. 4 Oct. 3, 1967 Oct. 24, 1968 Dec. 17, 1968 May 9, 1969 Mar. 5, 1970 Jun. 2, 1970 Oct. 14, 1970 Mar. 9, 1971 Oct. 14, 1971 May 25, 1972 Aug. 1974 Aug. 5, 1975 NASA pilot John Manke lands X-24B lifting body on Edwards runway, showing that a space shuttle-like vehicle can be landed safely on a designated runway after returning from orbit. NASA Flight Research Center is dedicated in honor of the late Hugh L. Dryden. NASA personnel number more than 560. First free-flight of the space shuttle Enterprise from the top of the Boeing 747 shuttle carrier aircraft. Enterprise piloted by Fred Haise and Gordon Fullerton. The 747 SCA was piloted by Fitz Fulton and Tom McMurtry, Vic Horton and Lewis (Skip) Guidry were the flight engineers. Last of 13 captive and free-flight tests with space shuttle prototype Enterprise, proving the shuttle glide and landing characteristics. Last research flight of the NASA YF-12 research program, with Fitz Fulton as pilot and Victor Horton flight test engineer on a YF-12A, one of three YF-12s flown during the program. Nearly 300 research flights explored high-speed, high-altitude flight, and yielded information on thermal stress, aerodynamics, high-altitude environment, and propulsion and flight control systems. AD-1 first flight is flown by NASA research pilot Thomas C. McMurtry. Three hundred and twenty thousand people at Edwards watch Columbia, the first orbital space shuttle, land. Dryden VIPs number 20,000, and 300,000 are at the East Shore public viewing site. Dryden is consolidated with Ames Research Center, Moffett Field, California, to become the Ames-Dryden Flight Research Facility. Position of Dryden director is renamed site manager and John Manke is selected for the post. NASA personnel number 491. President Ronald Reagan heads list of 45,000 guests at Dryden watching the fourth space shuttle landing. Crowd of 500,000 watches from East Shore public viewing site. A modified Schweizer SGS1-36 sailplane is flown in controlled, stabilized flight at 72 degrees angle of attack by research pilot Einar Enevoldson. NASA Dryden retires its oldest aircraft, the C-47 that towed the M2-F1 lifting body aloft during that program’s early days and was used to support many other projects. A remotely-piloted Boeing 720 test aircraft used in the joint FAA/NASA Controlled Impact Demonstration erupts in flames as it slides through the impact site on the dry lakebed, demonstrating that, contrary to expectations, an anti-misting fuel additive did not substantially inhibit fuel fires. Steve Ishmael is first NASA pilot to fly the X-29 research aircraft investigating forward-swept wings, composite construction concepts, and integrated flight controls. Data Analysis Facility opens as new home for general computer and associated engineering support and flight data operations. 5 Mar. 26, 1976 Aug. 12, 1977 Oct. 26, 1977 Oct. 31, 1979 Dec. 21, 1979 Apr. 14, 1981 Oct. 1, 1981 Jul. 4, 1982 Sept. 22, 1983 Oct. 30, 1984 Dec. 1, 1984 Apr. 2, 1985 Sept. 9, 1985 Jul. 10, 1986 F-111 Mission Adaptive Wing research aircraft flies Mach 1 for the first time, with Rogers Smith part of two-person crew. This program tested the wing with no ailerons, flaps, or slats. Camber changed mechanically in flight based on performance and mission. Groundbreaking held for $16.1 million Integrated Test Facility featuring interdependent systems testing, systems troubleshooting, and rapid pre-and postflight systems checkout on several aircraft simultaneously. Ed Schneider flies the 100th mission in the F-18 High Angle of Attack Research aircraft in Phase 1 of the three-phase program investigating the high angle-of-attack, “alpha,” regime. First self-repairing flight control system demonstrated on the F-15 HiDEC (Highly Integrated Digital Electronic Control) aircraft, with Jim Smolka as pilot. System identifies control surface failures or damage then automatically repositions other control surfaces to allow the pilot to continue the mission or land safely. First of three SR-71s arrive at Dryden for a program to investigate a host of disciplines to help development of future high-speed civil and military aircraft. (Three YF-12s, prototypes for a fighter-interceptor version of the SR-71, were flown at Dryden from 1969 to 1979 in an earlier high-speed program.) Pegasus® space booster is successfully air-launched from the NASA B-52 in one of the first successful flights of a commercially developed space launch vehicle placing a payload into earth orbit. The launch was off the California coast, with a NASANavy payload placed in a polar orbit 320 miles high. First flight in NASA’s first program to investigate laminar flow control at supersonic speeds. Program uses the only two F-16XL prototypes to investigate passive and active methods of reducing turbulence on wing surfaces at high speeds. Final test in a series of eight using B-52 No. 008 to validate drag chute deployment system for use on space shuttle to improve their landing efficiency. The tests with 008 were on the lakebed and main runway. Position of Dryden site manager redesignated as director in reorganization that strengthens Dryden’s role as a national flight research installation, with Ken Szalai, chief of Dryden’s Research Engineering Division, named to new position. Dryden personnel number 430. Full-scale X-30 structural test component, representing a wing control surface, arrives at Dryden’s Thermostructural Research Laboratory for loads and temperature testing. First flight of F-18 High Alpha (Angle-of-Attack) Research Vehicle (HARV) with thrust vectoring system engaged to enhance control and maneuvering at high angles of attack; 104th flight of the HARV, which arrived at Dryden Oct. 22, 1984, and initially flew a series of missions without thrust vectoring to obtain experience with aerodynamic measurements at high angles of attack and to develop the flight research techniques needed for this measurement. Dec. 1, 1987 Sept. 18, 1989 Dec. 18, 1989 Feb. 15, 1990 Apr. 5, 1990 May 3, 1990 Oct. 25, 1990 Dec. 3, 1990 May 15, 1991 Jul. 12, 1991 ® Pegasus is a registered trademark of Orbital Sciences Corporation, Fairfield, Virginia. 6 Aug. 14, 1991 First all-NASA SR-71 flight with research pilots Steve Ishmael and Rogers Smith in the cockpit. It was the first Mach 3 mission flown at Dryden since the final YF-12 flight Oct. 31, 1979. Seven-year X-29 Advanced Technology Demonstrator program ends after 362 research missions with the two forward-swept wing aircraft. No. 1 aircraft was flown 242 times to validate design concepts. X-29 No. 2 was flown 120 times in high-angle-of-attack studies. Joint USAF/NASA program later flies No. 2 in vortex control study. Dryden aeronautical engineer Marta Bohn-Meyer becomes first female crewmember to fly in an SR-71. Tests of pressure-sensitive luminescent paint end, opening door for a new method of measuring surface pressures on aircraft. F-18 high-angle-of-attack research aircraft, with pilot Ed Schneider, achieves design point of roughly 70 degrees angle of attack. First flight of an X-31 aircraft from Dryden following relocation of X-31 International Test Organization from Air Force Plant 42 in Palmdale, in a DOD study of thrust vectoring for air combat at high angles of attack. Maiden landing of the space shuttle Endeavour, built to replace Challenger. Landing is viewed by an estimated 125,000 people, including 2,500 school students. Single-day Dryden record of six missions tied by X-29 No. 2 after the aircraft returns to flight for a 60-flight USAF/NASA study using vortex flow controls on nose to study improved control at high angles of attack. Integrated Test Facility (ITF) officially opens, giving Dryden a unique capability to carry out interdependent systems testing, systems troubleshooting, and rapid pre-and post-flight systems checks on several aircraft simultaneously. Flights begin with Dryden’s CV-990 Landing Systems Research Aircraft, equipped with a space shuttle landing gear fixture that later led to increased orbiter cross wind landing limits at the Kennedy Space Center, and aided in the decision to resurface the Kennedy runway. Judy Janisse Child Development Center is dedicated. The $700,000 facility is named after a former NASA employee, killed in a commercial airline accident, who was instrumental in the development of the center. NASA SR-71 flies on first science mission, taking a JPL ultraviolet camera to 85,000 feet for night photo studies. Flight was also first SR-71 night mission. The F-15 Highly Integrated Digital Electronic Control (HiDEC) is landed using only engine power to turn, climb, and descend. Gordon Fullerton is pilot on milestone event. The thrust-vectored X-31 executes a minimum radius 180-degree turn—the “Herbst Maneuver”—while flying at more than 70-degrees angle of attack, well beyond the aerodynamic limits of any other aircraft. 7 Sept. 30, 1991 Oct. 3, 1991 Nov. 1, 1991 Dec. 12, 1991 Apr. 23, 1992 May 16, 1992 Jul. 1, 1992 Oct. 24, 1992 Dec. 22, 1992 Jan. 4, 1993 Mar. 9, 1993 Apr. 21, 1993 Apr. 29, 1993 May 21, 1993 First research flight with Dryden’s F-18 Systems Research Aircraft checks out an electric actuator that monitors and controls one of the aircraft’s ailerons, and becomes a testbed for advanced electric and fiber optics components. Modified F-15 called ACTIVE—Advanced Control Technology for Integrated Vehicles—replaces the HiDEC as Dryden’s integrated systems aircraft. The ACTIVE F-15 features forward canards and will be fitted with thrust-vectoring nozzles to study their use for pitch and yaw control. Replica of X-15 rocket research aircraft, displayed at the corner of Lilly and Lakeshore Drives, is dedicated. The space shuttle Columbia, on mission STS-58, lands at 7:06 a.m. (PST), the last planned landing of a shuttle at Edwards. Nearly 35,000 people, including about 5000 Dryden guests, view the event. The Perseus remotely piloted aircraft flies for the first time in a project to develop technologies to be used to construct and fly unpiloted vehicles on high-altitude science missions. Final flight of an F-104 at Dryden, a symbolic farewell with NASA 826, is piloted by Tom McMurtry, Chief, Flight Operations Division. First acquired in 1956, 11 F-104s flew at Dryden over a 38-year period as chase and research aircraft. Last research mission with NASA 826 was Jan. 31, 1994. The other remaining F-104, NASA 825, was flown on its last research mission Jan. 24, 1994. Dryden named a NASA Center again. Transition period to institute independent administrative functions ends Sept. 30, 1994. Ten thousandth research mission is logged by Dryden’s Western Aeronautical Test Range, a flight with the F-18 HARV. Facility was developed in the 1950s to support the X-15 program. Twenty-fifth anniversary of Apollo 11 features salute to Dryden’s Lunar Landing Research Vehicle (LLRV), used to develop moon-landing training techniques. Sixth and last Pegasus® mission using NASA B-52 No. 008 as the launch vehicle is successful. Future airborne launches to be with an L-1011 owned and operated by Pegasus developer, Orbital Sciences Corp. X-31 logs 438th flight, new record for experimental aircraft. Record holder had been X-29, set on its last flight in 1992. Dryden assumes full Center status, as NASA’s Center of Excellence in Flight Research. NASA personnel number 465. X-31 completes final research flights, making a total of 555 for the program. The NASA B-52, No. 008, turns 40 years old. Based at Dryden since mid-1959, it is the oldest B-52 still flying. Jun. 15, 1993 Jun. 24, 1993 Nov. 1, 1993 Dec. 21, 1993 Feb. 3, 1994 Mar. 1, 1994 Mar. 18, 1994 Jul. 20, 1994 Aug. 3, 1994 Aug. 4, 1994 Oct. 1, 1994 May 13, 1995 Jun. 11, 1995 8 Aug. 11, 1995 CV-990 Landing Systems Research Aircraft completes study of space shuttle landing gear, with a total of 155 research flights. Final tests subjected orbiter wheels to total failure modes on lakebed surface and concrete runway. Aided by NASA-developed propulsion controlled aircraft (PCA) system, an MD-11 makes first-ever, safe landings of an actual transport aircraft using only engine power for control. Pathfinder sets new altitude record for solar-powered aircraft. The remotely controlled, unpiloted prototype attained an altitude 50,500 ft during a nearly 12-hour flight. Solar cells on the top surface of the all-wing aircraft power six electric, propeller-turning motors for propulsion. Pathfinder is part of the NASA Environmental Research Aircraft and Sensor Technology (ERAST) Program. First flight of the F-16XL with the active glove installed. The two-seat F-16 XL was piloted by Dana Purifoy, and begins a program researching laminar flow at supersonic speeds using a suction panel that covers 60 percent of the wing chord. Previous studies with the single-seat F-16XL used a glove that covered only 20 percent of the chord. Center Director Ken Szalai renames the Integrated Test Facility as the Walter C. Williams Research Aircraft Integration Facility. Improved software enables an MD-11 to make a final landing at Edwards without the need for the pilot to manipulate the flight controls while using only engine power for control. NASA announces award of X-33 contract to Lockheed Martin Corp. to design, build, and fly a vehicle that will demonstrate advanced technologies to dramatically increase reliability and lower the cost of putting a pound of payload in space. The test vehicle was projected to fly from DFRC in the year 2000. F-18 High Alpha Research Vehicle (HARV) makes final flight in 385-flight research program that investigated improved maneuverability of tactical aircraft at high angles of attack. F-15 ACTIVE research aircraft conducts first thrust vectoring of engine exhaust at speeds approaching Mach 2. First flight of Tu-144LL flying laboratory inaugurates year-long flight test program in support of NASA’s High-Speed Research program. Year-long Supersonic Laminar-Flow Control program concludes with 45th flight on highly modified F-16XL research aircraft. Program proved that laminar–or smooth–airflow could be obtained over a significant portion of an aircraft wing’s chord at speeds of Mach 2 by use of a suction system pulling turbulent boundary-layer air through tiny holes in the wing skin. Aug. 29, 1995 Sept. 11, 1995 Oct. 13, 1995 Nov. 17, 1995 Nov. 30, 1995 Jul. 2, 1996 Sept. 6, 1996 Nov. 1, 1996 Nov. 24, 1996 Nov. 26, 1996 More information on the history of Dryden Flight Research Center can be found in the references 1 through 9 on the following page. 9 REFERENCES 1. Hallion, R. P.: On the Frontier: Flight Research at Dryden 1946–1981, NASA SP-4303, 1984. 2. Wallace, Lane E.: Flights of Discovery 50 Years at the NASA Dryden Flight Research Center, NASA SP-4309, 1997. 3. Stillwell, Wendell H.: X-15; Research Results with a Selected Bibliography, NASA SP-60, 1965. 4. Thompson, Milton O.: At the Edge of Space, Smithsonian Institution Press, 1992, TL789.8.U6X578. 5. Powers, Sheryll Goecke: Women in Flight Research at NASA Dryden Flight Research Center From 1946 to 1995, Monograph in Aerospace History No. 6, 1997. 6. Day, Richard E.: Coupling Dynamics in Aircraft: A Historical Perspective, NASA SP-532, 1997. 7. Saltzman, Edwin J.; and Ayers, Theodore G.: Selected Examples of NACA/NASA Supersonic Flight Research, NASA SP-513, May 1995. 8. Reed, R. Dale (with Darlene Lister; foreward by Chuck Yeager): Wingless Flight: The Lifting Body Story, NASA SP-4220, 1997. 9. Kempel, Robert W.; Painter, Weneth D.; and Thompson, Milton O.: Developing and Flight Testing the HL-10 Lifting Body: A Precursor to the Space Shuttle, NASA RP-1332, April 1994. 10 ACRONYMS AND ABBREVIATIONS AAS AASE ACTIVE AEDC AFB AFFTC AFWAL AGARD AHS AIAA AICHE AIR AIST ARC ARS ASCE ASEE ASME BuAero CAI CASI CP CR DGLR American Astronautical Society, Washington, D. C. Department of Aerophysics and Aerospace Engineering, Mississippi State University, Starkville, Mississippi Advanced Control Technology for Integrated Vehicles–a NASA Dryden test program Arnold Engineering Development Center, Arnold Air Force Base, Tennessee Air Force Base Air Force Flight Test Center, Edwards, California Air Force Wright Aeronautical Laboratory, Dayton, Ohio Advisory Group for Aeronautical Research & Development, Paris, France American Helicopter Society, Alexandria, Virginia American Institute of Aeronautics and Astronautics, Washington, D. C. American Institute of Chemical Engineers, New York, New York Aerospace Information Report–a military report Agency of Industrial Science and Technology, Japan Ames Research Center, Moffett Field, California American Rocket Society–became AIAA in 1963 with IAS American Society of Civil Engineers, Washington, D. C. American Society for Engineering Education, Washington, D. C. American Society of Mechanical Engineers, New York, New York Bureau of Aeronautics, Department of the Navy, 1921–1959 Canadian Aeronautical Institute, Ottawa, Ontario NASA Center for Aerospace and Scientific Information, Hanover, Maryland Conference Proceedings, NASA report Contractor Report, NASA report Deutsche Gesellschaft fuer Luft-und Raumfahrt, Germany 11 DOD DTIC ERAST FAA FED FRC HARV HiDEC HSCT HSFRS IAA IAS ICAS ICASE ICASSP IEEE IES IFAC ISA ITC ITEA ITF IUTAM JPL JSASS Department of Defense, Washington, D. C. Defense Technical Information Center, Ft. Belvoir, Virginia Environmental Research Aircraft and Sensor Technology program—a NASA Dryden test program Federal Aviation Agency / Administration, Washington, D. C. Fluids Engineering Division (ASME), New York, New York Flight Research Center, NASA facility at Edwards AFB 1959–1976 F-18 high alpha research vehicle F-15 highly integrated digital electronic control High speed civil transport High-Speed Flight Research Station International Aerospace Abstracts–a monthly publication of recent international and AIAA Conference information in abstract form, AIAA, New York, New York Institute of Aeronautical Sciences, Inc.—became AIAA in 1963 with ARS International Council of Aeronautical Sciences, Les Mureaux Cedex, France Institute for Computer Applications in Science and Engineering, Hampton, Virginia International Conference on Acoustics, Speech, and Signal Processing Institute of Electrical and Electronics Engineers, New York, New York Institute of Environmental Sciences, Mount Prospect, Illinois International Federation of Automatic Control, Uppsala, Sweden Instrument Society of America, Triangle Park, North Carolina International Telemetering Conference International Test and Evaluation Association, Fairfax, Virginia Integrated Test Facility, at Dryden since 1987 International Union of Theoretical and Applied Mechanics, Stuttgart, Germany Jet Propulsion Laboratory, Pasadena, California Japan Society for Aeronautical and Space Sciences, Japan 12 LLRV MIT NACA NASA NIWC NTIS PCA RAE RECON lunar landing research vehicle Massachusetts Institute of Technology, Cambridge, Massachusetts National Advisory Committee for Aeronautics–a NASA Precursor National Aeronautics and Space Administration, Washington, D. C. National Institute of Materials and Chemical Research, Japan National Technical Information Service, Springfield, Virginia propulsion controlled aircraft Royal Aircraft Establishment, England A CASI database to archive scientific and technical reports RECONplus A CASI database to archive scientific and technical reports RM RP SAE SCA SETP SFTE SIAM SP SPIE TM TN TP USAF USN Research Memo, a NACA report Reference Publication, NASA Report Society of Automotive Engineers, Warrendale, Pennsylvania shuttle carrier aircraft Society of Experimental Test Pilots, Lancaster, California Society of Flight Test Engineers, Lancaster, California Society for Industrial and Applied Mathematics, Philadelphia, Pennsylvania Special Publication, a NASA report The International Society for Optical Engineering, Bellingham, Washington Technical Memorandum, a NASA report Technical Note, a NACA or NASA report Technical Paper, a NACA or NASA report United States Air Force United States Navy USN/NAWC United States Navy/Naval Air Warfare Center 13 ORDERING INFORMATION Ordering sources for the different types of materials given below: Sources Document Delivery/AIAA Dispatch 800-662-1545 816-926-8794 FAX dispatch@lhl.lib.mo.us http://www.lhl.lib.mo.us/pubserv/ AIAA/dispatch.htm National Technical Information Service (NTIS) 703-487-4650 703-321-8547 FAX orders@ntis.fedworld.gov http://ntis.gov NASA Center for Aerospace Information (CASI) 301-621-0390 301-621-0134 FAX help@sti.nasa.gov http://www.sti.nasa.gov Defense Technical Information Center (DTIC) 800-225-3842 703-767-8228 FAX help@dtic.mil http://dticam.dtic.mil/ NASA libraries or NASA CASI Libraries Types of material AIAA papers and worldwide literature from conferences and periodicals available through AIAA. Accession number examples: A90-12345 90A12345 AIAA paper no. 97-1234 Report literature having no distribution limitations. N95-12345 95N12345 19970012345 AD-A123456 Report literature having no distribution limitations. Report literature with some type of distribution limitation. Report literature from U. S. Government Agency or AGARD. N95-12345 95N12345 19970012345 X97-12345 97X12345 AD-A123456 Pre-1962 reports and papers. Books 87H12345 93R12345 TL123.C66.D7 14 KEY TO CITATIONS Typical Citation and Abstract 1993. Burcham, Frank W., Jr.; Maine, Trindel; and 3 Wolf, Thomas. Flight Testing and Simulation of an F15 Airplane Using Throttles for Flight Control. 6 7 5 4 NASA TM-104255, H-1826, NAS 1.15:104255, AIAA 8 Paper 92-4109. Presented at the AIAA Flight Test 9 Conference, Hilton Head, SC, 24 Aug. 1992. August 11 12 10 1992, 92N32864, # (see also 2004). Flight tests and simulation studies using the throttles of an F-15 airplane for emergency flight control have been conducted at the NASA Dryden Flight Research Center. The airplane and the simulation are capable of extended up-and-away flight, using only throttles for flight path control. Initial simulation results showed that runway landings using manual throttles-only control were difficult, but possible with practice. Manual approaches flown in the airplane were much more difficult, indicating a significant discrepancy between flight and simulation. Analysis of flight data and development of improved simulation models that resolve the discrepancy are discussed. An augmented throttle-only control system that controls bank angle and flight path with appropriate feedback parameters has also been developed, evaluated in simulations, and is planned for flight in the F-15. 1 2 1 Chronological number of citation 2 Author(s) 3 Title 4 NASA publication number 5 NASA Dryden production number 6 GPO number 7 Assigned conference publication number 8 Conference name, place, and date 9 Date of publication (underlined) 10 Accession number 11 Available on microfiche (#) 12 Chronological number of cross-reference citation 15 BIBLIOGRAPHY 1947 Technical Publications 1. Beeler, De E.; and Gerard, George: Wake Measurements Behind a Wing Section of a Fighter Airplane in Fast Dives. NACA TN 1190, March 1, 1947, 93R11438. Wake measurements made in a vertical plane behind a wingsection of a fighter airplane are presented for a range of Mach number up to 0.78. Since evidences of reverse flow were found in a large part of the surveys—possibly because of interference of the rake support—the computed profile-drag coefficients are considered to be only qualitative. The results showed that the large increase in drag coefficient beyond critical Mach number indicated by wind-tunnel tests was also obtained under flight conditions and that the wake width was extended sharply when shock was encountered. The wake extension occurred first at the upper surface since the highest local velocity was obtained on that surface. The large increase in drag coefficient for the wing section tested did not occur until after the critical Mach-number had been exceeded by approximately 0.05. Comparison of the profile-drag measurements with total airplane drag measurements showed that the large increases in drag in both cases started to occur at the same value of Mach number. The results further indicated that wake measurements made in three dimensional-flow after shock had occurred cannot, in general, be interpreted in terms of section profile-drag coefficient because of the existence of the strong lateral flow indicated by tuft behavior in the dead-air region behind the shock. 2. Beeler, De E.: Air-Flow Behavior Over the Wing of an XP-51 Airplane as Indicated by WingSurface Tufts at Subcritical and Supercritical Speeds. NACA RM L6L03, April 24, 1947, 86H23988. Results are presented in this report of the air-flow behavior over the wing of an XP-51 airplane including photographs of tuft attached to the wing surface and chordwise pressure distributions. A comparison of tuft studied is made of the flight results with those obtained from wind-tunnel tests. The results indicate that steady flow is obtained over the wing until the critical speed has been exceeded by about 0.04 to 0.05 in Mach numbers. At higher Mach numbers the flow is unsteady and becomes very rough and turbulent over the rear 50 percent of the chord after the limit maximum pressure coefficient has been reached. Observation of surface tufts alone without benefit of prevailing pressure distributions may indicate separated flow before separation actually occurs. Comparisons made of the flight and wind-tunnel data show a similar tuft behavior throughout the Mach number range. 16 1948 Technical Publications 3. Beeler, De E.; and Mayer, John P.: Measurements of the Wing and Tail Loads During the Acceptance Tests of the Bell XS-1 Research Airplane. NACA RM L7L12, April 1948, 87H24183. During the acceptance tests of the XS-1 airplane, strain-gage measurements were made of wing and tail loads up to a Mach number of 0.80. The maximum lift and buffet boundaries were also determined. The loads encountered were well within the design loads and showed fairly good agreement with wind-tunnel and calculated data. EC72-3431 XS-1 (X-1) Airplane 4. Williams, Walter C.; and Beeler, De Elroy: Results of Preliminary Flight Tests of the XS-1 Airplane (8-Percent Wing) to a Mach Number of 1.25. NACA RM L8A23A, April 6, 1948, 86H18960. The data obtained in flight with the XS-1 airplane with 8-percent-thick wing up to and beyond the speed of sound at an altitude of 37,000 feet and above show that most of the trim and force changes expected in the transonic range have been experienced. Although conditions are not normal, the airplane can be flown under control through a Mach number of 1 at altitudes of 37,000 feet and above. In detail, the following has been noted: (1) Buffeting has been experienced in level flight but has been mild. The horizontal-tail loads associated with the buffeting have been small. (2) The airplane has experienced longitudinal trim changes in the speed range from 0.8 up to 1.25. The largest control force associated with these trim changes was 25 pounds. The pilot has been able to control the airplane. The relatively small magnitude of the control force may be attributed to the small size of the elevator and the high altitude of the flight. (3) The elevator effectiveness has decreased more than 50 percent in going from a Mach number of 0.7 to 0.87. There is evidence of further reduction in elevator effectiveness above a Mach number of 0.87. This loss in elevator effectiveness has affected the magnitude of the trim changes as noted by the pilot but the actual trim changes for the most part have been caused by changes in the wing-fuselage moment. (4) No aileron buzz or associated phenomena have been experienced. The airplane becomes right wing heavy with increasing Mach number up to a Mach number of 1.10, but can be trimmed with the ailerons. 5. McLaughlin, Milton D.; and Clift, Dorothy C.: Results Obtained During a Dive Recovery of the Bell XS-1 Airplane to High Lift Coefficients at a Mach Number Greater Than 1.0. NACA RM L8C23A, April 6, 1948, 86H31136. Measured quantities are presented which were obtained on the Bell XS-1 airplane with an 8-percent-thick wing and a 6-percent-thick horizontal tail during a dive recovery at a Mach number greater than 1.0. The data obtained show that it is possible to obtain fairly high load factors with the airplane at Mach numbers greater than 1.0 if the stabilizer is used for longitudinal control. Lift coefficients approaching low-speed maximum-lift values have been obtained at a Mach number of 1.1 with no indication that these values are the maximum obtainable for the airplane. At the Mach number and lift coefficient reported, there was little or no buffeting. 6. Drake, Hubert M.; McLaughlin, Milton D.; and Goodman, Harold R.: Results Obtained During Accelerated Transonic Tests of the Bell XS-1 Airplane in Flights to a Mach Number of 0.92. NACA RM L8AO5A, April 19, 1948, 86H18537. An accelerated flight program using the Bell XS-1 airplane has been undertaken to explore the transonic-speed range. The flying was done by an Air Force pilot, and the data reduction and analysis were made from NACA instrumentation by NACA personnel. This paper presents the results of tests obtained up to a Mach number of 0.92 at altitudes around 30,000 feet. The data obtained show that the airplane has experienced most of the difficulties expected in the transonic range, but that it can be flown satisfactorily to a Mach number of at least 0.92 at altitude above 30,000 feet. Longitudinal trim changes have been experienced but the forces involved have been small. The elevator effectiveness decreased about one-half with increase of Mach number from 0.70 to 0.87. Buffeting has been experienced in level flight but it has been mild and the associated tail loads have been small. No aileron buzz or other flutter phenomena have been noted. 7. Williams, Walter C.; Forsyth, Charles M.; and Brown, Beverly P.: General Handling-Qualities Results Obtained During Acceptance Flight Tests of the Bell XS-1 Airplane. NACA RM L8A09, April 19, 1948, 86H18538. During the acceptance tests conducted by the Bell Aircraft Corporation on the Bell XS-l transonic research airplane, NACA instruments were installed to measure the stability and 17 control characteristics and the aerodynamic loads. Tests were made in gliding flight and in powered flight. Two Bell XS-l airplanes were used during the program: one airplane had wing and tail thicknesses of 10 percent and 8 percent of the chord, respectively, and the other airplane had wing and tail thicknesses of 8 percent and 6 percent of the chord, respectively. The results for the stability and control measurements are presented in this paper. 8. Williams, Walter C.: Limited Measurements of Static Longitudinal Stability in Flight of Douglas D-558-1 Airplane (BuAero No. 37971). NACA RM L8E14, June 24, 1948, 86H17021. During airspeed calibration flights of the D-558-1 airplane being used by NACA for high-speed-flight research, some measurements were obtained of the static longitudinal stability up to a Mach number of 0.85. These data showed that the airplane possessed positive static longitudinal stability up to a Mach number of 0.80. A trim change in the nose-down direction occurred for Mach numbers above 0.82. E49-090 D-558-1 Airplane 9. Beeler, De E.; McLaughlin, Milton D.; and Clift, Dorothy C.: Measurements of the Chordwise Pressure Distributions Over the Wing of the XS-1 Research Airplane in Flight. NACA RM L8G21, August 4, 1948, 86H25805. Measurements of the chordwise pressure distribution over the 8-percent-thick wing of the XS-1 research airplane have been made at a section near the midspan of the left wing. Data presented are for a Mach number range of 0.75 to 1.25 at a normal-force coefficient of about 0.33 and for normal-force coefficients up to 0.93 at a Mach number of approximately 1.16. The results show that there is a rearward shift of section center of load with an increase in Mach number due to the rearward movement of shock with a corresponding extension of the region of supersonic flow. The load center moves from about 25 to 51 percent of the chord as the Mach number is increased for 0.75 to 1.25. During the rearward movement of load from the forward to rearward limit position, there is a rapid and large shift of the center of load within these limits for a Mach number range of 0.82 to 0.88. It is expected that large changes in trim, with corresponding large changes in load factor at low altitude, may occur within this Mach number range. 10. Drake, Hubert M.; Goodman, Harold R.; and Hoover, Herbert H.: Preliminary Results of NACA Transonic Flights of the XS-1 Airplane With 10-PercentThick Wing and 8-Percent-Thick Horizontal Tail. NACA RM L8I29, October 13, 1948, 87H24234. Contains results of exploratory flights at altitudes of about 40,000 feet to a maximum Mach number of 1.06. Data are presented showing the longitudinal trim changes, elevator effectiveness in producing acceleration, and rudder effectiveness as a function of Mach number. Data on lateral oscillations are also presented. 11. Goodman, Harold R.; and Drake, Hubert M.: Results Obtained During Extension of U.S. Air Force Transonic-Flight Tests of XS-1 Airplane. NACA RM L8128, November 16, 1948, 86H18957. Limited data covering extension of the U.S. Air Force transonic-flight tests of the XS-1 airplane are presented. These data show that successful flight to a Mach number of 1.35 has been achieved at altitudes above 40,000 feet. Longitudinal trim changes were experienced to the highest Mach numbers attained, with the wheel forces remaining small and the pilot able to control the airplane with ease. The airplane becomes right-wing heavy above a Mach number of 0.8 but can be trimmed with the aileron. No aileron buzz or flutter phenomena have been encountered. Buffeting has been light, when encountered in the range of Mach number and lift coefficient covered by these data. 12. Matthews, James T., Jr.: Effect of Downwash on the Estimated Elevator Deflection Required for Trim of the XS-1 Airplane at Supersonic Speeds. NACA RM L8H06A, November 1948. This report contains the results of an investigation to determine from linearized theory, which has recently become available, the downwash at supersonic speeds at the tail of the XS-1 airplane and the effect of the downwash on the elevator deflection required for trim. The results are presented in the form of curves showing the variation of downwash angle with angle of attack and elevator deflection required for trim plotted against Mach number. The calculations indicate that increasing up-elevator deflection is required with increasing Mach number (unstable variation) in level flight between Mach numbers of 1.1 and 1.6. A slight reduction in upelevator deflection occurs between Mach numbers of 1.6 and 2.0. The stabilizer angle has a similar variation, that is, unstable up to a Mach number of about 1.6 and then becoming slightly stable up to a Mach number of 2.0. The reduction of downwash with increasing Mach number is not the main cause of the increase in up-elevator deflection. The main reasons for this trend are that the pitching-moment 18 coefficients due to the wing camber, the wing lift, and the lift of the stabilizer are all in a nose-down direction, and as the Mach number increases, these pitching-moment coefficients apparently decrease less rapidly than the elevator effectiveness. 1949 Technical Publications 13. Drake, Hubert M.: Stability and Control Data Obtained From First Flight of X-4 Airplane. NACA RM L9A31, February 7, 1949, 86H17466. NACA instrumentation has been installed in the X-4 airplanes to obtain stability and control data during the Northrop conducted acceptance tests. The results of the first flight of the X-4 number 1 airplane are presented in this report. These data were obtained for a center-of-gravity position of about 22 percent of the mean aerodynamic chord. A maximum indicated airspeed and pressure altitude of 290 miles per hour and 11,000 feet, respectively, were obtained during the flight. Results of the flight indicated that the airplane is slightly unstable, stick fixed, in gear-up, flapsup configuration for a center-of-gravity position at 21.4 percent of the mean aerodynamic chord. The pilot reported that it was difficult to maintain steady flight in this configuration. There was no indication of a snaking or lateral oscillation for the speed range covered. For gear-down configuration at low lift coefficients with the center of gravity at 22.4 and 21.6 percent of the mean aerodynamic chord the airplane was longitudinally stable; however, at high lift coefficients, it was indicated that the airplane was longitudinally unstable. The rudder effectiveness appeared to be low in the gear-down, low-speed condition. The maximum rate of rudder motion of 25 degrees per second available with the present control system was considered by the pilot to be too slow. E-359 X-4 Airplane 14. Williams, Walter C.: Flight Measurement of the Stability Characteristics of the Douglas D-558-1 Airplane (BuAero No. 37971) in Sideslips. NACA RM L8E14A, April 18, 1949, 86H17022. Measurements have been made of the stability characteristics of the D-558-1 airplane in steadily increasing sideslips at various Mach numbers from 0.50 to 0.80 at 10,000 feet altitude and at Mach numbers from 0.50 to 0.84 at 30,000 feet altitude. The results of these tests show that the apparent directional stability of the airplane is high and increases with increasing Mach number and dynamic pressure. The dihedral effect is positive at all speeds, there is little or no change in pitching moment with sideslip, and the cross-wind force is positive. 15. Barlow, William H.; and Lilly, Howard C.: Stability Results Obtained With Douglas D-558-1 Airplane (BuAero No. 37971) in Flight Up to a Mach Number of 0.89. NACA RM L8K03, April 22, 1949, 86H17024. Measurements have been made of some of the high-speed characteristics of the D-558-1 airplane up to a Mach number of 0.89. The results of these tests showed that the stabilizer incidence drastically affected the longitudinal trim characteristics above a Mach number of 0.80. With a stabilizer incidence of 2.3 degrees, the airplane became nose heavy above a Mach number of 0.8. With a stabilizer incidence of 1.4 degrees, the airplane became tail heavy above a Mach number of 0.83. The airplane also became right-wing heavy above a Mach number of 0.84 and the airplane felt uncertain laterally to the pilot. The longitudinal stability in accelerated flight was positive throughout the speed range from a Mach number of 0.50 to 0.80 and increased above a Mach number of 0.675. The buffet boundary was defined up to a Mach number of 0.84 and was similar to that for the Bell XS-1 airplane with the same wing section, 65-110. 16. Drake, Hubert M.: Measurement of the Dynamic Lateral Stability of the Douglas D-558-1 Airplane (BuAero No. 37971) in Rudder Kicks at a Mach Number of 0.72. NACA RM L9D06A, May 31, 1949, 86H17025. Contains results of flight measurements of the dynamic lateral stability of the Douglas D-558-1 airplane at a Mach number of 0.72 and a pressure altitude of 8500 feet. 17. Drake, Hubert M.: Measured Characteristics of the Douglas D-558-1 Airplane (BuAero No. 37971) in Two Landings. NACA RM L9D20A, June 3, 1949, 86H18900. Records were obtained of two landings of the Douglas D-558-1 airplane made during the stability and control investigation. These two records show that the maximum normal-force coefficient used during the landings, 0.95, was considerably below the maximum, 1.2, estimated to be available. The approaches were made at 150 percent of the possible minimum speed, and the actual contacts were at about 115 percent of minimum speed. The rate of descent in the approach was 1200 to 1800 feet per minute at the start of the landing flare. 18. Drake, Hubert M.: Measurements of Aileron Effectiveness of Bell X-1 Airplane Up to a Mach Number of 0.82. NACA RM L9D13, June 20, 1949, 86H27619. Abrupt, rudder-fixed aileron rolls have been made with the Bell X-1 airplane having a 10-percent-thick wing in glides to a Mach number of 0.82 at about 30,000 feet pressure altitude. Aileron movements were between one-fourth and one-half of full deflection. These aileron rolls indicate that Mach number has little effect on the aileron effectiveness up to a Mach number of 0.82. 19. Williams, Walter C.: Results Obtained From Second Flight of X-4 Airplane (USAF No. 46-676). NACA RM L9F21, July 18, 1949, 86H17468. NACA instrumentation has been installed in the X-4 airplanes to obtain stability and control data during the Northrop-conducted acceptance tests. The results of the second flight of the X-4 number 1 airplane are presented in this report. This flight was made with the center of gravity at 19.7 percent of the mean aerodynamic chord and with the rudder-boost system removed. The results of the flight showed that the longitudinal stability was positive in the clean condition and in the gear-down, flaps-up condition. Records taken during landing approach and in a steady run at 170 miles per hour showed that the lateral oscillation is poorly damped. The pilot reported that the rudder control was adequate. 20. Drake, Hubert M.; and Wall, Helen L.: Preliminary Theoretical and Flight Investigation of the Lateral Oscillation of the X-1 Airplane. NACA RM L9F07, July 19, 1949, 86H17642, 87H24281. A small-amplitude, undamped, lateral oscillation has been encountered in flight tests of the X-1 airplane. The oscillation occurs in subsonic and supersonic flight, in maneuvers, and power on and off. The calculations indicate that a change, in the positive direction, of the inclination of the principal axis with respect to the flight path should have a considerable stabilizing effect. 19 E49-059 D-558-1 Airplane 21. Goodman, Harold R.; and Yancey, Roxanah B.: The Static-Pressure Error of Wing and Fuselage Airspeed Installations of the X-1 Airplanes in Transonic Flight. NACA RM L9G22, July 22, 1949, 86H63671. Measurements were made in the transonic speed ranges of the static-pressure position error at a distance of 0.96 chord ahead of the wing tip of both the 8-percent-thick-wing and the l0-percent-thick-wing X-l airplanes, and at a point 0.6 maximum fuselage diameter ahead of the fuselage nose of the X-l airplanes. 22. Drake, Hubert M.: Measurements of Aileron Effectiveness of the Bell X-1 Airplane at Mach Numbers Between 0.9 and 1.06. NACA RM L9G19A, August 4, 1949, 87H24295. Presents results of flight measurements of aileron effectiveness of the X-1 airplane up to a Mach number of 0.94. The data indicate a 75 percent loss of aileron effectiveness between M = 0.82 and M = 0.94. 23. Valentine, George M.: Stability and Control Data Obtained From Fourth and Fifth Flights of the Northrop X-4 Airplane (USAF No. 46-676). NACA RM L9G25A, August 4, 1949, 86H17473. NACA instrumentation has been installed in the Northrop X-4 airplane to obtain stability and control data during the Northrop-conducted acceptance tests. The results of the fourth and fifth flights of the Northrop X-4 number 1 airplane are presented in this paper. These data were obtained for a center-of-gravity position of approximately 19.5 percent of the mean aerodynamic chord. The results of this flight showed that the directional stability as measured in steadily increasing sideslips was positive and high and that the effective dihedral was positive. The results also show the airplane to be longitudinally stable, stick fixed, with the center of gravity at 19.5 percent of the mean aerodynamic chord. 24. Sjoberg, Sigurd A.: Preliminary Measurements of the Dynamic Lateral Stability Characteristics of the Douglas D-558-II (BuAero No. 37974) Airplane. NACA RM L9G18, August 18, 1949, 86H17031. This paper presents some data on the dynamic lateral stability characteristics of the Douglas D-558-II (BuAero No. 37974) airplane. For the airplane in the clean condition, the lateral oscillations are lightly damped. In the landing condition, the airplane performs a constant-amplitude lateral oscillation. E-1441 D-558-II Airplane 25. Williams, Walter C.: Results Obtained from Third Flight of Northrop X-4 Airplane (USAF No. 46-676). NACA RM L9G20A, September 9, 1949, 86H17471. NACA instrumentation has been installed in the Northrop X-4 airplane to obtain stability and control data during the Northrop-conducted acceptance tests. The results of the third flight of the X-4 number 1 airplane are presented in this paper. The results of this flight showed that the directional stability as measured in steadily increasing sideslips was positive and high and that the lateral stability was positive. 26. Goodman, Harold R.: The Static-Pressure Error of a Wing Airspeed Installation of the McDonnell XF-88 Airplane in Dives to Transonic Speeds. NACA RM SL9I12, September 23, 1949. Measurements were made, in dives to transonic speeds, of the static-pressure position error at a distance of one chord ahead of the wing tip of the McDonnell XF-88 airplane. The airplane incorporates a wing which is swept back 35° along the 0.25-chord line and utilizes a 65-series airfoil with a 9-percent-thick section perpendicular to the 0.25-chord line. The section in the stream direction is approximately 8 percent thick. Data up to a Mach number of about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Data at Mach numbers above about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.38. 27. Sjoberg, S. A.; and Champine, R. A.: Preliminary Flight Measurements of the Static Longitudinal Stability and Stalling Characteristics of the Douglas D-558-II Research Airplane (BuAero No. 37974). NACA RM L9H31A, October 18, 1949, 86H17086. Contains results of brief flight measurements of the static longitudinal stability and stalling characteristics of the Douglas D-558-II (BuAero No. 37974) research airplane. 20 1950 Technical Publications 28. Matthews, James T.: Results Obtained During Flights 1 to 6 of the Northrop X-4 Airplane (USAF No. 46-677). NACA RM L9K22, January 12, 1950, 86H17475. NACA instruments were installed in the Northrop X-4 number 2 airplane (A.F. No. 46-677) to obtain stability and control data during the acceptance tests conducted by the Northrop Company. The results of flights 1 to 6 are presented in this report. These data were obtained for a center-ofgravity position of about 19.5 to 20.0 percent of the mean aerodynamic chord. The data presented include a time history of a complete pull-up, time histories of several level and accelerated flight runs, and the effect of dive-brake extension on the longitudinal and lateral trim. The pilot reports the mechanical trim device to be unsatisfactory for any stick-free or dynamic stability and control analysis because the stick force cannot be trimmed to zero sufficiently well to permit the stick to be released during a maneuver without the airplane performing a divergence. In addition, the trim device is inoperative when more than 8.0 degrees up elevon angle is required for trim. A short-period longitudinal oscillation with relatively poor damping was present, but this oscillation was not objectionable to the pilot. The airplane has a stable variation of longitudinal-control angle with normal-force coefficient for the indicated airspeed ranges of 180 to 300 miles per hour at about 30,000-foot pressure altitude. Extension of the dive brakes up to ±30 degrees has no appreciable effect on the longitudinal trim at indicated airspeeds of 160 miles per hour with landing gear down, and at airspeeds of 300 miles per hour with the landing gear up, at altitudes of 8,500 and 10,000 feet, respectively. A slight tendency to roll to the left was indicated in the landing-geardown case. 29. Sadoff, Melvin; and Sisk, Thomas R.: Stall Characteristics Obtained From Flight 10 of Northrop X-4 No. 2 Airplane (USAF No. 46-677). NACA RM A50A04, February 27, 1950, 86H30240. NACA instrumentation has been installed in the X-4 airplanes to obtain stability and control data during the acceptance tests conducted by the Northrop Aircraft Corporation. This report presents data obtained on the stalling characteristics of the airplane in the clean and gear-down configurations. The center of gravity was located at approximately 18 percent of the mean aerodynamic chord during the tests. The results indicated that the airplane was not completely stalled when stall was gradually approached during nominally unaccelerated flight but that it was completely stalled during a more abruptly approached stall in accelerated flight. The stall in accelerated flight was relatively mild, and this was attributed to the nature of the variation of lift with angle of attack for the 0010-64 airfoil section, the plan form of the wing, and to the fact that the initial sideslip at the stall produced (as shown by wind-tunnel tests of a model of the airplane) a more symmetrical stall pattern. 21 30. Sjoberg, S. A.: Flight Measurements With the Douglas D-558-II (BuAero No. 37974) Research Airplane. Static Lateral and Directional Stability Characteristics as Measured in Sideslips at Mach Numbers Up to 0.87. NACA RM L50C14, May 19, 1950, 86H17088. Flight measurements were made in sideslips of the static lateral and directional stability characteristics of the Douglas D-558-II (BuAero No. 37974) research airplane. The directional stability of the airplane was positive in both the clean and landing conditions at all test speeds. About 2 degrees of rudder deflection were required to produce 1 degree of sideslip in both the clean and landing conditions. There was no decrease in the effectiveness of the rudder in producing sideslip up to the highest Mach number reached (0.87). 31. Thompson, Jim Rogers; Roden, William S.; and Eggleston, John M.: Flight Investigation of the Aileron Characteristics of the Douglas D-558-I Airplane (BuAero No. 37972) at Mach Numbers Between 0.6 and 0.89. NACA RM L50D20, May 26, 1950, 86H27648. Abrupt, rudder-fixed aileron rolls have been made with the Douglas D-558-I airplane (BuAero No. 37972) at Mach numbers between 0.6 and 0.89. Rolls were made at aileron deflections between one-eighth and one-half the maximum available deflection. The results obtained indicate that the aileron effectiveness is independent of Mach number and deflection within the range investigated. Limited information on the lateral trim and handling qualities of the airplane at high Mach numbers is presented. 32. Sadoff, Melvin; and Sisk, Thomas R.: Longitudinal-Stability Characteristics of the Northrop X-4 Airplane (U.S.A.F. No. 46-677). NACA RM A50D27, June 29, 1950, 86H48899, 86H17213. The results obtained from several recent flights on the Northrop X-4 No. 2 airplane are presented. Information is included on the longitudinal-stability characteristics in straight flight over a Mach number range of 0.38 to about 0.63, the longitudinal-stability characteristics in accelerated flight over a Mach number range of 0.43 to about 0.79, and the short-period longitudinal-oscillation characteristics at Mach numbers of 0.49 and 0.78. It was shown that the stickfixed and stick-free static longitudinal stability, as measured in straight flight, were positive over the test speed range with the center of gravity located at about 18.0 percent of the mean aerodynamic chord. During the longitudinal-stability tests in accelerated flight an inadvertent pitch-up of the airplane occurred at a Mach number of about 0.79 and a normal-force coefficient of about 0.45 (normal acceleration factor, the ratio of the net aerodynamic force along the airplane Z axis to the weight of the airplane = 5), in which the acceleration built up rapidly to the ratio of the net aerodynamic force along the airplane Z axis to the weight of the airplane = 6.2 (which was in excess of the load factor, 5.2, required for demonstration of the airplane) before recovery could be initiated. A comparison of the experimentally determined elevon angles required for balance and the elevon-angle gradients with values estimated from limited wind-tunnel data showed fairly good agreement. Wind-tunnel data, however, were not available in the region where the pitch-up occurred so that an evaluation in this regard was not possible. The short-period oscillation was lightly damped and did not meet the Air Force requirements for satisfactory handling qualities. The pilot, however, did not object to the low damping characteristics of this airplane for small-amplitude oscillations. Theory predicted the period of the short-period longitudinal oscillation fairly well; however, the damping evaluated from the theory indicated considerably greater damping than was actually measured in flight, especially at the higher Mach numbers. 33. Mayer, John P.; Valentine, George M.; and Mayer, Geraldine C.: Flight Measurements with the Douglas D-558-II (BuAero No. 37974) Research Airplane. Determination of the Aerodynamic Center and Zero-Lift Pitching-Moment Coefficient of the Wing-Fuselage Combination by Means of Tail-Load Measurements in the Mach Number Range From 0.37 to 0.87. NACA RM L50D10, July 11, 1950, 86H18984. Determination of the aerodynamic center and zero-lift pitching-moment coefficient of the wing-fuselage combination by means of tail-load measurements in the Mach number range from 0.37 to 0.87. 34. Wilmerding, J. V.; Stillwell, W. H.; and Sjoberg, S. A. Flight Measurements with the Douglas D-558-II (BuAero No. 37974) Research Airplane: Lateral Control Characteristics as Measured in Abrupt Aileron Rolls at Mach Numbers Up to 0.86. NACA RM L50E17, July 20, 1950, 93R15435. Flight measurements were made of the lateral control characteristics of the Douglas D-558-II airplane in abrupt rudder-fixed aileron rolls. In the Mach number range from 0.50 to 0.86 the aileron rolling effectiveness is substantially constant and the rate of change of the maximum wing-tip helix angle with total aileron deflection (rate of change of maximum wing-tip helix angle with total aileron deflection, radians per degree) has a value of 0.0027 radian per degree. Extrapolated data indicate that in this Mach number range full aileron deflection of 30 degrees will produce a maximum wing-tip helix angle pb/2V of about 0.08 radian. As the speed is reduced below a mach number of 0.50 a marked decrease occurs in the maximum value of pb/2V obtainable with a given aileron deflection. This decrease pb/2V occurs because the dihedral effect increases with decrease in speed and the adverse sideslip angles reached in the rolls at low speed are larger. At an indicated airspeed of 150 miles per hour in the landing condition, full aileron deflection will produce a maximum pb/2V of 0.04 radian, which for standard sea-level conditions corresponds to a rolling velocity of 40 degrees per second. In the opinion of the pilots this rolling velocity is sufficiently high for the landing condition with this airplane. 22 It is the opinion of several NACA pilots that the maximum usable rolling velocity is on the order of 2.5 radians per second. In the Mach number range from 0.42 to 0.86 at an altitude of 15,000 feet rolling velocities greater than 2.5 radians per second can be obtained with less than full aileron deflection. The data indicate that in going from high to low lift coefficient the yawing moment due to rolling changes direction. At high lift coefficients the sideslip due to roll is in the same direction as the roll (right roll produces right sideslip), but at low lift coefficients the opposite tendency is present. 35. Gates, Ordway B.; and Sternfield, Leonard: Effect of an Autopilot Sensitive to Yawing Velocity on the Lateral Stability of the Douglas D-558-II Airplane. NACA RM L50F22, August 17, 1950. A theoretical investigation has been made to determine the effect on the lateral stability of the Douglas D-558-II airplane of an autopilot sensitive to yawing velocity. The effects of inclination of the gyro spin axis to the flight path and of time lag in the autopilot were also determined. The flight conditions investigated included landing at sea level, approach condition at 12,000 feet, and cruising at 12,000 feet at Mach numbers of 0.80 and 1.2. The results of the investigation indicated that the lateral stability characteristics of the D-558-II airplane for the flight condition discussed should satisfy the Air Force - Navy period-damping criterion when the proposed autopilot is installed. Airplane motions in sideslip subsequent to a disturbance in sideslip are presented for several representative flight conditions in which a time lag in the autopilot of 0.10 second was assumed. 36. Mayer, John P.; and Valentine, George M.: Flight Measurements With the Douglas D-558-II BuAero No. 37974 Research Airplane. Measurements of the Buffet Boundary and Peak Airplane Normal Force Coefficients at Mach Numbers Up to 0.90. NACA RM L50E31, August 28, 1950, 86H91836. Measurements of the buffet boundary and peak airplane normal force coefficients at Mach numbers up to 0.90. 37. Stillwell, W. H.; Wilmerding, J. V.; and Champine, R. A.: Flight Measurements With the Douglas D-558-II BuAero No. 37974 Research Airplane. Low-Speed Stalling and Lift Characteristics. NACA RM L50G10, September 5, 1950, 86H92789. The low-speed stalling and lift characteristics of the D-558-II airplane were measured in a series of 1-g stalls in four different airplane configurations. With the slats locked closed and the flaps up or down, the airplane was unstable at angles of attack greater than about 9 degrees. With the flaps up this corresponds to a normal-force coefficient of about 0.8 and with the flaps down, about 1.07. Because of this instability, the airplane tended to pitch to high angles of attack; at these high angles of attack, violent rolling and yawing motions sometimes occurred. In one case with the flaps down and the slats locked the airplane went into a spin after pitching up to high angles of attack. The pilots considered the stalling characteristics of the airplane with the slats locked to be very objectionable. No data are presented in this paper on the stalling characteristics in maneuvering flight, but the pilots considered the longitudinal instability particularly objectionable in maneuvering flight. With the slats unlocked and the flaps up or down the airplane was unstable at angles of attack greater than about 23 degrees. Uncontrolled-for rolling and yawing motions due to stalling were present when the airplane was unstable in the high angle-of-attack range. With the slats unlocked and the flaps and landing gear up or down, there was adequate stall warning in the form of buffeting and lateral oscillations of the airplane. With the slats locked, slight buffeting of the airplane occurred at a normalforce coefficient slightly less than the normal-force coefficient at which the airplane became longitudinally unstable. With the flaps up and the slats locked, the highest normal-force coefficient obtained was 1.13 at an angle of attack of about 17.5 degrees. The highest normal-force coefficient obtained with the flaps up and the slats unlocked was 1.46 at an angle of attack of 36 degrees, and in the angleof-attack range from 23 degrees to 30 degrees the normalforce coefficient had a substantially constant value of 1.32. At the lower angles of attack with the slats locked or unlocked deflecting the flaps produced an increment in normal-force coefficient at a given angle of attack of about 0.26. The highest normal-force coefficient obtained with the flaps down and the slats locked or unlocked was about 1.65. This value was attained at an angle of attack of about 35.5 degrees with the slats locked and at an angle of attack of about 38 degrees with the slats unlocked. However, in the angle-of-attack range from 12 degrees to 32 degrees considerably greater normalforce coefficients were obtained with the slats unlocked than with the slats locked. 38. Carner, H. Arthur; and Knapp, Ronald J.: Flight Measurements of the Pressure Distribution on the Wing of the X-1 Airplane (10-Percent-Thick Wing) Over a Chordwise Station Near the Midspan, in Level Flight at Mach Numbers from 0.79 to 1.00 and in a Pull-Up at a Mach Number of 0.96. NACA RM L50H04, September 12, 1950, 86H89804. Measurements of the chordwise pressure distribution over the 10-percent-thick wing of the X-1 research airplane have been made at a section near the midspan of the left wing. Data presented are for a Mach number range from 0.79 to 1.00 at a section normal-force coefficient of about 0.32 and for section normal-force coefficients up to 1.00 at a Mach number of approximately 0.96. The results show that the section center of load moves aft from about 32 percent chord at Mach number 0.79 to 40 percent chord at Mach number 0.84, and then forward to 18 percent chord at Mach number 0.89. The section center of load moves aft to 45 percent chord at Mach number 0.95 and then remains approximately constant at Mach numbers up to 1.00. At a section normal-force coefficient of 0.32 a shock exists on the upper surface at the lowest test Mach number of 0.79 and supersonic flow exists over approximately 50 percent of the chord on the upper 23 surface. The first indication of a shock on the lower surface occurs at a Mach number of about 0.84. At Mach numbers above 0.95 the shocks on both surfaces occur near the trailing edge and the pressure distribution over both surfaces is quite similar. An increase in the normal-force coefficient at a Mach number of approximately 0.96 causes a slight increase in the section stability at the higher normal-force coefficients. 39. Carner, H. Arthur; and Payne, Mary M.: Tabulated Pressure Coefficients and Aerodynamic Characteristics Measured on the Wing of the Bell X-1 Airplane in Level Flight at Mach Numbers from 0.79 to 1.00 and in a PullUp at Mach Number of 0.96. NACA RM L50H25, September 18, 1950, 86H91513. Tabulated pressure coefficients and aerodynamic characteristics are presented for six spanwise stations on the left wing of the Bell X-1 research airplane. The data were obtained in level flight at Mach numbers from 0.79 to 1.00 and in a pull-up to an airplane normal-force coefficient of 0.91 at a Mach number of approximately 0.96. 40. Drake, Hubert M.; and Carden, John R.: ElevatorStabilizer Effectiveness and Trim of the X-1 Airplane to a Mach Number of 1.06. NACA RM L50G20, November 1, 1950, 87H24582. The relative elevator-stabilizer effectiveness of the X-1 has been determined to decrease from a value of 0.25 at a Mach number of 0.78 to a value of 0.05 at a Mach number of 1.0. At supersonic speeds the effectiveness increases. The various stabilizer settings are caused by the variation in effectiveness and the fact that the effectiveness is nonlinear at Mach numbers between 0.94 and 0.97. It was found that, with the elevator fixed at zero, only about 0.5 degrees of stabilizer movement would be required to trim through the Mach number range from 0.78 to 1.02. 41. Knapp, Ronald J.; and Wilken, Gertrude V.: Tabulated Pressure Coefficients and Aerodynamic Characteristics Measured on the Wing of the Bell X-1 Airplane in Pull-Ups at Mach Numbers From 0.53 to 0.99. NACA RM L50H28, November 1, 1950, 86H93182. Tabulated pressure coefficients and aerodynamic characteristics are presented for six spanwise stations on the left wing of the Bell X-l research airplane. The data were obtained in 10 pull-ups at Mach numbers from 0.53 to 0.99. 42. Angle, Ellwyn; and Holleman, Euclid C.: Determination of Longitudinal Stability of the Bell X-1 Airplane From Transient Responses at Mach Numbers Up to 1.12 at Lift Coefficients of 0.3 and 0.6. NACA RM L50I06A, November 7, 1950, 86H91094. A number of free-flight transient responses resulting from small stabilizer movements were obtained during flight tests of the Bell X-1 airplane (8-percent-thick wing and 6-percentthick tail). Responses were analyzed to obtain a measure of the longitudinal stability characteristics of the airplane over the Mach number range from 0.72 to 1.12 at lift coefficients of 0.3 and 0.6. The data presented indicate three significant features: (1) The damping varies greatly with Mach number, maximum damping occurring at Mach numbers of 0.82 and 1.08 and a minimum damping at about 0.93; (2) some uncertainty of damping between Mach numbers of 0.91 to 0.95 appears although good agreement with model tests exists throughout the Mach number range covered; and (3) the static stability of the airplane increases with Mach number to a Mach number of about 0.93 and decreases with further increasing Mach number. Data above a Mach number of 0.90 indicate some lift-coefficient effects. Agreement of the fullscale flight data and model data over the Mach number range is good. 43. Drake, Hubert M.: Effects on the Lateral Oscillation of Fixing the Rudder and Reflexing the Flaps on the Bell X-1 Airplane. NACA RM L50I05, December 11, 1950, 86H93173. Flight tests have been made on the Bell X-1 airplane having the 10-percent-thick wing and the 8-percent-thick tail to evaluate the effects of fixing the rudder and changing the inclination of the principal axes of inertia by reflexing the landing flaps on the snaking which has been encountered over practically the entire range of Mach number and normal-force coefficient. The data were obtained during power-off glides at altitudes between 32,000 and 16,000 feet. The results showed that fixing the rudder reduced the amplitude of snaking, but did not eliminate it at a Mach number of 0.84. It was also found that reflexing the flaps to change the inclination of the principal axis of inertia 1 and 3/4 nose up increased the dynamic lateral stability, but had only a small effect on the snaking oscillation at a Mach number of 0.85. 44. Sadoff, Melvin; and Sisk, Thomas R.: Summary Report of Results Obtained During Demonstration Tests of the Northrop X-4 Airplanes. NACA RM A50I01, December 13, 1950, 86H93149. Results obtained during the demonstration flight tests of the Northrop X-4 No. 1 and No. 2 airplanes are presented. Information is included on the static and dynamic longitudinal- and lateral-stability characteristics, the stalling characteristics, and the buffet boundary. The data indicated that the airplane was almost neutrally stable in straight flight at low Mach numbers with the center of gravity located at about 21.4 percent of the mean aerodynamic chord for the clean configuration. In accelerated flight over a Mach number range of about 0.44 to 0.84 the airplane was longitudinally stable up to a normal-force coefficient of about 0.4. At higher values of normal-force coefficient and at the higher (approximately Mach 0.8) Mach numbers a longitudinal instability was experienced. The X-4 airplane does not satisfy the Air Force specifications for damping of the short-period longitudinal oscillation. The pilot, however, did not consider the low damping characteristics of the airplane objectionable 24 for small disturbances. An objectionable undamped oscillation about all three axes was experienced, however, at the highest test Mach number of 0.88. Theory predicted the period of the short-period longitudinal oscillation fairly well, while, in general, the theoretical damping indicated a higher degree of stability than was actually experienced. This discrepancy was traced to a considerable error in the estimation of the rotational damping factor. The directional stability of the X-4 airplane as measured in steady sideslips was high and essentially constant over the speed range covered, while the dihedral effect decreased considerably with an increase in airspeed. The damping of the lateral oscillation does not meet the Air Force requirements for satisfactory handling qualities over the Mach number range covered. The data indicated decreased damping as the flight Mach number was increased above about 0.5, and at high Mach numbers (M>0.8) and at high altitudes the X-4, in common with other transonic research airplanes, experienced a small amplitude undamped lateral oscillation. The dynamic lateral-stability characteristics were estimated fairly well by theory at low Mach numbers and at a pressure altitude of 10,000 feet. At 30,000 feet, however, at Mach numbers above about 0.6, the theory again indicated a higher degree of stability than was actually obtained. For the conditions covered in these tests the stalling characteristics of the X-4 airplane, as measured in stall approaches in straight flight and in an accelerated stall to about 1.6g, were, in general, satisfactory. Both the stall approaches and the stall were characterized by a roll-off to the right. The X-4 buffet boundary showed a sharp drop-off in the normal-force coefficient for the onset of buffeting as the flight Mach number exceeded 0.8. The boundary was almost identical to that obtained for the D-558-II research airplane at comparable Mach numbers. 45. Keener, Earl R.; and Pierce, Mary: Tabulated Pressure Coefficients and Aerodynamic Characteristics Measured in Flight on the Wing of the Douglas D-558-I Airplane for a 1-G Stall, a Speed Run to a Mach Number of 0.90, and a Wind-Up Turn at a Mach Number of 0.86. NACA RM L50J10, December 15, 1950, 86H91866. Tabulated pressure coefficients and aerodynamic characteristics are presented unanalyzed for six spanwise stations on the right wing of the Douglas D-558-I research airplane (BuAero No. 37972). The data were obtained in a 1 g stall at subcritical Mach numbers, in a speed run to a Mach number of 0.90 and in a wind-up turn at a Mach number of 0.86. 46. Mayer, John P.; Valentine, George M.; and Swanson, Beverly J.: Flight Measurements With the Douglas D-558-II (BuAero No. 37974) Research Airplane: Measurements of Wing Loads at Mach Numbers Up to 0.87. NACA RM L50H16, December 26, 1950, 86H91869. Measurements of wing loads at Mach numbers up to 0.87. 1951 Technical Publications 47. Drake, Hubert M.; and Clagett, Harry P.: Effects on the Snaking Oscillation of the Bell X-1 Airplane of a Trailing-Edge Bulb on the Rudder. NACA RM L50K01A, January 16, 1951, 86H94153. A rudder bulb was installed on the trailing edge of the rudder of the Bell X-1 airplane having the 8-percent-thick wing and 6-percent-thick tail. Several flights were made to investigate the effects of the bulb on the snaking oscillation at Mach numbers between 0.75 and 1.0. It was found that the rudder bulb had no noticeable effect on the snaking oscillation over the Mach number range tested. 48. Sjoberg, S. A.; Peele, James R.; and Griffith, John H.: Flight Measurements With the Douglas D-558-II (BuAero No. 37974) Research Airplane: Static Longitudinal Stability and Control Characteristics at Mach Numbers Up to 0.87. NACA RM L50K13, January 17, 1951, 86H94865. The paper presents the results of flight measurements of the longitudinal stability and control characteristics of the Douglas D-558-II research airplane. Data are presented in the speed range from the stalling speed of the airplane up to a maximum Mach number of 0.87. 49. Mayer, John P.; and Valentine, George M.: Flight Measurements With the Douglas D-558-II (BuAero No. 37974) Research Airplane: Measurements of the Distribution of the Aerodynamic Load Among the Wing, Fuselage, and Horizontal Tail at Mach Numbers Up to 0.87. NACA RM L50J13, January 19, 1951, 86H91625, 93R15887. Flight measurements of the aerodynamic wing and tail loads have been made on the Douglas D-558-II airplane from which the distribution of the aerodynamic load among the wing, fuselage, and horizontal tail has been determined at Mach numbers up to 0.87. These measurements indicate that, for normal-force coefficients less than 0.7, the distribution of air load among the airplane components does not change appreciably with Mach number at Mach numbers up to 0.87. The measurements also indicate that, for all flight configurations, the increase in airplane normal-force coefficient above the angle of attack at which the wing reaches its maximum normal-force coefficient is due principally to the contribution of the fuselage to the airplane normal-force coefficient. 50. Keener, Earl R.; Peele, James R.; and Woodbridge, Julia B.: Tabulated Pressure Coefficients and Aerodynamic Characteristics Measured in Flight on the Wing of the Douglas D-558-I Airplane Throughout the Normal-Force-Coefficient Range at Mach Numbers of 0.67, 0.74, 0.78, and 0.82. NACA RM L50L12A, January 29, 1951, 86H92973. 25 Tabulated pressure coefficients and aerodynamic characteristics measured in flight are presented for six spanwise stations on the right wing of the D-558-I research airplane (BuAero No. 37972). The data were obtained throughout the normal-force-coefficient range at Mach numbers of 0.67, 0.74, 0.78, and 0.82. This paper supplements similar tabulated data which have been presented in NACA RM L50J10. (See also 45.) 51. Carman, L. Robert; and Carden, John R.: Lift and Drag Coefficients for the Bell X-1 Airplane (8-PercentThick Wing) in Power-Off Transonic Flight. NACA RM L51E08, June 1951, 87H24264, 86H96412. Drag coefficients have been determined by the accelerometer method for the Bell X-1 airplane with 8-percent-thick wing and 6-percent-thick tail in power-off flight over a Mach number range of 0.64 to 1.14 and at lift coefficients from 0.1 to 1.2. 52. Stillwell, W. H.; and Wilmerding, J. V.: Flight Measurements With the Douglas D-558-II (BuAero No. 37974) Research Airplane: Dynamic Lateral Stability. NACA RM L51C23, June 18, 1951, 87H24199, 86H96642. The paper presents flight measurements of the dynamic lateral stability of the D-558-II (BuAero No. 37974) research airplane. Data are presented for a range of calibrated airspeed from 167 miles per hour to 474 miles per hour. 53. Smith, Lawrence A.: Tabulated Pressure Coefficients and Aerodynamic Characteristics Measured on the Wing of the Bell X-1 Airplane in an Unaccelerated Stall and in Pull-Ups at Mach Numbers of 0.74, 0.75, 0.94, and 0.97. NACA RM L51B23, June 19, 1951, 87H24185. Presents tabulated pressure coefficients and aerodynamic characteristics measured on the wing of the Bell X-1 research airplane in an unaccelerated stall and in pull-ups at Mach numbers of approximately 0.74, 0.75, 0.94, and 0.97. 54. Sadoff, Melvin; Roden, William S.; and Eggleston, John M.: Flight Investigation of the Longitudinal Stability and Control Characteristics of the Douglas D-558-I Airplane (BuAero No. 37972) at Mach Numbers Up to 0.89. NACA RM L51D18, June 1951, 87H24220, 86H97408. Results and analysis pertaining to the longitudinal stability and control characteristics of the Douglas D-558-I airplane (BuAero No. 37972) are presented. The results indicated that large and rapid changes in elevator deflection and force were required for balance above a Mach number of 0.84. Analysis indicated that a major part of these changes were due to a loss in elevator effectiveness. A large increase in the apparent stick-fixed stability parameter d delta (sub) e/dC (sub) N was also noted due to a loss in elevator effectiveness combined with an increase in airplane stability. Presents tabulated pressure coefficients and aerodynamic characteristics measured on the wing of the Bell X-1 research airplane in an unaccelerated low-speed stall, in push-overs at Mach numbers of 0.83 and 0.99, and in a pull-up at a Mach number of 1.16. 58. Drake, Hubert M.; Carden, John R.; and Clagett, Harry P.: Analysis of Longitudinal Stability and Trim of the Bell X-1 Airplane at a Lift Coefficient of 0.3 to Mach Numbers Near 1.05. NACA RM L51H01, October 1951, 87H24599. An analysis has been made of the flight test data obtained with two X-1 airplane shaving 10-percent-thick wings and an 8-percent-thick tail and 8-percent-thick wings and a 6-percent-thick tail. The variation with Mach number of the rate of change of downwash angle of attack, the static stability, and the airplane trim were obtained at a lift coefficient of 0.3. 980009 D-558-I Airplane, Three-View Drawing 55. Keener, Earl R.; and Bandish, Rozalia M.: Tabulated Pressure Coefficients and Aerodynamic Characteristics Measured in Flight on the Wing of the D-558-I Research Airplane Through a Mach Number Range of 0.80 to 0.89 and Throughout the Normal-ForceCoefficient Range at Mach Numbers of 0.61, 0.70, 0.855, and 0.88. NACA RM L51F12, August 1951, 87H24327, 86H98165. Presents tabulated pressure coefficients and aerodynamic characteristics obtained in flight from pressure distributions over six chordwise rows of orifices on a wing of the Douglas D-558-I research airplane (BuAero No. 37972). It includes data obtained throughout a Mach number range of 0.80 to 0.89 and throughout the normal-force-coefficient range at M = 0.61, 0.70, 0.855, and 0.88. 56. Beeler, De E.; Bellman, Donald R.; and Griffith, John H.: Flight Determination of the Effects of Wing Vortex Generators on the Aerodynamic Characteristics of the Douglas D-558-I Airplane. NACA RM L51A23, August 1951, 87H24172, 86H93919. Tests were made to determine the effects of wing vortex generators on the handling and buffeting characteristics of the Douglas D-558-I airplane. Measurements of the chordwise pressure distribution over one section of the wing, the totalhead losses in a portion of the wing wake, the total airplane drag, and the buffeting and handling characteristics were made with the basic configuration and with vortex generators of an arbitrary size, shape, and location installed on the wing. 57. Knapp, Ronald J.: Tabulated Pressure Coefficients and Aerodynamic Characteristics Measured on the Wing of the Bell X-1 Airplane in an Unaccelerated Low-Speed Stall, in Push-Overs at Mach Numbers of 0.83 and 0.99, and in a Pull-Up at a Mach Number of 1.16. NACA RM L51F25, September 1951, 87H24339. 26 59. Sadoff, Melvin; Ankenbruck, Herman O.; and O’Hare, William: Stability and Control Measurements Obtained During USAF-NACA Cooperative Flight-Test Program on the X-4 Airplane (USAF No. 46-677). NACA RM A51H09, October 26, 1951, 93R16101. Results obtained during the Air Force testing of the Northrop X-4 airplane are presented. Information is included on the stalling characteristics, the static and dynamic longitudinaland lateral-stability characteristics, and the lateral-control characteristics. The data indicated that the stalling characteristics of the X-4 airplane in straight flight and in accelerated flight at low Mach numbers were satisfactory, but that at Mach numbers above 0.68, the airplane became longitudinally unstable at moderate lift coefficients. 980042 X-4 Airplane, Three-View Drawing 60. Williams, W. C.; and Crossfield, A. S.: Handling Qualities of High-Speed Airplanes. NACA Conference on High-Speed Airplane Aerodynamics, Langley Field, Virginia, December 4–5, 1951, pp. 171–188. This paper discusses the handling qualities and stability of the X-1, D-558-I, D-558-II, X-4, F-86A, and XF-92 airplanes. 61. Drake, H. M.; and Stillwell, W. H.: Landing Experience With Transonic Research Airplanes. NACA Conference on High-Speed Airplane Aerodynamics, Langley Field, Virginia, December 4–5, 1951, pp. 269–280. This paper discusses landing experiences with the D-558-I and the X-4 airplanes. 62. *Soule, Hartley A.; and Beeler, De. E.: Review of High-Speed Buffeting Problems. NACA Conference on High-Speed Airplane Aerodynamics, Langley Virginia., December 4–5, 1951, pp. 327–340. *NACA - Langley Aeronautical Laboratory. Field, This paper discusses buffeting results from the X-1, D-558-II, X-4, and XF-92A airplanes. 1952 Technical Publications 63. Williams, W. C.; and Crossfield, A. S.: Handling Qualities of High-Speed Airplanes. NACA RM L52A08, January 1952, 87H24815. Because there have been such drastic changes in the speed range and the configuration of airplanes in the past decade it becomes necessary to re-examine the requirements for satisfactory handling qualities as proposed by Gilruth in E-842 NACA High Speed Flight Station Research Aircraft, circa 1952 (clockwise from front center), Northrop X-4, Douglas D-558-I, Douglas D-558-II, Convair XF-92A, Bell X-5, and Bell X-1-2 27 1940. This paper does not attempt to describe completely the handling qualities of all the research airplanes but does attempt to describe the objectionable characteristics and those which indicate review of the requirements. The research airplanes discussed are the X-1, D-558-1, D-558-2, X-4, F-86A, and the XF-92A. Horizontal-tail load measurements were made during the Bell acceptance tests of a transonic speed research airplane having wings variable in flight between 20 degrees and 60 degrees sweepback. Load measurements were made during sweep changes in level flight from Mach numbers of 0.5 to 0.85, and during pull-ups at a Mach number of 0.83 at sweep angles of 20 degrees, 45 degrees, and 59 degrees. E-1326 F-86 Sabre Airplane 64. Angle, Ellwyn E.; and Holleman, Euclid C.: Longitudinal Frequency-Response Characteristics of the Douglas D-558-I Airplane as Determined From Experimental Transient-Response Histories to a Mach Number of 0.90. NACA RM L51K28, February 1952, 87H24557. Transient responses from elevator pulses of the Douglas D-558-I research airplane are analyzed by the Fourier transform to give the longitudinal frequency response of the airplane to a Mach number of 0.90 at altitudes between 30,000 and 37,000 feet. 65. Huss, Carl R.; Andrews, William H.; and Hamer, Harold A.: Time-History Data of Maneuvers Performed by a McDonnell F2H-2 Airplane During Squadron Operational Training. NACA RM L52B29, May 1952, 87H24976. Preliminary results of 276 maneuvers of all types performed by an F2H-2 jet fighter airplane during normal operational training are presented in time history form and are summarized as plots of load factors and angular accelerations against indicated airspeed. 66. Rogers, John T.; and Dunn, Angel H.: Preliminary Results of Horizontal-Tail Load Measurements of the Bell X-5 Research Airplane. NACA RM L52G14, August 15, 1952, 87H24757. 28 E-810 X-5 Airplane 67. Holleman, Euclid C.: Longitudinal FrequencyResponse and Stability Characteristics of the Douglas D-558-II Airplane as Determined From Transient Response to a Mach Number of 0.96. NACA RM L52E02, September 1952, 87H25163. By an application of the Fourier transformation to transientflight data the longitudinal frequency response of the Douglas D-558-II airplane has been determined over a Mach number range of 0.62 to 0.96 at altitudes between 21,000 and 43,000 feet; however, the results have been reduced to airplane stability derivatives which are presented as functions of Mach number. 68. Sisk, Thomas R.: Flight Investigation of the Aileron Effectiveness of the Republic XF-91 Airplane Over a Mach Number Range From 0.40. NACA RM L52E07A, September 1952, 87H25188. A flight investigation has been conducted to determine the aileron effectiveness of the Republic XF-91 airplane. The tests were conducted over a Mach number range from 0.40 to 0.91 at approximate altitudes of 13,000, 24,000, and 32,000 feet. temperatures that were obtained for the positions measured were not great enough to cause loss in structural strength. 1953 Technical Publications 71. Knapp, Ronald J.; and Johnson, Wallace E.: Flight Measurements of Pressures on Base and Rear Part of Fuselage of the Bell X-1 Research Airplane at Transonic Speeds, Including Power Effects. NACA RM L52L01, January 1953, 87H24836. Flight measurements of the pressure distribution over the base and rear portion of fuselage of the Bell X-1 rocketpropelled airplane at transonic speeds, including power effects, are presented. 72. Childs, Joan M.: Flight Measurements of the Stability Characteristics of the Bell X-5 Research Airplane in Sideslips at 59-Degree Sweepback. NACA RM L52K13B, February 1953, 87H24781. Flight measurements of the stability characteristics of the Bell X-5 research airplane were made in steady sideslips at 59 degree sweepback at Mach numbers from 0.62 to 0.97 at altitudes varying from 35,000 to 40,000 feet. 73. Finch, Thomas W.; and Briggs, Donald W.: Preliminary Results of Stability and Control Investigation of the Bell X-5 Research Airplane. NACA RM L52K18B, February 1953, 87H24804. Results obtained during the acceptance tests of the X-5 airplane are presented. Information on the stalling characteristics, static longitudinal stability characteristics, and lateral control characteristics at various sweep angles is included. 74. Finch, Thomas W.; and Walker, Joseph A.: Flight Determination of the Static Longitudinal Stability Boundaries of the Bell X-5 Research Airplane With 59-Degree Sweepback. NACA RM L53A09B, February 1953, 87H24925. Results obtained during flights of the Bell X-5 airplane with 59 degree sweepback are presented showing the variation with Mach number of the normal-force coefficient at the longitudinal reduction in stability. Results are given for a Mach number range of 0.67 to 0.98. 75. Bellman, Donald R.: Lift and Drag Characteristics of the Bell X-5 Research Airplane at 59-Degree Sweepback for Mach Numbers From 0.60 to 1.03. NACA RM L53A09C, February 1953, 87H24926. Lift and drag coefficients for the 59-degree sweptback configuration of the Bell X-5 airplane were determined from flight tests covering the Mach number range 0.60 to 1.03. A 29 E-475 XF-91 Airplane 69. Day, Richard E.; and Stillwell, Wendell H.: First Landing of Bell X-2 Research Airplane. NACA RM L52I11, October 1, 1952, 93R16608. The Bell X-2 supersonic research airplane is equipped with a skid main landing gear and a nose wheel. Pending completion of the rocket engine, glide flights are being performed to determine low-speed handling qualities of the airplane and the landing characteristics with the ski type landing gear. The present paper presents data obtained during the approach and landing of X-2 airplane on its first flight. E-2822 X-2 Airplane 70. Jones, Ira P. Jr.: Measurements of Aerodynamic Heating Obtained During Demonstration Flight Tests of the Douglas D-558-II Airplane. NACA RM L52I26A, November 1952, 87H24932. Measurements of transient skin and canopy-glass temperature and stagnation temperature were made on the Douglas D-558-II research airplane up to a Mach number of 1.89 and to an altitude of about 77,000 feet. The maximum brief comparison is made between the 20-degree and 59-degree sweptback configuration. 76. Ankenbruck, Herman O.; and Dahlen, Theodore E.: Some Measurements of Flying Qualities of a Douglas D-558-II Research Airplane During Flights to Supersonic Speeds. NACA RM L53A06, March 1953, 87H24908. Results of measurements of lateral and longitudinal flying qualities of the Douglas D-558-II research airplane in flight to a Mach number of 1.87 and an altitude of 67,000 feet are presented. 77. Sisk, Thomas R.; and Mooney, John M.: Preliminary Measurements of Static Longitudinal Stability and Trim for the XF-92A Delta-Wing Research Airplane in Subsonic and Transonic Flight. NACA RM L53B06, March 1953, 87H24972. Preliminary longitudinal-trim and static longitudinal stability measurements were made on the XF-92A delta-wing research airplane from Mach number of 0.18 to 0.97 at altitudes from 11,000 to 40,000 feet. coefficient at which the stability decreases and pitch-up starts have been determined for both airplane configurations at Mach numbers up to about 0.94. 79. Dahlen, Theodore E.: Maximum Altitude and Maximum Mach Number Obtained With the Modified Douglas D-558-II Research Airplane During Demonstration Flights. NACA RM L53B24, April 1953, 87H24989. The maximum values of Mach number and altitude obtained with the Douglas D-558-II all-rocket research airplane and determined by the radar-phototheodolite method are presented in this paper. 80. Baker, Thomas F.: Some Measurements of the Buffet Region of a Swept-Wing Research Airplane During Flights to Supersonic Mach Numbers. NACA RM L53D06, May 1953, 87H25036. Limited measurements have been made of the region in which buffeting has been experienced by the swept-wing Douglas D-558-II research airplane during flights to supersonic Mach numbers. Buffet intensities and frequencies are given. 81. Johnson, H. I.: The Background of Flying or Handling Qualities. Presented to the Flight Test Panel of AGARD, May 1953, 65N88424. 82. Holleman, Euclid C.; Evans, John H.; and Triplett, William C.: Preliminary Flight Measurements of the Dynamic Longitudinal Stability Characteristics of the Convair XF-92A Delta-Wing Airplane. NACA RM L53E14, June 1953, 87H26860. Longitudinal airplane oscillations obtained during U.S. Air Force performance tests of the Convair XF-92A airplane have been analyzed to give limited static stability and damping measurements for a Mach number range of 0.59 to 0.94. 83. Knapp, Ronald J.; and Jordan, Gareth H.: FlightDetermined Pressure Distributions Over the Wing of the Bell X-1 Research Airplane (10-Percent-Thick Wing) at Subsonic and Transonic Speeds. NACA RM L53D20, June 1953, 87H25070. Aerodynamic section characteristics for various span locations, as determined by pressure distribution measurements in flight to high lift at Mach numbers between 0.30 and 1.19, are presented for the 10-percent-thick wing of the Bell X-1 research airplane. 84. Drake, Hubert M.; and McKay, John B.: Aileron and Elevator Hinge Moments of the Bell X-1 Airplane Measured in Transonic Flight. NACA RM L53E04, June 1953, 87H26857. 30 E-953 XF-92A Airplane 78. Fischel, Jack; and Nugent, Jack: Flight Determination of the Longitudinal Stability in Accelerated Maneuvers at Transonic Speeds for the Douglas D-558-II Research Airplane Including the Effects of an Outboard Wing Fence. NACA RM L53A16, March 1953, 87H24937. The results of transonic flight measurements of the longitudinal stability characteristics of the Douglas D-558-II research the wings are presented. The levels of normal-force Hinge moments have been measured on the aileron and elevator of the Bell X-1 airplane having the 10-percent-thick wing and 8-percent-thick tail. The aileron measurements were made by means of strain gages and pressure distributions while the elevator measurements were made by means of the wheel-force strain gases. The elevator hingemoment characteristics were determined to a Mach number of 1.18 and the aileron hinge moments to a Mach number of 1.13. 85. Drake, Hubert M.; Robinson, Glenn H.; and Kuhl, Albert E.: Loads Experienced in Flights of Two SweptWing Research Airplanes in the Angle-of-Attack Range or Reduced Stability. NACA RM L53D16, June 1953, 87H25066. Loads imposed upon the Bell X-5 and Douglas D-558-II swept-wing research airplanes during flights in which reductions of longitudinal stability followed attempted moderate-lift maneuvers are discussed. Information regarding horizontal- and vertical-tail loads, wing loads, and normal-force coefficients obtained during the high-angle-ofattack, unstable portions of the flights are presented. 86. Day, Richard E.: Measurements Obtained During the Glide-Flight Program of the Bell X-2 Research Airplane. NACA RM L53G03a, July 30, 1953, 93R17091. Results obtained during the glide-flight program of the Bell X-2 research airplane are presented. Landing characteristics and limited data evaluating static longitudinal stability at low speeds are included. 87. Martin, James A.: Longitudinal Flight Characteristics of the Bell X-5 Research Airplane at 59-Degree Sweepback With Modified Wing Roots. NACA RM L53E28, August 10, 1953, 87H26865, 93R16939. In an attempt to improve the longitudinal stability characteristics of the Bell X-5 research airplane at 59-degrees sweepback the wing-root leading edge was modified, the original 52.5-degrees sweptback leading-edge fillets being replaced by rounded leading-edge fillets. The two fillet configurations are compared in this paper on the basis of results obtained from maneuvers into the region of reduced stability at pressure altitudes from 28,000 to 40,000 feet and in the Mach number range from that for stall approach in the clean condition to a Mach number of 0.97. 88. Baker, Thomas F.: Results of Measurements of Maximum Lift and Buffeting Intensities Obtained During Flight Investigation of the Northrop X-4 Research Airplane. NACA RM L53G06, August 1953, 87H26873. The variation of the intensity of buffeting experienced throughout the operational region of the semitailless Northrop X-4 airplane and the values of maximum and peak normal-force coefficients in the Mach number range from 0.42 to 0.92 have been determined. The results are compared 31 with data obtained with the swept-wing Douglas D-558-II airplane. 89. Peele, James R.: Transonic Flight Measurements of the Aerodynamic Load on the Extended Slat of the Douglas D-558-II Research Airplane. NACA RM L53F29, August 1953, 87H24933. The results of transonic flight measurements of the aerodynamic load encountered over a partial-span leadingedge slat on the 35-degrees sweptback wing of the Douglas D-558-II research airplane are presented for the lift range at Mach numbers from 0.45 to 0.98. 90. Saltzman, Edwin J.: Flight Measurements of Lift and Drag for the Bell X-1 Research Airplane Having a 10-Percent-Thick Wing. NACA RM L53F08, September 1953, 87H24842. Lift and drag results have been obtained from power-off flight tests of the Bell X-1 (10-percent-thick wing) airplane for Mach numbers 0.68 to 1.01. Comparisons of drag are made with 8-percent-thick-wing flight tests and 10-percent-thickwing wind tunnel results. 91. Rogers, John T.: Horizontal-Tail Load Measurements at Transonic Speeds of the Bell X-1 Research Airplane. NACA RM L53F30, September 1953, 87H26871. Flight measurements of aerodynamic tail loads have been made on the Bell X-1 research airplane from which the balancing tail loads, the static-longitudinal stability parameter (dC sub M/dC sub L) WF, wing-fuselage aerodynamic center, and zero-lift wing-fuselage combination pitching moment coefficient have been determined from a Mach number of 0.7 to a Mach number greater than 1.0. A comparison was made of measured tail loads and loads calculated using available wing-tunnel data. 92. Baker, Thomas F.: Some Measurements of Buffeting Encountered by a Douglas D-558-II Research Airplane in the Mach Number Range From 0.5 to 0.95. NACA RM L53I17, November 1953, 87H24978. Measurements of the intensity of buffeting were made with a swept-wing Douglas D-558-II research airplane at subsonic speeds. The variation of buffet intensity with lift, angle of attack, and Mach number is presented. The results are compared with similar measurements made with another Douglas D-558-II airplane at high subsonic and supersonic speeds. 93. Knapp, Ronald J.; and Jordan, Gareth H.: Wing Loads on the Bell X-1 Research Airplane (10-PercentThick Wing) as Determined by Pressure-Distribution Measurements in Flight at Subsonic and Transonic Speeds. NACA RM L53G14, November 1953, 87H26874. The wing loads (including some wing-to-fuselage load carry over) and some aerodynamic characteristics as determined by pressure-distribution measurements in flight to high lift at Mach numbers from 0.50 to 1.19 for the 10-percent-thick wing of the Bell X-1 research airplane are presented. 94. Knapp, Ronald J.; Jordan, Gareth H.; and Johnson, Wallace E.: Fuselage Pressures Measured on the Bell X-1 Research Airplane in Transonic Flight. NACA RM L53I15, November 1953, 87H24970. Fuselage pressure distributions and integrated normal load on the fuselage of the Bell X-1 research airplane (10-percent-thick wing) in flight during pull-ups to near maximum lift at Mach numbers of about 0.78, 0.85, 0.88, and 1.02 are presented. 98. Fischel, Jack: Effect of Wing Slats and Inboard Wing Fences on the Longitudinal Stability Characteristics of the Douglas D-558-II Research Airplane in Accelerated Maneuvers at Subsonic and Transonic Speeds. NACA RM L53L16, February 1954, 87H25262. The results of subsonic and transonic flight measurements of the longitudinal stability characteristics of the Douglas D-558-II research airplane for several wing-slat and inboard wing-fence configurations are presented at Mach numbers up to 1.0. The improvement provided by fully extended slats, compared to the stability characteristics of the slats-retracted configuration, is shown; and the effects of inboard wing fences on the stability characteristics with slats fully extended are also discussed. Limited data obtained with slats half extended indicated the similarity of this configuration to the slats-retracted configuration. The effects of a bungee (which improved the stick-free stability characteristics) in alleviating the stability changes apparent to the pilot are also discussed. 99. Peele, James R.: Flight-Determined Pressure Measurements Over the Wing of the Douglas D-558-II Research Airplane at Mach Numbers Up to 1.14. NACA RM L54A07, March 12, 1954, 93R17395. A flight investigation of the section and panel characteristics and loads obtained from pressure measurements over a 35-degrees sweptback wing at level-flight lifts has been made through the Mach number range of 0.65 to 1.14. The section pressure distributions at the root, midspan, and tip stations varied from a subsonic type of distribution at a Mach number of 0.65 to a supersonic type of distribution at Mach numbers above 1.0. 100. Crossfield, A. Scott: Subjective Experiences and Reactions During Flight Testing in the Transonic Region. 4th General Assembly of AGARD, Schevenigen, The Netherlands, May 3–7, 1954. (See excerpts in Aerospace Engineering Review, Vol. 13, No. 9, Sept. 1954, pp. 49 and 87.) 101. Briggs, Donald W.: Flight Determination of the Buffeting Characteristics of the Bell X-5 Research Airplane at 58.7-Degrees Sweepback. NACA RM L54C17, May 24, 1954, 93R17481. Flight measurements were made of the buffeting characteristics of the Bell X-5 research airplane at 58.7-degrees sweepback in the Mach number range from 0.65 to approximately 1.03 at altitudes from 37,000 to 43,000 feet. Maximum airplane normal-force coefficients were attained for Mach numbers up to 0.96. 102. Finch, Thomas W.: A Flight Investigation of the Effects of Inclination of the Principal Axis of Inertia on the Dynamic Lateral Stability of the Republic XF-91 Airplane. NACA RM L53I28, July 1954, 87H25044. 32 1954 Technical Publications 95. Bellman, Donald R.; and Sisk, Thomas R.: Preliminary Drag Measurements of the Consolidated Vultee XF-92A Delta-Wing Airplane in Flight Tests to a Mach Number of 1.01. NACA RM L53J23, January 1954, 87H25168. Lift and drag data for the Consolidated Vultee XF-92A deltawing airplane were obtained for Mach numbers from 0.63 to 0.90. The drag coefficients for a lift coefficient of 0.08 are extended to a Mach number of 1.01. 96. Baker, Thomas F.: Measured Data Pertaining to Buffeting at Supersonic Speeds of the Douglas D-558-II Research Airplane. NACA RM L53L10, February 1954, 87H25252. Data pertaining to buffeting have been measured at supersonic speed and high lift with the Douglas D-558-II airplane. Buffeting was encountered at normal-force coefficients greater than about 0.7 in the Mach number range from 0.96 to 1.27 but at a Mach number of 1.57, a peak normal-force coefficient of 0.80 was attained with no indication of buffeting. Buffet intensities at normal-force coefficients up to 1.5 are given for low supersonic Mach numbers. Sample records of flight in rough air at supersonic speed are included. 97. Ankenbruck, Herman O.: Determination of Longitudinal Stability in Supersonic Accelerated Maneuvers for the Douglas D-558-II Research Airplane. NACA RM L53J20, February 1954, 87H25139. Flight tests were performed with the Douglas D-558-II research airplane to investigate the longitudinal stability of the airplane in accelerated flight at supersonic speeds to a Mach number of 1.67. This paper shows the conditions where instability occurs at supersonic speeds. A flight investigation has been conducted to determine the effect of variable wing incidence angle on the dynamic lateral stability of the Republic XF-91 airplane. The tests were conducted over a Mach number range of 0.3 to 0.9 at altitudes of 10,000, 20,000, 30,000 and 37,500 feet at wing incidence angles of –2 degrees, 2 degrees, 4 degrees, and 5.65 degrees. 103. Sadoff, Melvin; and Crossfield, A. Scott: A Flight Evaluation of the Stability and Control of the X-4 SweptWing Semitailless Airplane. NACA RM H54G16, August 1954, 87H24527. A flight evaluation of the handling qualities of the swept-wing semitailless X-4 airplane was made. Static and dynamic stability and control investigation covered a speed range from stall to Mach numbers of 0.92. Typical swept-wing instability at moderate lifts was encountered. Unsatisfactory and dangerous self-excited dynamic motions occurred at the highspeed end of the range investigated. 104. Nugent, Jack: Lift and Drag Characteristics of the Douglas D-558-II Research Airplane Obtained in Exploratory Flights to a Mach Number of 2.0. NACA RM L54F03, August 4, 1954, 93R17490. A flight investigation was made of the Douglas D-558-II swept-wing airplane in the slats-retracted configuration. Lift and drag were determined for Mach numbers up to 2.0. The lift-coefficient range extended from below 0.1 to about 0.7. 105. Keener, Earl R.: Wing Pressure Distribution at Low Lift for the XF-92A Delta-Wing Airplane at Transonic Speeds. NACA RM H54H06, October 1954, 87H24530. Wing pressure distribution from dives at transonic speeds for the Convair XF-92A delta-wing airplane are presented. The data were obtained from five chordwise rows of orifices on the left wing throughout the Mach number range of 0.74 to 1.01 at an airplane normal-force coefficient of about 0.09, for which the left deflection was about 2 degrees up. 106. Fischel, Jack; and Brunn, Cyril D.: Longitudinal Stability Characteristics in Accelerated Maneuvers at Subsonic and Transonic Speeds of the Douglas D-558-II Research Airplane Equipped With a Leading-Edge Wing Chord-Extension. NACA RM H54H16, October 1954, 87H24531. On the basis of improved longitudinal stability characteristics exhibited in wind-tunnel model tests, the Douglas D-558-II research airplane was modified to include wing at Mach numbers up to about 1.0. The results of subsonic and transonic flight measurements of the longitudinal stability characteristics of the airplane are presented. The levels of normal-force coefficient at which the stick-fixed stability decays and pitch-up starts have been determined through speed range tested as have the variation of the stability 33 parameters d delta (sub) e/dC (sub) N (sub) A, dF (sub) e/dn, and dC (sub) N (sub) A/d sigma with Mach number. Comparisons of these data with comparable data for the unmodified airplane are also presented. 107. Crossfield, A. Scott: Flying Techniques With the Research Airplanes. Preprint 497, Inst. Aero. Sci., 1954, CAI-IAS International Meeting, Montreal, Canada, October 14–15, 1954 (see also in Aerospace Engineering Review, Vol. 14, No. 1, Jan. 1955, pp. 56–59). 108. Ankenbruck, Herman O.: Determination of Longitudinal Handling Qualities of the D-558-II Research Airplane at Transonic and Supersonic Speeds to a Mach Number of About 2.0. NACA RM H54G29A, November 1954, 87H24529. Flight tests were performed with the Douglas D-558-II research airplane to investigate the longitudinal handling qualities and trim characteristics at transonic and supersonic speeds. This paper describes the changes with Mach number of the lift, maneuvering, and trim characteristics with elevator and stabilizer to a Mach number of about 2.0. 109. Bellman, Donald R.; and Murphy, Edward D.: Lift and Drag Characteristics of the Douglas X-3 Research Airplane Obtained During Demonstration Flight to a Mach Number of 1.20. NACA RM H54I17, December 1954, 87H24532. Lift and drag data for the Douglas X-3 airplane were obtained during some of the demonstration flights. The data extend over the Mach number range from 0.82 to 1.20 and for certain constant Mach numbers the lift coefficient range from 0 to 1.0 is covered. A comparison of the flight data with wind-tunnel and rocket-model tests shows that the model tests satisfactorily predict the performance of the airplane. E55-01994 X-3 Airplane 110. Ankenbruck, Herman O.; and Wolowicz, Chester H.: Lateral Motions Encountered With the Douglas D-558-II All-Rocket Research Airplane During Exploratory Flights to a Mach Number of 2.0. NACA RM H54I27, December 1954, 87H24533. Flight tests were performed with the Douglas D-558-II research airplane to investigate the lateral motions obtained during exploratory flights at supersonic speeds. This paper describes the effects of Mach number and angle of attack on the lateral handling qualities during oscillations at supersonic speeds. Some calculations of period and damping are included. Also shown are the results of some measurements of the variation of rudder hinge moments with sideslip and the effects of power on the variation. 1955 Technical Publications 111. McKay, John B.: Rolling Performance of the Republic YF-84F Airplane as Measured in Flight. NACA RM H54G20A, January 1955, 87H24528. 980051 Flight measurements of the rolling performance of the Republic YF-84F airplane were made at altitudes of 10,000, 25,000, and 40,000 feet. The tests were conducted over a Mach number range from 0.35 to 0.95. XF-92A Airplane, Three-View Drawing 113. Holleman, Euclid C.; and Triplett, William C.: Flight Measurements of the Dynamic Longitudinal Stability and Frequency-Response Characteristics of the XF-92A Delta-Wing Airplane. NACA RM H54J26A, January 1955, 87H24535. Results of dynamic longitudinal flight test conducted with the XF-92A delta-wing airplane over a Mach number range of 0.42 to 0.94 at an altitude of about 30,000 feet are presented. The data were analyzed by measuring the airplane oscillatory characteristics, by matching the airplane system with an analog computer, and by determining the frequency-response characteristics of the airplane. Wherever possible, stability derivatives were computed and are presented as a function of Mach number. 114. Saltzman, Edwin J.: The Effect of the BluntTrailing-Edge Elevons on the Longitudinal and Lateral Handling Qualities of the X-4 Semitailless Airplane. NACA RM H54K03, January 1955, 87H24537. The effects of thickening the trailing edges of the elevons on the static longitudinal stability and control, lateral control for the X-4, a swept-wing semitailless airplane, are presented. The results of this study are compared with similar tests for the X-4 with conventional elevon trailing edges. 115. Sisk, Thomas R.; and Muhleman, Duane O.: Longitudinal Stability Characteristics in Maneuvering Flight of the Convair XF-92A Delta-Wing Airplane Including the Effects of Wing Fences. NACA RM H54J27, January 1955, 87H24536. The longitudinal maneuvering stability characteristics are evaluated on the Convair F-92A delta-wing airplane in windup turns over the Mach number range from 0.70 to 0.95 at 34 (Photo courtesy of Langley Research Center) L-84926 YF-84F Airplane 112. Johnson, Clinton T.; and Kuhl, Albert E.: Flight Measurements of Elevon Hinge Moments on the XF-92A Delta-Wing Airplane. NACA RM H54J25A, January 1955, 87H24534. Elevon hinge-moment measurements were made during flight tests of the Convair XF-92A delta-wing airplane over the Mach number range from 0.70 to 0.95. Hinge moments were measured during longitudinal elevon pulses, aileron rolls, and wind-up turns. Data are presented giving the variation of C (sub) h (sub) delta and C (sub) h (sub) alpha as determined from these tests. altitudes between 22,000 and 39,000 feet. A longitudinal stability reduction evidenced as a pitch-up encountered over the Mach number range tested is evaluated along with the airplane behavior in the region of reduced stability. Two wing fence configurations are evaluated and compared with the basic airplane configuration characteristics. 116. Stillwell, Wendell H.: Results of Measurements Made During the Approach and Landing of Seven HighSpeed Research Airplanes. NACA RM H54K24, February 1955, 87H24267. Measurements made during the approach and landing of the X-1, X-3, X-4, X-5, D-558-I, D-558-II, and XF-92A research airplanes are presented. Data are also presented for the effect of lift drag ratio on the landing characteristics of the X-4 airplane. 117. Anon.: Flight Experience With Two High-Speed Airplanes Having Violent Lateral-Longitudinal Coupling in Aileron Rolls. NACA RM H55A13, February 4, 1955, 93R17938. During flight tests of two high-speed airplane configurations, violent cross-coupled lateral and longitudinal motions were encountered following abrupt rudder-fixed aileron rolls. The speeds involved ranged from a Mach number of 0.7 to 1.05. The motions were characterized by extreme variations in angles of attack and sideslip which resulted in load factors as large as 6.7g (negative) and 7g (positive) normal acceleration and 2g transverse acceleration. 118. Robinson, Glenn H.; Cothren, George E.; and Pembo, Chris: Wing-Load Measurements at Supersonic Speeds of the Douglas D-558-II Research Airplane. NACA RM H54L27, March 1955, 87H24539. Flight measurement of the aerodynamic wing loads on the D-558-II airplane have been made in the Mach number range from 1.0 to 2.0. Results of measurements of the wing-panel normal-force, bending moment, and pitching-moment coefficients, normal-force-curve slope, lateral center of pressure, and chordwise center of pressure are presented. 119. Jordan, Gareth H.; and Keener, Earl R.: FlightDetermined Pressure Distributions Over a Section of the 35-Degree Swept Wing of the Douglas D-558-II Research Airplane at Mach Numbers Up to 2.0. NACA RM H55A03, March 1955, 87H24287. Presented are pressure distributions and section characteristics for a wing-midsemispan station perpendicular to the 30-degree common-chord line of the 35-degree sweptback wing of the Douglas D-558-II research airplane at Mach numbers from 1.17 to 2.00. 120. *Gates, O. B., Jr.; Weil, J.; and *Woodling, C. H.: Effect of Automatic Stabilization on the Sideslip and Angle of Attack Disturbances in Rolling Maneuvers. In NACA Conf. on Autom. Stability and Control of Aircraft, March 30, 1955, pp. 25–41, (see N72-73193 12-99), 72N73195. Time histories are presented that illustrate the large motions which have been encountered in flight tests of some of the present-day fighter airplanes. Results of some analog studies are discussed which indicate that variations in certain of the airplane stability derivatives could have an appreciable effect on these undesirable motions. *Langley Aeronautical Laboratory, Hampton, Virginia. 121. Cole, H. A., Jr.; Brown, S. C.; and Holleman, E. C.: Effects of Flexibility on the Longitudinal and Lateral Dynamic Response of a Large Airplane. NACA RM A55D14, NACA Conf. on Autom. Stability and Control of Aircraft, March 1955, pp. 57–70, (see N72-73193 12-99), 72N73197. In recent years the desire to increase the range and speed of large airplanes has led to sweptback wings of high aspect ratio, thin airfoils, and fuselages of high fineness ratio. The dynamic effects are especially important in the design of automatic control systems because structural modes may introduce instabilities which would not arise with a rigid airplane. It is important for the automatic-control designer to consider the effects of flexibility on control systems. 122. Jordan, Gareth H.; and Hutchins, C. Kenneth: Preliminary Flight-Determined Pressure Distribution Over the Wing of the Douglas X-3 Research Airplane at Subsonic and Transonic Mach Numbers. NACA RM H55A10, April 1955, 87H24540. Preliminary flight-measured chordwise pressure distributions have been obtained at a wing midsemispan station of the Douglas X-3 research airplane through an angle-of-attack range at Mach numbers of 0.61, 0.78, 0.94, and 1.10. The results of the investigation indicate that the maximum section normal-force coefficient increased from about 0.7 at the lower Mach numbers to about 1.2 at a Mach number of 1.10. The pressure distributions at Mach numbers of about 0.61, 0.78, and 0.94 showed good agreement with wind-tunnel results. 35 E-3995 D-558-II Airplane, Top View At Mach numbers of 0.94 and 1.10 leading-edge flap normalforce and hinge-mount coefficients increased with increase in angle of attack throughout the angle-of-attack range tested and resulted in high normal-force and hinge-moment coefficients at the higher angles of attack. The lateral stability and control characteristics of the Convair XF-92A delta-wing airplane are determined for sideslips, aileron rolls, and rudder pulses. A limited amount of data with wing fences installed at 60 percent of the wing semispan is presented for comparison with the basic airplane configuration. 126. Finch, Thomas W.: Flight Determination of the Longitudinal Stability and Control Characteristics of the Bell X-5 Research Airplane at 58.7 Degrees Sweepback. NACA RM H55C07, May 1955, 87H24301. Flight tests were performed with the Bell X-5 research airplane at 58.7 degrees sweepback to measure the longitudinal stability control characteristics from elevator and stabilizer maneuvers at 40,000 feet and elevator maneuvers at 25,000 feet and 15,000 feet. Results are presented for Mach numbers up to 1.0 and include trim characteristics, apparent stability parameters, relative control effectiveness, stick-force gradient, normal-force-curve slope, and effect of dynamic pressure and engine power. Comparison is made with windtunnel data. 127. Kuhl, Albert E.; and Johnson, Clinton T.: Flight Measurements of Wing Loads on the Convair XF-92A Delta-Wing Airplane. NACA RM H55D12, May 1955, 87H24304. Aerodynamic loads were obtained from strain-gage measurements during the NACA flight test program of the XF-92A research airplane. Wing-panel loads were measured during longitudinal pulses and wind-up turns over the mach number rang from 0.43 to 0.95. The wing-panel loads due to elevon deflection and angle of attack were determined for the Mach number range of these tests. 128. Cooney, T. V.; Andrews, William H.; and McGowan, William A.: Preliminary Results From Flight Measurements in Gradual-Turn Maneuvers of the Wing Loads and the Distribution of Load Among the Components of a Boeing B-47A Airplane. NACA RM L55B02, June 27, 1955, 93R17884. 129. Day, Richard E.; and Fischel, Jack: Stability and Control Characteristics Obtained During Demonstration of the Douglas X-3 Research Airplane. NACA RM H55E16, July 1955, 87H24543. Results obtained from flights of the manufacturer’s demonstration program and from U.S. Air Force evaluation flights of the Douglas X-3 research airplane are presented. Data evaluation includes static longitudinal directional, and lateral stability and control for Mach numbers up to 1.21 at pressure altitudes from 12,800 feet to 34,000 feet. Comparisons are made with wind-tunnel and rocket-model tests. 130. Weil, Joseph; Gates, Ordway B.; Banner, Richard D.; and Kuhl, Albert E.: Flight Experience of Inertia 36 980041 X- 3 Airplane, Three-View Drawing 123. Banner, Richard D.; Reed, Robert D.; and Marcy, William L.: Wing-Load Measurements of the Bell X-5 Research Airplane at a Sweep Angle of 58.7 Degrees. NACA RM H55A11, April 1955, 87H24541. A flight investigation has been made over an altitude and lift range to determine the wing loads of the Bell X-5 research airplane at a sweep angle of 58.7 degrees at subsonic and transonic Mach numbers. The wing loads were nonlinear over the angle-of-attack range from zero to maximum wing lift. The nonlinear trends were more pronounced at angles of attack above the “pitch-up” where there is a reduction in the wing lift and an inboard and forward movement in the center of load. No apparent effects of altitude on the wing loads were evident from the data obtained in these tests. 124. Baker, Thomas F.; and Johnson, Wallace E.: Flight Measurements at Transonic Speeds of the Buffeting Characteristics of the XF-92A Delta-Wing Research Airplane. NACA RM H54L03, April 1955, 87H24538. Measurements were made on the XF-92A delta-wing airplane of buffet-induced fluctuations in normal acceleration at the airplane center of gravity and of fluctuations in wing structural shear load in the Mach number range from 0.6 to 0.96 at altitudes from 25,000 to 38,000 feet. Airplane normal force coefficients in the order of 0.7 were attained at Mach numbers less than 0.9. Buffet frequencies and the variations with Mach number, lift, and angle of attack of buffet intensity are given. 125. Sisk, Thomas R.; and Muhleman, Duane O.: Lateral Stability and Control Characteristics of the Convair XF-92A Delta-Wing Airplane as Measured in Flight. NACA RM H55A17, May 1955, 87H24299. Coupling in Rolling Maneuvers. NACA RM H55E17B, July 1955, 87H24318. A brief discussion is presented of the flight tests of two airplanes which have exhibited strong coupling between their lateral and longitudinal motions. Results are presented which indicate the effect of directional stability and vertical tail size on the tail shear loads encountered in rolling maneuvers. 131. Reed, Robert D.: Flight Measurements of Horizontal-Tail Loads on the Bell X-5 Research Airplane at a Sweep Angle of 58.7 Degrees. NACA RM H55E20A, July 1955, 87H24544. Flight measurements of the horizontal tail loads on the Bell X-5 research airplane have been made in accelerated maneuvers at Mach numbers from 0.61 to 1.00 at an altitude of 40,000 feet. At this altitude the aerodynamic and balancing tail loads were determined at all normal-force coefficients up to near maximum lift. Comparisons were made with flight and wind-tunnel data obtained at 25,000 feet and 15,000 feet over a limited lift range. 132. Fischel, Jack; and Reisert, Donald: Effect of Several Wing Modifications on the Low Speed Stalling Characteristics of the Douglas D-558-II Research Airplane. NACA RM H55E31A, July 1955, 87H24545. The low-speed stalling and lift characteristics of the DouglasD-558-II research airplane were measured in a series of 1g stalls performed with several wing modifications designed to alleviate swept-wing instability and pitch-up. The various configurations investigated include the basic wing configurations and two wing-fence configurations in combination with retracted free-floating, or extended slats, and a wing leading-edge chord extension configuration. All configurations were investigated with flaps and landing gear retracted and extended. number and altitude capabilities of the Bell X-1A research airplane. On two flights of the X-1A airplane, one reaching a Mach number of about 2.44, the other a geometric altitude of about 90,000 feet, lateral stability difficulties were encountered which resulted in uncontrolled rolling motions of the airplane at Mach numbers near 2.0. Analysis indicates that this behavior apparently results from a combination of low directional stability and damping in roll and may be aggravated by high control friction and rocket motor misalignment. The deterioration of directional stability with increasing Mach number can lead to severe longitudinallateral coupling at low roll rates. The misalignment of the rocket motor could induce sufficiently high roll velocities to excite these coupled motions. Adequate control of these motions was virtually impossible because of the high control friction. In the absence of rolling, poor lateral behavior might be expected at somewhat higher Mach numbers because wind-tunnel data indicate neutral directional stability at about M = 2.35. E55-01799 X-1A Airplane 133. Holleman, Euclid C.: Flight Measurements of the Lateral Response Characteristics of the Convair XF-92A Delta-Wing Airplane. NACA RM H55E26, August 1955, 87H24323. Rudder pulse maneuvers were performed with the XF-92A 60 degrees delta-wing airplane at 30,000 feet over a Mach number range of 0.52 to 0.92. Tests were conducted with and without a wing fence. Representative data were analyzed to give airplane stability derivatives and frequency responses. 134. Drake, Hubert M.; and Stillwell, Wendell H.: Behavior of the Bell X-1A Research Airplane During Exploratory Flights at Mach Numbers Near 2.0 and at Extreme Altitudes. NACA RM H55G25, September 1955, 93R18135. A flight program has been conducted by the U. S. Air Force consisting of exploratory flights to determine the Mach 37 135. Drake, Hubert M.; Finch, Thomas W.; and Peele, James R.: Flight Measurements of Directional Stability to a Mach Number of 1.48 for an Airplane Tested With Three Different Vertical Tail Configurations. NACA RM H55G26, October 1955, 87H24547. Directional stability characteristics have been determined from the measured period and damping of a fighter-type airplane over the Mach number range from 0.72 to 1.48 at altitudes of 40,000 feet and 30.000 feet. Three different vertical tails of varying aspect ratio or area, or both, were employed. 136. Sisk, Thomas R.; and Andrews, William H.: Flight Experience With a Delta-Wing Airplane Having Violent Lateral-Longitudinal Coupling in Aileron Rolls. NACA RM H55H03, October 1955, 87H24548. A time-history presentation is made of aileron rolls performed by a high-speed delta-wing airplane between Mach numbers of 0.7 and 0.8 including a one-half deflection aileron roll at a Mach number of 0.75 where violent crosscoupled lateral and longitudinal motions were experienced. 137. Videan, Edward N.: Flight Measurements of the Dynamic Lateral and Longitudinal Stability of the Bell X-5 Research Airplane at 58.7 Degrees Sweepback. NACA RM H55H10, October 1955, 87H24549. Longitudinal and lateral dynamic-response characteristics to elevator and rudder pulse deflections have been measured on the Bell X-5 research airplane at 58.7 degrees sweepback. Flight records were obtained at altitudes of 40,000 feet and 25,000 feet. Period and damping, including nonlinear damping effects, are presented, and comparison is made with U.S. military lateral dynamic stability criteria. Engine gyroscopic coupling effects are discussed, and frequency response calculations are presented. 139. Williams, Walter C.: Flight Research at High Altitudes and High Speeds With Rocket-Propelled Research Airplanes. SAE Paper 601, October 1955, 87H28633. 140. Thompson, Jim Rogers; Bray, Richard S.; and Cooper, George E.: Flight Calibration of Four Airspeed Systems on a Swept-Wing Airplane at Mach Numbers Up to 1.04 by the NACA Radar-Phototheodolite Method. NACA TN 3526, November 1955, 93R13513. The calibrations of four airspeed systems installed in a North American F-86A airplane have been determined in flight at Mach numbers up to 1.04 by the NACA radarphototheodolite method. The variation of the static-pressure error per unit indicated impact pressure is presented for three systems typical of those currently in use in flight research, a nose boom and two different wing-tip booms, and for the standard service system installed in the airplane. A limited amount of information on the effect of airplane normal-force coefficient on the static-pressure error is included. The results are compared with available theory and with results from wind-tunnel tests of the airspeed heads alone. Of the systems investigated, a nose-boom installation was found to be most suitable for research use at transonic and low supersonic speeds because it provided the greatest sensitivity of the indicated Mach number to a unit change in true Mach number at very high subsonic speeds, and because it was least sensitive to changes in airplane normal-force coefficient. The static-pressure error of the nose-boom system was small and constant above a Mach number of 1.03 after passage of the fuselage bow shock wave over the airspeed head. 141. Keener, Earl R.; and Jordan, Gareth H.: Wing Pressure Distributions Over the Lift Range of the Convair XF-92A Delta-Wing Airplane at Subsonic and Transonic Speeds. NACA RM H55G07, November 1955, 87H24546. Chordwise and spanwise pressure distributions are presented for the left wing of the Convair XF-92A delta-wing airplane at Mach numbers from 0.30 to 0.93. Reynolds number based on the mean aerodynamic chord of the wing varied between 22 x 10 to the 6 power and 49 x 10 to the 6 power. The data cover the lift range from level flight to near-maximum lift. Effects of wing section stall upon the elevon-section loads are included. 980043 X-5 Airplane, Three-View Drawing 138. Johnson, Clinton T.: Flight Measurements of the Vertical-Tail Loads on the Convair XF-92A Delta-Wing Airplane. NACA RM H55H25, October 1955, 87H24551. Vertical-tail loads as obtained from strain-gage measurements during the NACA flight test program of the Convair XF-92A research airplane were measured during rudder pulses rudder-fixed oscillations, and gradually increasing sideslips over the Mach number range from 0.50 to 0.87. The vertical-tail loads resulting from rudder deflection and sideslip angle were determined for the Mach number range of the these tests. 38 1956 Technical Publications 142. Williams, Walter C.; and Phillips, William H.: Some Recent Research on the Handling Qualities of Airplanes. NACA RM H55L29A, February 1956, 87H24556. Results of recent research on the handling qualities of airplanes are reviewed. Among the subjects considered are dynamic longitudinal stability, transonic trim changes, pitchup due to decreasing airspeed, dynamic lateral stability, aileron control, and mechanical characteristics of power control systems. 143. Brunn, Cyril D.; and Stillwell, Wendell H.: Mach Number Measurements and Calibrations During Flight at High Speeds and at High Altitudes Including Data for the D-558-II Research Airplane. NACA RM H55J18, March 1956, 87H24552. This paper contains data concerning research equipment and techniques pertaining to measurements of airspeed and altitude, with particular reference to high Mach number and altitude. Computations are presented for errors in determining Mach number for altitudes of 40,000 feet to 140,000 feet. Illustrative examples are included for maximum Mach number and altitude flights of the Douglas D-558-II research airplane. number range of 0.73 to 1.39 at altitudes of 40,000 feet and 30,000 feet with three different vertical tails of varying aspect ratio or area, or both. Increasing the tail area 27 percent and the tail aspect ratio 32 percent greatly improved the rolling behavior. The adverse sideslip during the rolls decreased with increasing speeds to negligible values near Mach numbers of 1.00 to 1.05, then increased in the favorable direction at higher speeds. Stabilizer motion during the rolls greatly affected the rolling behavior. 147. Stephenson, Harriet J.: Flight Measurements of Horizontal-Tail Loads on the Douglas X-3 Research Airplane. NACA RM H56A23, April 1956, 87H24562. Horizontal-tail loads were obtained from strain-gage measurements during flight tests of the Douglas X-3 research airplane over a Mach number range of 0.65 to 1.16. The horizontal-tail-panel lift-curve slope was obtained from stabilizer pulses. Balancing-tail loads, downwash, and total airplane pitching moment were obtained from wind-up turns and pull-ups. 148. Weil, Joseph; and Day, Richard E.: An Analog Study of the Relative Importance of Various Factors Affecting Roll Coupling. NACA RM H56A06, April 1956, 87H24558. An analog study of the roll coupling problem has been made for a representative swept-wing and a tailless delta-wing configuration. The investigation, conducted primarily for subsonic flight conditions, included determination of the effects of wide variations in many of the pertinent aerodynamic derivatives on the motions developed in rolling maneuvers. The influence of large changes in principal axis inclination and mass distribution was also considered. 149. Weil, Joseph; Campbell, George S.; and Diederich, Margaret S.: An Analysis of Estimated and Experimental Transonic Downwash Characteristics as Affected by Plan Form and Thickness for Wing and Wing-Fuselage Configurations. NACA TN 3628, April 1956, 87H24920. This paper presents a summary of the effects of changes in wing plan form and thickness ratio on the downwash characteristics of wing and wing-fuselage configurations in the Mach number range between 0.6 and 1.1. Data obtained by the transonic-bump technique at two tail heights have been compared with theoretical estimations made in the subsonic and supersonic Mach number range. 150. Finch, Thomas W.; and Walker, Joseph A.: Flight Determination of the Lateral Handling Qualities of the Bell X-5 Research Airplane at 58.7 Degrees Sweepback. NACA RM H56C29, May 1956, 87H24570, 93R19517. Flight tests were performed with the Bell X-5 research airplane at 58.7 degrees sweepback to measure the lateral handling qualities over a Mach number range up to 0.97 at altitudes of 40,000, 25,000, and 15,000 feet. The dynamic 39 980010 D-558-II Airplane, Three-View Drawing 144. Weil, Joseph: A Brief Review of NACA Flight Research Relating to Roll Coupling. Symposium on Roll Coupling, Dayton, Ohio, March 1, 1956. 145. Bellman, Donald R.; and Kleinknecht, Kenneth S.: Operational Experience With Rocket Propelled Airplanes. IAS-F75-038, 11th Annual Flight Propulsion Meeting of IAS, Cleveland, Ohio, March 9, 1956, 87H28957. 146. Finch, Thomas W.; Peele, James R.; and Day, Richard E.: Flight Investigation of the Effect of Vertical Tail Size on the Rolling Behavior of a Swept-Wing Airplane Having Lateral-Longitudinal Coupling. NACA RM H55L28A, April 1956, 87H24554. The rolling behavior of a swept-wing airplane having lateral longitudinal coupling has been determined over a Mach characteristics were influenced by aerodynamic and engine gyroscopic coupling, and the damping was nonlinear at the higher Mach numbers. The positive apparent directional stability and the high apparent effective dihedral increased rapidly at higher Mach numbers. The low aileron effectiveness was adversely affected by the high effective dihedral. Directional divergence and aileron overbalance occurred at high lifts. An abrupt wing-dropping tendency existed and single-degree-of-freedom flutter occurred on the rudder at low supersonic Mach numbers. 151. Wolowicz, Chester H.: Time-Vector-Determined Lateral Derivatives of a Swept-Wing Fighter Type Airplane With Three Different Vertical Tails at Mach Numbers Between 0.70 and 1.48. NACA RM H56C20, June 1956, 87H24567. The time-vector method was used to obtain the lateral stability derivatives C (sub) Y (sub) beta, C (sub) n (sub) beta, C (sub) iota (sub) beta, C (sub) iota (sub) p, (C sub n sub r – C sub n sub beta) of a swept-wing fighter-type airplane. The airplane was tested over a Mach number range of 0.71 to 1.48 at altitudes extending from 30,000 to 43,000 feet to obtain static and dynamic lateral stability characteristics. Four configurations were employed: three different vertical tails and an extended wing. Available wind-tunnel data and theoretical calculations were used for comparison purposes. 152. Fischel, Jack; and Reisert, Donald: Effect of Several Wing Modifications on the Subsonic and Transonic Longitudinal Handling Qualities of the Douglas D-558-II Research Airplane. NACA RM H56C30, June 1956, 87H24575, 93R19515. The subsonic and transonic longitudinal handling qualities of the Douglas D-558-II research airplane were measured with several wing modifications designed to alleviate swept-wing instability and pitch-up. Airplane configurations investigated include the basic wing configuration and two wing-fence configurations in configurations in combination with retracted, free-floating, or extended slats, and a wing leadingedge chord-extension configuration. Results indicated that the comparative effects of these wing modifications on airplane pitch-up and trim-stability and on the stability parameters d delta (sub) e/dC (sub) N, dFe/da (sub) n, and C (sub) N (sub) sigma were essentially negligible. The various modifications had some measurable effect on airplane buffeting characteristics. 153. Nugent, Jack: Lift and Drag of the Bell X-5 Research Airplane in the 45-Degrees Sweptback Configuration at Transonic Speeds. NACA RM H56E02, July 1956, 87H24585. Lift and drag coefficients for the 45-degrees sweptback configuration of the Bell X-5 research airplane were determined from flight tests covering the Mach number range from 0.61 to 1.01 and were compared to data obtained with the 59-degree sweptback configuration. Below the drag rise the 45-degree configuration had a zero lift drag coefficient of 40 0.020 as compared with 0.0175 for the 59-degree sweptback configuration. The lift-drag ratio for the 45-degree configuration exceeded that for the 59-degree configuration for a Mach number range from 0.61 to 0.88 with a maximum difference of about 0.7 at a Mach number of 0.82. 154. Saltzman, Edwin J.; Bellman, Donald R.; and Musialowski, Norman T.: Flight-Determined Transonic Lift and Drag Characteristics of the YF-102 Airplane with Two Wing Configurations. NACA RM H56E08, July 1956, 87H24588. The flight lift and drag characteristics of the YF-102 airplane, a 60-degree delta-wing interceptor, were determined for a symmetrical wing configuration and for a configuration with cambered wing and reflexed tips. The Mach number range extended from 0.6 to 1.17 and altitude varied from 25,000 feet to 50,000 feet. The cambered wing configuration experienced a considerable reduction of drag-due-to-lift, resulting in about 0.01 lower drag coefficient at about 0.3 lift coefficient and an increase of about 20 percent in lift-drag ratio. Comparable wing-tunnel data are included. E-2550 YF-102 Airplane 155. Beeler, De E.; Bellman, Donald R.; and Saltzman, Edwin J.: Flight Techniques for Determining Airplane Drag at High Mach Numbers. NACA TN 3821, AGARD Report #84, presented to the Flight Test Panel of Advisory Group for Aeronautical Research and Development, Brussels, Belgium, August 27–31, 1956, August 1956, 93R13811. The accelerometer method has proven to be the most satisfactory method for the flight measurement of total airplane drag during research investigations of high-speed airplanes by the NACA High-Speed Flight Station. The method requires special instrumentation and measuring techniques which are described in detail. Method for separating the flight measured overall drag into drag components and for comparing the flight data with windtunnel-model data are presented. 156. Williams, Walter C.; Drake, Hubert M.; and Fischel, Jack: Comparison of Flight and Wind-Tunnel Measurements of High-Speed-Airplane Stability and Control Characteristics. NACA TN 3859, August 1956, 93R13946. Comparisons of wind-tunnel and flight-measured values of stability and control characteristics are of considerable interest to the designer, since the wind-tunnel method of testing is one of the prime sources upon which estimates of the characteristics of a new configuration are based. In this paper comparisons are made of some of the more important stability and control characteristics of three swept-wing airplanes as measured in flight and in wind tunnels. Windtunnel data from high-speed closed-throat tunnels, a slottedthroat transonic tunnel, and a supersonic tunnel are used. The comparisons show that, generally speaking, the wind tunnels predict all trends of characteristics reasonably well. There are, however, differences in exact values of parameters, which could be attributed somewhat to differences in the model caused by the method of support. The small size of the models may have some effect on measurements of flap effectiveness. When nonlinearities in derivatives occur during wind-tunnel tests, additional data should be obtained in the region of the nonlinearities in order to predict more accurately the flight characteristics. Also, nonlinearities in static derivatives must be analyzed on the basis of dynamic motions of the airplane. Aeroelastic corrections must be made to wind-tunnel data for models of airplanes which have thin surfaces and are to be flown at high dynamic pressures. Inlet effects can exert an influence on the characteristics, depending upon air requirements of the engine and location of the inlets. 157. Williams, Walter C.; Drake, Hubert M.; and Fischel, Jack: Some Correlations of Flight-Measured and WindTunnel Measured Stability and Control Characteristics of High-Speed Airplanes. NACA RM H56AG62, August 1956, 87H29107, (also AGARD Report 62). Comparisons of wind-tunnel and flight-measured values of stability and control characteristics are of considerable interest to the designer, since the wind-tunnel method of testing is one of the prime sources upon which estimates of the characteristics of a new configuration are based. In this paper comparisons are made of some of the more important stability and control characteristics of three swept-wing airplanes as measured in flight and in wind tunnels. Windtunnel data are used from high-speed closed-throat tunnels, a slotted-throat transonic tunnel, and a supersonic tunnel. The comparison shows that, generally speaking, the wind tunnels predict all trends of characteristics reasonably well. There 41 are, however, differences in exact values of parameters, which could be attributed somewhat to differences in the model caused by the method of support. The small size of the models may have some effect on measurements of flap effectiveness. When non-linearities in derivatives occur during wind-tunnel tests, additional data should be obtained in the region of the non-linearities. Also, non-linearities in static derivatives must be analyzed on the basis of dynamic motions of the airplane. Aeroelastic corrections must be made to the wind-tunnel data for models of airplanes which have thin surfaces and are to be flown at high dynamic pressures. Inlet effects can exert an influence on the characteristics, depending upon air requirements of the engine and location of the inlets. 158. Weil, Joseph; and Day, Richard E.: Correlation of Flight and Analog Investigations of Roll Coupling. NACA RM H56F08, September 1956, 87H24591, 93R19575. A brief review of NACA flight experience relating to the rollcoupling problem is presented. Conditions rated by pilots and intolerable, marginal, and good are discussed and correlated with calculated results. A suggested flight test procedure for roll-coupling investigations and a discussion of several other items of general interest are also presented. 159. Drake, Hubert M.: Flight Experience With Present Research Airplanes. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, Langley Aeronautical Laboratory, Langley Field, Virginia, October 26, 1956, 93R21716. Declassified per NASA ccn 14, dated 25 April 1967. The North American X-15 airplane is being designed for speeds and altitudes considerably greater than those presently being encountered by airplanes. This paper explores the status of flight research with the current research airplanes to see what experience and planned research are pertinent to the X-15 project. E-5251 X-15 Airplane 160. Banner, Richard D.; and Malvestuto, Frank S., Jr.: Skin and Structural Temperature Measurements on Research Airplanes at Supersonic Speeds. ResearchAirplane-Committee Report on Conference on the Progress of the X-15 Project, Langley Aeronautical Laboratory, Langley Field, Virginia, October 26, 1956, 93R21723. Declassified per NASA ccn 14, dated 25 April 1967. Skin and structural temperatures of airplanes in flight at supersonic speeds have been determined by use of thermocouples and temperature resistance gages installed on various research airplanes. Such data have recently been obtained on two research airplanes, the Bell X-2 and the Bell X-1B. The object of this paper is to show some of the actual magnitudes and trends in the structural temperatures that exist in an airplane experiencing the effects of aerodynamic heating. 162. Taback, I.; and Truszynski, G. M.: Instrumentation for the X-15. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, Langley Aeronautical Laboratory, Langley Field, Virginia, October 26, 1956, 93R21729. Declassified per NASA ccn 14, dated 25 April 1967. The development of a research airplane which extends manned flight into regions where extremes of temperature and pressure are reached requires the simultaneous development of new instrumentation technique not only to ensure safe operation of the aircraft but also to derive a maximum of research data throughout the operational range of the aircraft. The instrumentation required for the North American X-15 airplane project consists of ground range and for research measurements. This paper outlines a plan for a ground range, which is based upon developed equipment already in use, and also discusses the airborne instrumentation and some of the special airborne devices which are made necessary by the extended performance capabilities of this airplane. 163. Marcy, William L.; Stephenson, Harriet J.; and Cooney, Thomas V.: Analysis of the Vertical-Tail Loads Measured During a Flight Investigation at Transonic Speeds of the Douglas X-3 Research Airplane. NACA RM H56H08, November 1956, 87H24605. An analysis of the vertical-tail loads obtained from straingage measurements during a flight investigation of the Douglas X-3 research airplane in the transonic speed range is presented. Data from rudder pulses, gradually increasing sideslips, and rudder-fixed aileron rolls were used to obtain rudder effectiveness and the effective lift-curve slope of the vertical tail. The variation of airplane yawing-moment coefficient with sideslip as determined from the measured vertical-tail loads is also presented. 164. Weil, Joseph: Review of Recent Rate of Roll Investigations at the NACA High-Speed Flight Station. Symposium on Roll Requirements, November 14, 1956, 87H29700. 165. Keener, Earl R.; and Jordan, Gareth H.: Wing Loads and Load Distributions Throughout the Lift Range of the Douglas X-3 Research Airplane at Transonic Speeds. NACA RM H56G13, December 1956, 87H24598. Wing loads and load distributions were obtained in flight by differential-pressure measurements between the upper and lower surfaces of the left wing of the Douglas X-3 research airplane. The effects of angle of attack and Mach number on the wing characteristics at transonic Mach numbers are shown. The wing has an aspect ratio of 3.09 and a modified 4.5 percent-thick hexagonal section. Data cover the range from near-zero lift to maximum lift, over a Mach number range of 0.71 to 1.15. Reynolds number based on the mean aerodynamic chord of the wing varied between 16 x 10 to the 6 power and 26 x 10 to the 6 power. 42 E-2539 X-1B Airplane 161. Stillwell, Wendell H.: Control Studies. Part B: Studies of Reaction Controls. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, Langley Aeronautical Laboratory, Langley Field, Virginia, October 26, 1956, 93R21728. Declassified per NASA ccn 14, dated 25 April 1967. The attitude-control method selected for the North American X-15 for flight at extremely low and zero dynamic pressures utilizes the reaction forces developed by small-rocket units located on the airplane to produce rolling, pitching, and yawing moments. An investigation of reaction control similar to those selected for the X-15 has shown that unique control problems exist for flight at the low dynamic pressures where this type of control is used. Although the Bell X-1B configuration was utilized for this investigation, a range of variables was covered to determine the significant effects of various factors on flight with reaction controls. It was also of interest to determine fuel requirements for the rocket units. The investigation consisted of analog-computer studies and ground-simulator tests. The significant results of this investigation is discussed in this paper. 166. Andrews, William H.; Sisk, Thomas R.; and Darville, Robert W.: Longitudinal Stability Characteristics of the Convair YF-102 Airplane Determined From Flight Tests. NACA RM H56I17, December 1956, 87H26846. The longitudinal stability and trim characteristics for the cambered-wing configuration of the Convair YF-102 airplane are determined up to M = 1.18 at altitudes of 25,000, 40,000, and 50,000 feet from level-flight speed runs, stall approaches, wind-up turns, and elevator pulses. Trim data are also included for the symmetrical-wing version. The trim characteristics are conventional. The static stability more than double between M = 0.60 and 1.16 and there is a loss of 50 percent in control effectiveness between M = 0.90 and 1.0. No severe pitch-up was exhibited except in cases resulting from speed change in the trim region. A preliminary analysis of the artificial-feel system was made as a result of the poor stick-force characteristics exhibited around 1.5g and 2.0g. 1957 Technical Publications 167. Sisk, Thomas R.; Andrews, William H.; and Darville, Robert W.: Flight Evaluation of the Lateral Stability and Control Characteristics of the Convair YF-102 Airplane. NACA RM H56G11, January 1957, 87H24594. The lateral stability and control characteristics of the Convair YF-102 delta-wing airplane with cambered-reflexed wings are determined from side-slips, aileron rolls, rudder pulses, trim runs, and wind-up turns. Violent inertial coupling has been encountered on this airplane and a summary of the rolling and sideslip characteristics is presented. The relation of the reciprocal of the cycles to damp to one-half amplitude with phi/Ve (Military Specification) varies from unsatisfactory to marginally satisfactory. Comparison is shown between the pilot’s rating of the rudder pulse E-2889 NACA High Speed Flight Station Research Aircraft, circa 1957 (clockwise from lower left), Bell X-1A, Douglas D-558-I, Convair XF-92A, Bell X-5, Douglas D-558-II, Northrop X-4, and Douglas X-3 (center) 43 maneuvers and the military requirement. A directional divergence that was encountered at high lift coefficient (alpha = 20 degrees) is presented. 168. Larson, Terry J.; Stillwell, Wendell H.; and Armistead, Katharine H.: Static Pressure Error Calibrations for Nose-Boom Airspeed Installation of 17 Airplanes. NACA RM H57A02, March 1957, 87H26847. Static-pressure error calibrations made for nose-boom airspeed installations of 17 airplanes are presented. The calibrations are given in the form of true Mach number against indicated Mach number, Mach number error, and static-pressure error per recorded impact pressure. Staticpressure errors are compared and are shown to be dependent on nose-boom length, fuselage diameter, and nose fineness ratio. Information is presented to provide a useful means for predicting the static-pressure errors for similar airspeed installation. Designation Airplane A, D-558-I; B, F-86A; C, F-86F; D, X-5; E, F-100A; F, XF-92A; G, YF-102; H, X-4; I, F-89; J, F-89; K, X-1; L, X-1; M, X-1A; N, D-558-II; O, D-558-II; P, X-3; and Q, B-47A. 170. Wolowicz, Chester H.: Dynamic Longitudinal Stability Characteristics of a Swept-Wing Fighter Type Airplane at Mach Numbers Between 0.36 and 1.45. NACA RM H56H03, April 1957, 87H24602. Longitudinal pulse maneuvers were conducted on a sweptwing fighter-type airplane for an original-wing and an extended wing-tip configuration at altitudes from 10,000 to 40,000 feet over a Mach number range from 0.36 to 1.45. Variations of the period, damping and the derivatives C (sub) L (sub) sigma, C (sub) m (sub) sigma, and (C sub m sub q + C sub m sub sigma) are presented as functions of Mach number. Comparisons are made with wind-tunnel data. Some consideration is given to pilot opinion in regard to the dynamic longitudinal behavior of the airplane during simulated combat maneuvers. 171. McGowan, William A.; and Cooney, T. V.: An Analysis of Vertical-Tail Loads Measured in Flight on a Swept-Wing Bomber Airplane. NACA RM L57B19, May 7, 1957, 93R18908. An analysis is presented of vertical-tail loads measured on a swept-wing bomber airplane at altitudes to 35,000 feet and Mach numbers to 0.82. Flight data obtained from rudder-step, rudder-pulse, aileronroll, and steady-sideslip maneuvers were used in the analysis to determine lift-curve slopes, centers of pressure, and wing-fuselage, tail, and airplane staticdirectional-stability parameters. Results are compared, where possible, with values used in design and with theoretical values. Theoretical values of the lift-curve slopes were in agreement with flight values when fuselage flexibility was considered. 172. Saltzman, Edwin J.: Flight-Determined InductionSystem and Surge Characteristics of the YF-102 Airplane With a Two-Spool Turbojet Engine. NACA RM H57C22, June 1957, 87H29841. Total-pressure recovery and distortion at the compressor face have been recorded for a twin-side inlet, two-spool turbojet engine combination during turns, sideslips, and speed runs at altitudes between 33,000 and 50,000 feet. In addition, conditions prior to several compressor surges have been recorded. The Mach number range covered extends from about 0.6 to 1.1. The investigation showed that engine surge as experienced is not related to distortion at the compressor face. Mismatching existed for most flight conditions. 173. Jordan, Gareth H.; Keener, Earl R.; and Butchart, Stanley P.: Airplane Motions and Loads Induced by Flying Through the Flow Field Generated by an Airplane at Low Supersonic Speeds. NACA RM H57D17A, June 1957, 87H24646. Data are presented for the maximum sideslip angles and vertical-tail loads induced on a swept wing fighter-type airplane as a result of flying through the flow field generated E55-02094 F-100A Super Sabre Airplane 169. Matranga, Gene J.; and Peele, James R.: FlightDetermined Static Lateral Stability and Control Characteristics of a Swept-Wing Fighter Airplane to a Mach Number of 1.39. NACA RM H57A16, March 1957, 87H26848. Flight tests were performed with a swept-wing fighter-type airplane at an altitude of 40,000 feet over a Mach number range from 0.72 to 1.39 to determine the lateral stability and control characteristics. Results are presented for three different vertical tails and two different wing areas, and include plots of lateral stability derivatives against Mach number. 44 by a similar airplane at low supersonic Mach numbers. These data were obtained during side-by-side passes at various passing rates (5 fps to 50 fps) and interval separation distances. Significant airplane sideslip angles and vertical-tail loads were obtained during close-proximity passes at a passing rate near the natural period of the airplane in yaw. 174. Banner, Richard D.: Flight Measurements of Airplane Structural Temperatures at Supersonic Speeds. NACA RM H57D18B, June 1957, 93R19049. Skin and structural temperature distributions were obtained during transient supersonic flights of the X-1B and X-1E airplanes at Mach numbers up to approximately 2.0. Extensive temperature measurements were obtained on the X-1B. No critical temperatures were experienced over the range of the test. The measured temperatures were compared with simplified calculations. Mach number range from 0.48 to 1.03. The addition of stores increased the drag at all Mach numbers tested. Below the drag rise the increase was of about the same magnitude as the increase in wetted area caused by the addition of the stores. The peak lift-drag ratio was reduced by about 14 percent, and for lift coefficients of 0.2 and 0.4 a reduction in drag-rise Mach number was noted. Little change in lift-curve slope was observed for Mach numbers less than about 0.8. 177. Saltzman, Edwin J.; and Asher, William P.: Transonic Flight Evaluation of the Effects of Fuselage Extension and Indentation on the Drag of a 60 Degrees Delta-Wing Interceptor Airplane. NACA RM H57E29, September 1957, 87H26849, 71N72659. The flight lift and drag characteristics of a 60 degrees delta wing interceptor airplane incorporating fuselage extension and indentation were determined over the Mach number range from 0.7 to 1.15 and the altitude range from 25,000 to 50,000 feet. Comparison is made with a similar airplane which did not utilize fuselage extension or indentation. The results indicate that the modifications (extension and indentation) reduced the transonic drag coefficient about 50 drag counts (0.0050) at a Mach number of about 1.1. Three sets of comparable low Reynolds number data are included which indicate reductions in transonic drag coefficient ranging from about 0.0025 to 0.0045 at a Mach number of about 1.1. 178. Fischel, Jack; Darville, Robert W.; and Reisert, Donald: Effects of Wing-Mounted External Stores on the Longitudinal and Lateral Handling Qualities of the Douglas D-558-II Research Airplane. NACA RM H57H12, October 1957, 87H24664. E-1927 X-1E Airplane 175. Malvestuto, Frank S.; Cooney, Thomas V.; and Keener, Earl R.: Flight Measurements and Calculations of Wing Loads and Load Distributions at Subsonic, Transonic, and Supersonic Speeds. NACA RM H57E01, July 1957, 87H24650. Presented in this report is a summary of local and net angleof-attack wing-panel loads measured in flight on six airplanes. In addition, a comparison of these loads measured in flight with calculations based on simple theory is presented. 176. Nugent, Jack: Effect of Wing-Mounted External Stores on the Lift and Drag of the Douglas D-558-II Research Airplane at Transonic Speeds. NACA RM H57E15A, July 1957, 87H24654, 93R19584. Lift and drag measurements were made during a flight investigation with the Douglas D-558-II(145) airplane in the basic and 150-gallon DAC store configurations over a 45 The subsonic and transonic handling qualities of the Douglas D-558-II research airplane were investigated with several configuration of midsemispan external stores in the altitude region between 20,000 and 40,000 feet. The configurations tested consisted of an underslung pylon on each wing, pylons plus simulated DAC (Douglas Aircraft Co.) 1,000-pound bombs, and pylons plus DAC 150-gallon-fuel tanks. Comparisons of the results obtained were made with comparable data from the clean airplane. The trends exhibited in the characteristics measured with each configuration were generally the same as for the clean airplane; however, significant changes in the magnitude of the parameters measured with the pylon-tank configuration were sometimes apparent, particularly at the higher speeds tested. 179. Baker, Thomas F.; Martin, James A.; and Scott, Betty J.: Flight Data Pertinent to Buffeting and Maximum Normal-Force Coefficient of the Douglas X-3 Research Airplane. NACA RM H57H09, November 1957, 87H26850. The X-3 airplane, which has a straight, 4.5-percent-thick wing, was flown to maximum wing lift at Mach numbers from 0.7 to 1.1 at an average altitude of 30,000 feet. Airplane E57-2952 NACA High Speed Flight Station Century Series Fighter Aircraft, circa 1957 (clockwise from left), Lockheed F-104, McDonnell F-101, Convair F-102, and North American F-100 and wing maximum normal-force coefficients and buffeting characteristics were determined. Wing maximum normalforce coefficients at low supersonic speeds were almost twice the values at subsonic speed (M nearly equal to 0.8). At transonic speeds the buffet boundary abruptly increased, rather than decreased, with Mach number. The buffeting encountered did not constitute either an operational or a structural problem. The effective longitudinal maneuverability limit was defined by maximum wing lift. Limited data at subsonic speeds on the effects on lift and buffeting of deflecting the wing leading edge flaps 7 degrees are included. 180. Drake, Hubert M.; and Kincheloe, Iven C.: Flight Research at High Altitude. SLN: 63423, November 8, 1957, 87H29332. In the past ten years the Air Force, Navy, and the NACA, in the research airplane program have obtained considerable experience in flight research at high altitude. The present paper details some of this experience, discusses some of the problems encountered and being investigated, and a quick look into future is taken. 181. Fischel, Jack; Holleman, Euclid C.; and Tremant, Robert A.: Flight Investigation of the Transonic Longitudinal and Lateral Handling Qualities of the Douglas X-3 Research Airplane. NACA RM H57I05, December 1957, 87H26851. 46 A flight investigation was performed to determine the longitudinal and lateral handling qualities of the Douglas X-3 research airplane in the clean configuration and with wing leading-edge flaps deflected. Static and dynamic stability and control characteristics were determined during trimmed and maneuvering flight at an average altitude of 30,000 feet over a Mach number range from 0.7 to 1.16. Statically and dynamically determined stability and control derivatives are presented, as well as pilot evaluation of the airplane. 182. Saltzman, Edwin J.: In-Flight Gains Realized by Modifying a Twin Side-Inlet Induction System. NACA RM H57J09, December 1957, 87H24669. The effects of modifying a twin side-inlet duct system have been recorded and analyzed over an altitude range from about 25,000 to 51,000 feet and throughout the transonic region to a Mach number of about 1.2. The modification consisted primarily of redesigning the inlet lip, increasing the crosssectional area of the inlet and diffuser, and adding a region of duct contraction ahead of the engine. These changes greatly improved the pressure-recovery characteristics and provided a 50-percent reduction in compressor-face distortion (pressure-profile variation). 183. Drake, H. M.: Flight Research at High Altitude, Part 1. AGARD Proceedings of the Seventh AGARD General Assembly, 1957, pp. 74–75, (see N82-73409 12-01), 82N73414, #. Aerodynamic problems associated with flight at high Mach numbers and/or low dynamic pressures include reduction or directional stability, poor dynamic stability, low control effectiveness, and aerodynamic heating, and instrumentation problems. These are briefly indicated. Some of the highaltitude investigations on the X-1B aircraft are discussed. Including preliminary studies on an analogue computer used as a simulation. 184. Williams, Walter C.; and Drake, Hubert M.: The Research Airplane—Past, Present, and Future. IAS Summer Meeting, June 18, 1957, Preprint 750, 1957, 87H29153. E-1044 This paper discusses briefly the problems that have been studied by the NACA during the past ten years using research airplanes. 185. Beeler, De Elroy: Flight Loads Measurements on NACA Research Airplanes. AGARD Report 109, Structures and Materials Panel, 1957, 87H29112. Summarizes results from flight loads investigations conducted primarily for the purpose of confirming windtunnel results by comparing full-scale results with comparable wind-tunnel results. Data have been selected from three research aircraft having a wide range of configuration. (X-1, X-5, & XF-92A). Flight loads were determined by use of calibrated strain gauges. Pressure distribution measurements were also made of fuselage & wing of the X-1. 186. Cole, Henry Ambrose; Brown, Stuart C.; and Holleman, Euclid C.: The Effects of Flexibility on the Longitudinal Dynamic Response of the B-47 Airplane. Preprint IAS 678, 1957, 87H29016. (See also 121.) The frequency response of the B-47 to elevator control is presented for ranges of frequencies including the airplane oscillatory short period mode, the wing first-bending mode. Comparisons are made between response measured in flight and those predicted by pseudostatic analysis and by dynamical analysis. In the application of pseudostatic analysis, the importance of mass distribution is emphasized. Selection of dynamic degrees of freedom and the apparent disappearance of modes under forced oscillation is discussed. The frequency response at various points on the airplane is measured and calculated to find optimum locations for automatic control pickups. The points of low response to structural vibrations determined in flight are compared with those predicted by dynamical analysis and those measured in ground vibration tests. B-47 Airplane 187. Banner, R. D.: Flight Measurements of Airplane Structural Temperatures at Supersonic Speeds. In NACA Conference on Aircraft Loads, Structures, and Flutter, 1957, (see N71-75382), 71N75404. 188. Jordan, Gareth H.; Keener, Earl R.; and Butchart, Stanley P.: Airplane Motions and Loads Induced by Flying Through the Flow Field Generated by an Airplane at Low Supersonic Speeds. NACA Conference on Aircraft Loads, Structures, and Flutter, 1957, (see N71-75382), 71N75394. An exploratory flight investigation was conducted to determine the disturbances to an airplane while flying in formation with another airplane at low supersonic speeds. The most significant motions were encountered as a result of flying through the flow field of the lead airplane. Several of these supersonic passes were made using two sweptwing fighter-type airplanes in order to evaluate the gross effects of time to pass through the flow field, lateral distance, and altitude within a Mach number range from 1.2 to 1.3. 189. Malvestuto, Frank S.; Cooney, Thomas V.; and Keener, Earl R.: Flight Measurements and Calculations of Wing Loads and Load Distribution at Subsonic, Transonic, and Supersonic Speeds. NACA Conference on Aircraft Loads, Structures, and Flutter, 1957, (see N71-75382), 71N75385. 1958 Technical Publications 190. Holleman, Euclid C.; and Boslaugh, David L.: A Simulator Investigation of Factors Affecting the Design and Utilization of a Stick Pusher for the Prevention of Airplane Pitch-Up. NACA RM H57J30, January 1958, 87H26852. 47 Presented are the results of a simulator study of the factors affecting the design of a device, a stick pusher, for preventing a representative supersonic airplane from entering the pitchup region. The effects of varying the stick-pusher-activation boundaries, sensing parameters, and magnitude of stickpusher force on the controllability of the airplane pitch-up were investigated. The possible tactical importance of the loss in available supersonic maneuverability caused by angle-of-attack limiting in turns and zoom maneuvers is also discussed. 191. Williams, Walter C.; and Drake, Hubert M.: The Research Airplane—Past, Present, and Future. Aeronautical Engineering Review, Vol. 36, January 1958. 192. Marcy, William L.: High-Speed Landing Loads Measured on the Douglas X-3 Research Airplane. NACA RM H57L06, February 1958, 87H24675. Measured loads are shown for several landings of the Douglas X-3 at speeds from about 310 to 420 feet per second. Sinking speeds were between 2 and 5 feet per second. Some loads during taxiing and turning are also shown. 193. Matranga, Gene J.; and Armistead, Katharine H.: Flight Evaluation of the Effects of Leading-Edge-Slat Span on the Stability and Control Characteristics of a Sweptwing Fighter-Type Airplane During Accelerated Longitudinal Maneuvers at Transonic Speeds. NACA RM H58A03A, March 1958, 87H26853. Accelerated longitudinal maneuvers were performed at transonic speeds with a swept-wing fighter-type airplane having several slat-span configurations. The effects of these slat-span configurations on the stability and control characteristics of the airplane are discussed and compared with existing wind-tunnel data. 194. Drake, Hubert M.; Bellman, Donald R.; and Walker, Joseph A.: Operational Problems of Manned Orbital Vehicles. NACA Conference on High Speed Aerodynamics, March 18–20, 1958, pp. 89–102, (see N71-75285), 71N75292. 195. Banner, R. D.; McTigue, J. G.; and Petty, G., Jr.: Boundary-Layer Transition in Full-Scale Flight. NACA Conference on High-Speed Aerodynamics, March 18–20, 1958, pp. 467–475, (see N71-75285), 71N75319. Because of greatly increased need for knowledge of full-scale boundary-layer transition and the difficulty of simulating actual flight conditions, a program has been initiated to provide a better understanding of the boundary-layer flow as it exists in supersonic flight. This paper shows the results obtained in the early flight tests which determined the content of laminar flow that could be obtained with practical wingsurface conditions. 48 196. Taillon, Norman V.: An Analysis of Surface Pressures and Aerodynamic Load Distribution Over the Swept Wing of the Douglas D-558-II Research Airplane at Mach Numbers From 0.73 to 1.73. NACA RM H58A30, April 1958, 87H26196, 87H24678. The variation of the measured pressures and loads with lift is presented, covering the speed range at airplane normal-force coefficients from about 0 to 0.8. Spanwise load distributions are compared with theory at subsonic and supersonic Mach numbers. The wing is swept back 35 degrees, has a streamwise thickness ratio varying from 8.7 percent at the root to 10.4 percent at the tip, an aspect ratio of 3.57, and 3 degrees incidence. 197. Kuhl, A. E.; Little, M. V.; and Rogers, J. T.: Analysis of Flight-Determined and Predicted Effects of Flexibility on the Steady-State Wing Loads of the B-52 Airplane. NACA RM H57C25, April 23, 1958, 87H26193, 87H24615, 66N39617, #. An analysis is made of the steady-state wing loads measured during a flight investigation conducted on the Boeing B-52 airplane. The investigation covers the speed range of the airplane up to a Mach number of 0.90. The measured loads are analyzed in terms of the effects of Mach number, lift, and dynamic pressure. In addition, the measured loads are compared with predicted loads and the effects of varying some of the structural and aerodynamic properties in the predictions are presented. 198. McLeod, Norman J.; and Jordan, Gareth H.: Preliminary Flight Survey of Fuselage and BoundaryLayer Sound-Pressure Levels. NACA RM H58B11, May 1958, 87H26854, 87H26199. Presented are the results of a preliminary flight investigation of noise inside fuselages and in the fuselage boundary layer. The overall noise and frequency spectrum are presented for the B-47A at subsonic speeds. Measurements were made internally at three longitudinal locations and in the fuselage boundary layer at one location. The overall noise level is presented for one internal location of the D-558-II at subsonic and low supersonic speeds. The relative importance of engine noise and aerodynamic noise due to the boundary layer at various airspeeds and engine operating conditions is discussed for the locations at which measurements were made. 199. Johnson, Clinton T.: Flight Investigation of the Aerodynamic Forces on a Wing-Mounted External-Store Installation on the Douglas D-558-II Research Airplane. NACA RM H58B24, May 1958, 87H26203, 87H24680. Aerodynamic forces have been measured on an external-store arrangement on the Douglas D-558-II research airplane. Aerodynamic loads on the store-pylon combination and the pylon alone were determined from angle-of-attack and angle-of-sideslip maneuvers over the Mach number range from 0.50 to 1.03. Wing-panel aerodynamic loads have also been determined for the store configuration and the cleanwing configuration. A brief discussion is presented for the measured loads on the fins of the external store. 200. Larson, Terry J.; Thomas, George M.; and Bellman, Donald R.: Induction System Characteristics and Engine Surge Occurrence for Two Fighter-Type Airplanes. NACA RM H58C14, May 1958, 87H26206, 87H24681, 71N70251. Total-pressure recovery and distortion at the compressor face are presented as variations of angle of attack and mass-flow ratio for two fighter-type airplanes with similar two-spool turbojet engines, but with dissimilar inlets. One airplane has a normal-shock nose inlet, and the other has two engines and triangular-shaped inlets located in the wing roots. In addition, data are presented for engine surge occurrences of these two airplanes and also for a third single-engine airplane having two semicircular-shaped side inlets. 201. Pembo, Chris; and Matranga, Gene J.: Control Deflections, Airplane Response, and Tail Loads Measured on an F-100A Airplane in Service Operational Flying. NACA RM H58C26, June 1958, 87H26211, 87H24682. Results are presented from 20 hours of service operational flying of an instrumented North American F-100A fighter airplane. Air Force pilots at Nellis Air Force Base, Nev., performed air-to-air gunnery, simulated air-to-ground attacks, air-combat maneuvering, acrobatics, and transitiontype flights to altitudes slightly in excess of 50,000 feet and at Mach numbers up to 1.22. Measurements of pilot control deflections, throttle movement, airplane motions, and tail loads are presented primarily as envelope curves of maximum recorded values. An overall comparison of these results with results from earlier investigations with other airplanes is made. A limited analysis of the factors affecting horizontaland vertical-tail loads is included. 202. Drake, Hubert M.; Bellman, Donald R.; and Walker, Joseph A.: Operational Problems of Manned Orbital Vehicles. NACA RM H58D21, July 1958, 87H26216, 87H24686. Some of the operational problems of escape, piloting, orbit selection, flight termination, and range requirements are considered, including the effects of configuration. It is indicated that configuration materially affects operations, that survival considerations may preclude optimum procedures, and that use of the pilot can considerably simplify design and increase reliability. 203. Holleman, Euclid C.; and Stillwell, Wendell H.: Simulator Investigation of Command Reaction Controls. NACA RM H58D22, July 1958, 87H26855, 87H26220. 49 Reaction controls that command velocity and attitude have been investigated and are compared to controls that command acceleration. Velocity and attitude command systems facilitated the task of orientation and stabilization and minimized the effects of dynamic pressure. The simulation of an entry maneuver by dynamic-pressure buildup showed that successful entry could be made with any of the three systems, but the task was easier with the velocity or the attitude system. 204. Holleman, E. C.; and Stillwell, W. H.: Simulator Investigation of Command Reaction Controls. In NACA Conference on High-Speed Aerodynamics, 1958, pp. 157–165, (see N71-75285), 71N75297. (See also 203.) 205. Keener, Earl R.; McLeod, Norman J.; and Taillon, Norman V.: Effect of Leading-Edge-Flap Deflection on the Wing Loads, Load Distributions, and Flap Hinge Moments of the Douglas X-3 Research Airplane at Transonic Speeds. NACA RM H58D29, July 1958, 87H26221, 87H24690. Data were obtained in flight by differential-pressure measurements between the upper and lower surfaces of the wing, covering the range from near-zero lift to maximum lift over the Mach number range of 0.5 to 1.15 with flap undeflected and 0.5 to 0.9 with flap deflected. The unswept wing has an aspect ratio of 3.09, a taper ratio of 0.39, and a modified 4.5-percent-thick hexagonal section. The plain, constant-chord leading-edge flap extends from the wing root to the wing tip. Reynolds number based on the mean aerodynamic chord of the wing varied between 16 X 10 sup 6 and 26 X 10 sup 6. A brief comparison with wind-tunnel results is included. 206. Banner, Richard D.; McTigue, John G.; and Petty, Gilbert: Boundary-Layer-Transition Measurements in Full-Scale Flight. NACA RM H58E28, July 1958, 78N78570, 87H26856, 87H26226. (See NASA TM-79863). Chemical sublimation employed for boundary-layer flow visualization on the wings of a supersonic fighter airplane in level flight near a Mach number of 2.0 has shown that laminar flow can be obtained over extensive areas with practical wing surface conditions. Heated temperature resistance gages installed in a Fiberglas “glove” installation on one wing continuously monitored the conditions of the boundary layer. Tests were conducted at speeds from a Mach number of 1.2 to a Mach number of 2.0, at altitudes from 35,000 feet to 56,000 feet. Data obtained at all angles of attack, from near 0 degrees to near 10 degrees have shown that the maximum transition Reynolds number on the upper surface of the wing varies from about 2.5 x 10 to the 6 power at a Mach number of 1.2 to about 4 x 10 to the 6 power at a Mach number of 2.0. On the lower surface, the maximum transition Reynolds number varies from about 2 x 10 to the 6 power at a Mach number of 1.2 to about 8 x 10 to the 6 power at a Mach number of 2.0. 207. Fischel, Jack; Cooney, Thomas V.; and Bellman, Donald R.: Flight-Testing Techniques of Manned Hypersonic and Satellite Vehicles. IAS, Los Angeles, California, July 10–11, 1958, 87H29821. The desire of man to fly at ever increasing speeds and altitudes has resulted in a constantly accelerating performance growth. The technology has advanced at such a rate that future flight vehicles are planned not for small incremental performance increases as in the past, but rather, in multiples of present performance. With these increases in performance it is the problem of the flight test engineer to assure that his methodology is compatible with the vehicle. It is well, then, to take the case of a boost-glide vehicle capable of speed up to that required for orbital flight, examine its characteristics and problems, and outline the approach to the flight test of such a vehicle. 208. *Finch, Thomas W.; *Matranga, Gene J.; *Walker, Joseph A.; and *Armstrong, Neil A.: Flight and Analog Studies of Landing Techniques. This paper is from the Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project held at the IAS Building, Los Angeles, California on July 28–30, 1958, NACA-CONF30-Jul-58, July 30, 1958, pp. 83–93, 93R21698. Declassified per NASA ccn 14, 29 Nov. 1966. The approach and landing operation of unpowered rocket airplanes has always required considerable pilot concentration but has been completed without undue demands on piloting technique. The X-15 airplane will land in a range of lift-drag ratio l/d markedly lower than previous rocket airplanes have used. This paper presents the results of a flight and analog study of landing to assess the potential difficulty of landing the X-15 at low l/d and to determine whether different techniques would be required in the landing maneuver. *NACA High Speed Flight Station, Edwards, California. 209. *Rowen, Burt.: Aeromedical Support of the X-15 Program. Research-Airplane-Committee Report from Conference on the Progress of the X-15 project held at the IAS building, Los Angeles, California on July 28–30, 1958, NACA-CONF-30-Jul-58, July 30, 1958, pp. 129–146, 93R21680. Declassified per NASA ccn 14, 29 Nov. 1966. Report describes the human factors or aeromedical support program for the X-15. The overall objective is to obtain quantitative physiological data and to make the pilot’s actual flight task a realistic continuation of previous experience and training. During the flight phase of the X-15 aircraft, physiological data will be telemetered so that a flight surgeon observing the ground read-out can tell when the pilot is approaching the limit of his physiological tolerance. 210. Truszynski, G. M.; and Mace, W. D.: Status of High-Range and Flow-Direction Sensor. ResearchAirplane-Committee Report from Conference on the Progress of the X-15 project held at the IAS building, Los Angeles, California on July 28–30, 1958, NACA-CONF30-Jul-58, July 30, 1958, pp. 151–158, 93R21682. Declassified per NASA ccn 14, 29 Nov. 1966. This paper describes the two systems used to provide certain research measurements for the X-15 aircraft. These systems are: (1) a probe and associated system that will be capable of operating throughout the extreme temperature environment encountered on reentry to provide a measure of the angle of attack and sideslip to the pilot, and (2) an instrumented ground range capable of monitoring the flight of the airplane throughout its entire trajectory. 211. *Beeler, De E.: X-15 Research Objectives. Research-Airplane-Committee Report from Conference on the Progress of the X-15 project held at the IAS building, Los Angeles, California on July 28–30, 1958, NACA-CONF-30Jul-58,July 30, 1958, pp. 327–338, 93R21697. Declassified per NASA ccn 14, 29 Nov. 1966. Paper presents the areas of research interest for the most important and urgent problems at the present time. Indications are given of other types of data that will be obtained, as well as possible additional research uses of the X-15. In the course of conducting the flight research for the X-15, the emphasis will change from one area to another and problems of new and different significance will result. Those problems that are found to be real will be better understood as a result of the flight investigations and those problems that have been imagined will be replaced with unexpected or overlooked problems. *NACA High Speed Flight Station, Edwards, California. 980066 *Air Force Flight Test Center, Edwards AFB, California. 50 X-15 Airplane, Three-View Drawing 212. Wolowicz, Chester H.; and Rediess, Herman A.: Effects of Jet Exhaust on Flight-Determined Longitudinal and Lateral Dynamic Stability Characteristics of the Douglas D-558-II Research Airplane. NACA RM H57G09, August 1958, 87H24659. A flight investigation over a Mach number range from 0.67 to 1.61 was made to determine the longitudinal and lateral stability characteristics of the D-558-II airplane with particular reference to the jet exhaust effects of the rocket engine. Longitudinal stability was not affected because of the high horizontal tail. The rudder-fixed lateral stability was adversely affected in the supersonic region. Rudder-free lateral stability during power-on supersonic yawed flight was better than for rudder-fixed conditions. 213. Cooney, Thomas V.: The Service Pilot as a Test Pilot. SETP Quarterly Review, Summer 1958. 214. Stillwell, Wendell H.; and Drake, Hubert M.: Simulator Studies of Jet Reaction Controls for Use at High Altitude. NACA RM H58G18A, September 1958, 87H26229, 87H24697, 93R19525. An investigation has been made of the use of pilot-controlled jet reaction forces for vehicle attitude control in regions of extremely low dynamic pressure. The effects of various control configurations, control magnitudes, control techniques, dynamic pressure, and aerodynamic stability were investigated. The results of analog computer studies and mechanical simulator tests indicate that control techniques are somewhat different from those used with aerodynamic controls at normal flight speeds and that constant attention to the control task is required. Because of the ease of overcontrolling with large control powers, much lower control power than that required for aerodynamic controls was preferred. Moderate values of effective dihedral produced a noticeable increase in the amount of roll control required to maintain trim at dynamic pressures up to 20 pounds per square foot. 215. Wolowicz, Chester H.; and Holleman, Euclid C.: Stability-Derivative Determination From Flight Data. Presented to AGARD Flight Test Panel, Copenhagen, Denmark, October 20–25, 1958, Report 224 OTP-1958, October 1958, 87H29966, 87H29964. A comprehensive discussion of the various factors affecting the determination of stability and control derivatives from flight data is presented based on the experience of the NASA High-Speed Flight Station. Factors relating to test techniques, determination of mass characteristics, instrumentation, and methods of analysis are discussed. 216. Fischel, Jack; Butchart, Stanley P.; Robinson, Glenn H.; and Tremant, Robert A.: Flight Studies of Problems Pertinent to Low-Speed Operation of Jet Transports. NASA Conference on Some Problems Related to Aircraft Operations, November 5–6, 1958. 217. Cooney, Thomas V.: Motions and Vertical-Tail Loads Experienced by Jet Transport Aircraft in Rough Air. NASA Conference on Some Problems Related to Aircraft Operations, November 5–6, 1958. 218. Jordan, Gareth H.; and McLeod, Norman J.: Boundary-Layer Noise at Subsonic and Supersonic Speeds. NASA Conference on Some Problems Related to Aircraft Operations, November 5–6, 1958, 87H29996. Presents results of flight surveys of boundary-layer and engine noise levels in an attempt to establish the contribution and relative importance of boundary-layer and engine noise on the noise environment of aircraft in flight. The two airplanes used were the B-47A and the Douglas 558-2. 219. Armstrong, Neil A.: Future Range and Flight Test Area Needs for Hypersonic and Orbital Vehicles. Society of Experimental Test Pilots, Vol. 3 No. 2, (see F70-0705), Winter 1959. This paper points out some of the desirable and mandatory area and control requirements for the flight testing of such vehicles up to and including those of orbital velocity with indications of the needs of operational military and air carrier organizations. 1959 Technical Publications 220. Nugent, Jack: Lift and Drag of a Swept-Wing Fighter Airplane at Transonic and Supersonic Speeds. NASA Memo 10-1-58H, January 1959, 87H25781. Lift and drag measurements were made during a flight investigation of a swept-wing fighter airplane in the basic configuration and in a slats-locked-closed configuration over a Mach number range from about 0.63 to about 1.44. Negligible drag-coefficient difference existed between the 51 E-2578 Iron Cross Reaction Controls Simulator basic configuration and the slats-locked-closed configuration over a comparable test range. For the basic configuration at zero lift the supersonic drag-coefficient level was about three times as great as the subsonic drag-coefficient level, which was about 0.01. 221. Matranga, Gene J.: Roll Utilization of an F-100A Airplane During Service Operational Flying. NASA Memo 12-1-58H, January 1959, 87H25787. An instrumented North American F-100A fighter airplane was flown by Air Force pilots at Nellis Air Force Base, Nev., during 20 hours of service operational flying which included air-to-air gunnery, air-to-ground gunnery and bombing, aircraft combat maneuvering, and acrobatic-type maneuvers. During this investigation altitudes up to 50,000 feet and Mach numbers up to 1.22 were realized. In this paper the roll utilization for the various maneuvers performed and the roll utilization at supersonic speeds are analyzed. 222. Walker, Joseph A.: Outline of the X-15 Project. Presented at the IAS Meeting, San Francisco, California, January 15, 1959. This paper presents a brief discussion of the following phases of this project: history, cost, research objectives, protection and escape of the pilot, description of the aircraft, industry participation, flight-test range, role of the pilots, and flight testing. 223. Williams, Walter C.: X-15 Airplane as a Research Tool. Presented at the IAS 20th Annual Meeting, New York, January 26–29, 1959, IAS Report, pp. 59–79. 224. Messing, Wesley E.: Residual Fuel Expulsion From a Simulated 50,000-Pound-Thrust LiquidPropellant Rocket Engine Having a Continuous RocketType Igniter. NASA Memo 2-1-59H, February 1959, 87H25790. Tests were conducted to determine the starting characteristics of a simulated 50,000-pound-thrust rocket engine with a quantity of fuel lying dormant in the main thrust chamber. Ignition was alcohol-water and anhydrous ammonia was used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter was equipped with a sonic discharge nozzle. When the igniter was equipped with a supersonic exhaust nozzle, a slightly less effective expulsion rate was encountered. 225. Weil, Joseph: Summary of Planned X-15 Entry Research. NASA Conference on Review of NASA Research Related to Control Guidance and Navigation of Space Vehicles, NASA Ames Research Center, February 25–27, 1959. 52 226. Drake, Hubert M.: Energy Management Requirements of Entry Vehicles. NASA Conference on Review of NASA Research Related to Control Guidance and Navigation of Space Vehicles, NASA Ames Research Center, February 25–27, 1959. 227. Finch, Thomas W.: Flight and Analog Studies of Approach and Landing Characteristics of Low L/D Configurations. NASA Conference on Review of NASA Research Related to Control Guidance and Navigation of Space Vehicles, NASA Ames Research Center, February 25–27, 1959. 228. Fischel, Jack: Use of Air-Launched Techniques in Space Research. NASA Conference on Review of NASA Research Related to Control Guidance and Navigation of Space Vehicles, NASA Ames Research Center, February 25–27, 1959. 229. Holleman, Euclid C.: Utilization of Pilot During the Boost Stage of Multistaged Vehicles. NASA Conference on Review of NASA Research Related to Control Guidance and Navigation of Space Vehicles, NASA Ames Research Center, February 25–27, 1959. 230. Boslaugh, David L.: Investigation of Precise Attitude Control—Simulator Program. NASA Conference on Review of NASA Research Related to Control Guidance and Navigation of Space Vehicles, NASA Ames Research Center, February 25–27, 1959. 231. Funk, Jack; and Cooney, T. V.: Some Effects of Yaw Damping on Airplane Motions and Vertical-Tail Loads in Turbulent Air. NASA Memo 2-17-59L, March 1959. Results of analytical and flight studies are presented to indicate the effect of yaw damping on the airplane motions and vertical-tail loads in rough air. The analytical studies indicate a rapid reduction in the airplane motions and the loads on the vertical tail as the damping is increased up to the point of damping the lateral motions to 1/2 amplitude in one cycle. Flight measurements indicate that a yaw damper reduces the loads on the vertical tail due to turbulent air and that the pilot could provide a significant amount of lateral damping. 232. McKay, J. B.: Problems Associated with HighSpeed Flight. NASA TM X-56245. Presented at the Air Force Academy Instructors Workshop, Moore AFB, Mission, Texas, 18 March 1959, 1959, 65N83273. 233. Armstrong, Neil A.: Test Pilot’s View on Space Ventures. American Society of Mechanical Engineers Meeting in Los Angeles, California, March 11, 1959. 234. Andrews, William H.; and Rediess, Herman A.: Flight-Determined Stability and Control Derivatives of a Supersonic Airplane With a Low-Aspect-Ratio Unswept Wing and a Tee-Tail. NASA Memo 2-2-59H, April 1959, 87H25794. The longitudinal and lateral directional stability and control derivatives have been determined in the trim angle-of-attack range between Mach numbers of 0.88 and 2.08. The static and dynamic lateral directional derivatives were determined by the time-vector method, modified by using yawing velocity as a reference instead of sideslip angle. Generally, the flight data compared favorably with existing published and unpublished wind-tunnel data. 235. Fischel, Jack; Butchart, Stanley P.; Robinson, Glenn H.; and Tremant, Robert A.: Flight Studies of Problems Pertinent to Low-Speed Operation of Jet Transports. NASA Memo 3-1-59H, April 1959, 87H25799. The specific areas investigated include those of the take-off and landing, and relation of these maneuvers to the 1 g stall speed and stalling characteristics. The take-off studies included evaluation of the factors affecting the take-off speed and attitude, including the effects of premature rotation and of overrotation on ground run required. The approach and landing studies pertained to such factors as: desirable lateral-directional damping characteristics; lateral-control requirements; space-positioning limitations during approach under VFR or IFR conditions and requirements for glide-path controls; and evaluation of factors affecting the pilot’s choice of landing speeds. Specific recommendations and some indication of desirable characteristics for the jet transports are advanced to alleviate possible operational difficulties or to improve operational performance in the low-speed range. 236. Butchart, Stanley P.; Fischel, Jack; Tremant, Robert A.; and Robinson, Glenn H.: Flight Studies of Problems Pertinent to High-Speed Operation of Jet Transports. NASA Memo 3-2-59H, April 1959, 87H25803. Some of the specific areas investigated include: (1) an overall evaluation of longitudinal stability and control characteristics at transonic speeds, with an assessment of pitch-up characteristics, (2) the effect of buffeting on airplane operational speeds and maneuvering, (3) the desirable lateraldirectional damping characteristics, (4) the desirable lateralcontrol characteristics, (5) an assessment of overspeed and speed-spread requirements, including the upset maneuver, and (6) an assessment of techniques and airplane characteristics for rapid descent and slow-down. The results presented include pilot’s evaluation of the various problem areas and specific recommendations for possible improvement of jet-transport operations in the cruising speed range. 53 237. Bellman, Donald R.: A Summary of FlightDetermined Transonic Lift and Drag Characteristics of Several Research Airplane Configurations. NASA Memo 3-3-59H, April 1959, 87H25804. Flight-determined lift and drag data from transonic flights of seven research airplane configurations of widely varying characteristics are presented and compared with wind-tunnel and rocket-model data. The effects of some of the basic configuration differences on the lift and drag characteristics are demonstrated. 238. Williams, Walter C.: Pilot Considerations in the X-15 Research Airplane Program. Presented at the Annual Meeting of the American Psychiatric Association, Philadelphia, Pennsylvania, April 29, 1959. This report briefly discusses the aircraft, the pilots working space and environment, the pilots selected for flights, simulator programs, and the monitoring of the pilot’s physical condition during flight. 239. Taillon, Norman V.: Flow Characteristics About Two Thin Wings of Low Aspect Ratio Determined From Surface Pressure Measurements Obtained in Flight at Mach Numbers from 0.73 to 1.90. NASA Memo 5-1-59H, May 1959, 87H25805. The effects of Mach number and angle of attack on the flow about the X-3 wing (4.5-percent thickness ratio, aspect ratio 3.09) and the X-1E wing (4-percent thickness ratio, aspect ratio 4.0) as determined from surface pressure measurements are presented. The effect of the flow behavior on the section normal-force and moment coefficients of the two wings is also discussed. The detailed survey of pressures from which the data were selected is available in tabular form from the National Aeronautics and Space Administration. 240. Martin, James A.: Determination of Local Skin Friction by Means of Surface Total-Pressure Probes. OTP-1959, May 1959, 87H30092. Work by J. F. Naleid has further extended the method of Preston to the case of compressible flow with adverse pressure gradient and zero heat transfer. Tests were performed at M = 2. 241. Finch, Thomas W.: Results of the First X-15 Flight. Presented at the IAS Summer Meeting, Los Angeles, California, June 17, 1959. 242. McTigue, J. G.; Overton, J. D.; and Petty, G., Jr.: Two Techniques for Detecting Boundary-Layer Transition in Flight at Supersonic Speeds and at Altitudes Above 20,000 Feet. NASA TN D-18, August 1959, 78N78571, 87H26761. The location of transition was measured on a supersonic fighter-type airplane by resistance-thermometer and sublimation techniques. Application of these techniques required the use of only the external surface without disturbing the internal structure. Agreement between the two methods as achieved throughout this program is discussed. Also presented are possible extensions of the program to higher Mach numbers. 243. Matranga, Gene J.; and Armstrong, Neil A.: Approach and Landing Investigation at Lift-Drag Ratios of 2 to 4 Utilizing a Straight-Wing Fighter Airplane. NASA TM X-31, August 1959, 87H25161, 62N71855. A series of landings were performed with a straight-wing fighter airplane to evaluate the effect of low lift-drag ratios on landing. Landings with peak lift-drag ratios as low as 3 were achieved by altering the airplane configuration (extending speed brakes, flaps, and gear and reducing throttle setting). 244. McKay, James M.: Measurements of GroundReaction Forces and Vertical Acceleration at the Center of Gravity of a Transport Airplane Taxiing Over Obstacles. NASA TN D-22, September 1959, 89N70669, 87H26775. Results are presented of the effects of ground speed and obstacle width on the vertical and rearward drag groundreaction forces, the vertical acceleration at the center of gravity of the airplane, the shock-strut displacement, and the dynamic response of the upper mass of the airplane structure. The obstacles were 3.0 inches in height and 1 and 4 feet in width, and the investigation covered a range of ground speeds from 12 to 86 knots. 245. Saltzman, Edwin J.: Flight Investigation of the Effect of Distributed Roughness of Skin Drag of a FullScale Airplane. NASA TM X-36, September 1959, 87H25176. The change in drag caused by the addition of two sizes of distributed sand-type roughness to the wings and tail surfaces of a delta-wing airplane has been measured at Mach numbers near 0.8 and 1.1. The largest roughness, 0.006-inch mean effective diameter, caused an increase in drag comparable to the increase predicted by the low-speed drag law for a rough plate. The increase in drag caused by the addition of the smallest roughness, 0.002-inch mean effective diameter, was less than half that predicted by the low-speed drag law for a rough plate. 246. Day, Richard E.; and Reisert, Donald: Flight Behavior of the X-2 Research Airplane to a Mach 54 Number of 3.20 and a Geometric Altitude of 126,200 Feet. NASA TM X-137, September 1959, 87H25418. Flight-test data obtained with the X-2 research airplane are presented for the maximum performance flights. An analysis is made of the instability leading to the loss of the airplane. 247. Finch, Thomas W.; and Matranga, Gene J.: Launch, Low-Speed, and Landing Characteristics Determined from the First Flight of the North American X-15 Research Airplane. NASA TM X-195, September 1959, 62N72019, 87H25292. The primary areas of emphasis in the results of the first flight of the X-15 research airplane are the launch and landing characteristics. The launch characteristics were satisfactory and, in general, were qualitatively predicted by wind-tunnel studies. The landing characteristics were predicted qualitatively by analog and flight program, and the recommended technique of extending the flaps and gear at a minimum altitude appears to be a satisfactory method of landing the X-15 airplane. 248. Marcy, W. L.: Flight Investigation of Loads on a Tee-Tail at Transonic and Supersonic Speeds. NASA TM X-57, September 1959, 71N73446, 87H25224 Horizontal- and vertical-tail loads were measured on a supersonic fighter airplane. Lift-curve slopes, spanwise centers of load, and interference effects are shown for Mach numbers from 0.81 to 2.06 at altitudes from 20,000 to 55,000 feet. Flight results agreed well with wind-tunnel data and theoretical calculations. It was concluded that calculations of a preliminary-design type are adequate for the prediction of loads on tee-tails at small angles of attack and sideslip up to a Mach number of 2. 249. Jordan, Gareth H.: Some Aspects of Shock-Wave Generation by Supersonic Airplanes. AGARD Report 251, Flight Test Panel of AGARD, Aachen, Germany, September 21–25, 1959. 250. Keener, Earl R.: Pressure Measurements Obtained in Flight at Transonic Speeds for a Conically Cambered Delta Wing. NASA TM X-48, October 1959, 87H25203, 65N12688. Pressures were obtained over the wing of the Convair JF-102A airplane at Mach numbers up to 1.19. The wing has aspect and taper ratios of 2.08 and 0.023, respectively, and incorporates two fences, a reflexed tip, and an elevon-control surface. Wing Reynolds number varied between 23 X 10 sup 6 and 58 X 10 sup 6. The results are analyzed with regard to the effects of camber on the distribution of leading-edge pressures, on the span-load distribution, and on the flowseparation characteristics. Tabulated data are available upon request from the Administration. National Aeronautics and Space 253. McKay, James M.: Measurements Obtained During the First Landing of the X-15 Research Airplane. NASA TM X-207, October 1959, 87H25342, 62N72031. One purpose of the first glide first of the X-15 research airplane was to evaluate the effectiveness of the landing-gear system. Some results are presented of the landing-approach characteristics, the impact period, and the runout phase of the landing maneuver. 254. Walker, Joseph A.: Some Concepts of Pilot’s Presentation. Presented at the SETP Symposium, October 8–10, 1959, OTP-1959, CC-H-149, 1959, 87H30656. Contents include a discussion of the pilot’s instrument presentation and control system. The X-15 is used as an example. 255. Finch, Thomas W.: X-15 Flight Test Program and Significant Flight Results. Scientific Advisory Board Panel on Aerospace Sciences, October 23, 1959. 256. Day, Richard E.: Training Considerations During the X-15 Development. Presented to the National Security Industrial Association Training Advisory Committee Meeting, Los Angeles, California, November 17, 1959, NASA CC-H-157 OTP-1959, 1959. This paper reviews briefly some of the early uses of pilot training aids in research investigations. A pertinent flight trajectory of the X-15 research airplane is summarized and the various training aids that have been, and are being, used in preparing the pilot for flying this trajectory are indicated. 257. Campbell, George S.; and Weil, Joseph: The Interpretation of Nonlinear Pitching Moments in Relation to the Pitch-Up Problem. NASA TN D-193, December 1959, 89N70901, 87H26510. Equations are presented for calculating the dynamic behavior of airplanes having arbitrarily nonlinear aerodynamic characteristics. Application of the methods derived is directed toward a study of some of the factors affecting the severity of pitch-up motions encountered by airplanes having regions of reduced stability: pitching-moment shape, control movement, dynamic pressure, airplane moment of inertia, and aerodynamic damping. Brief consideration is also given to the effectiveness of automatic stabilization devices in reducing pitch-up severity. 258. Love, James E.; and Stillwell, Wendell H.: The Hydrogen-Peroxide Rocket Reaction-Control System for the X-1B Research Airplane. NASA TN D-185, December 1959, 89N70595, 87H26473. A fixed-thrust on-off H2O2-rocket reaction-control system was installed in the X-1B research airplane as a means of control at high altitude. The design considerations, fabrication, installation, and ground testing of this system are 55 E-2551 JF-102A Airplane 251. Nugent, Jack: Interaction of Nonsteady TwinInlet Flow and Airplane Directional Motions at a Mach Number of Approximately 1.9. NASA TM X-54, October 1959, 87H25217, 71N73436. Flight tests of a twin-duct propulsion system performed at a Mach number of about 1.9 have indicated a direct interaction between an asymmetric shock configuration at the inlets and airplane directional motion. The asymmetric shock configuration was produced at reduced mass flows and was aggravated by sideslip angle. Installation of a duct splitter plate at the engine face alleviated, but did not eliminate, the interaction phenomenon. 252. Matranga, Gene J.; and Menard, Joseph A.: Approach and Landing Investigation at Lift-Drag Ratios of 3 to 4 Utilizing a Delta-Wing Interceptor Airplane. NASA TM X-125, October 1959, 87H25396. A series of landings was performed with a delta-wing interceptor airplane, having an average test wing loading of 35 pounds per square foot, to evaluate the effect of landing at low lift-drag ratios. Landings with peak effective lift-drag ratios as low as 3.75 were achieved by altering the airplane configuration. The reduction in lift-drag ratio resulted in an increase in the pertinent flare parameters. The pilot also flew several calculated landing patterns which he reported were easy and comfortable to fly. described. The operational characteristics and some problems associated with system operations are discussed. 259. Beeler, De Elroy: The Supersonic Transport. A Technical Summary. OTP-1959, 1959. Contents: State of the art - performance. Some noise problems of the supersonic transport. Structures and materials problems associated with supersonic transports. Structural loads on supersonic transports. Flying qualities of supersonic transports. Runway and braking requirements. Airway traffic control and operations. Variable geometry for transports. Effect of variable sweep on structural weight. Possible performance improvements. Proposed groundsimulation studies and flight investigation pertinent to the supersonic transport. 260. Walker, Joseph A.: Piloting Research Aircraft. OTP-1959, 1959, 87H29770. This paper cites some examples of control problems gained from the experience of the author in flying advanced research and production aircraft at the NACA High-Speed Flight Station. 263. Reed, Robert D.: Vertical-Tail Loads Measured in Flight on Four Airplane Configurations at Transonic and Supersonic Speeds. NASA TN D-215, February 1960, 89N70902, 87H26629. Aerodynamic loads were obtained from the X-5 airplane, an F-100 with a small vertical tail, F-100 with a large vertical tail, and the X-1E airplane. Effects of sideslip angle, rudder deflection, and Mach number are presented for trim flight at altitudes from 40,000 feet to 70,000 feet and at Mach numbers from 0.70 to 2.08. Comparisons are made with simple theoretical methods of estimating the loads. Also, the total directional stability for each airplane as calculated by use of vertical-tail loads is shown. 264. Walker, Joseph A.: A Pilot’s Look at the X-15 Program. Presented at the IAS, Texas Section, Dallas, Texas, April 28–30, 1960. OTP-1960 265. Armstrong, N. A.: X-15 Operations: Electronics and the Pilot. Astronautics, vol. 5, no. 5, May 1960, pp. 42–43, 76–78. Electronic equipment figures prominently in X-15 flight and ground systems, but this hypersonic vehicle is an instrument of the pilot, depending on him for control and flight success. 266. Bellman, Donald R.; and Toll, Thomas A.: Aeronautical Operating Problems—Supersonic Transport. Presented at the NASA Conference, June 2–3, 1960. 267. Weil, Joseph; and Matranga, Gene J.: Review of Techniques Applicable to the Recovery of Lifting Hypervelocity Vehicles. NASA TM X-67563. Presented at the Joint Conference on Lifting Manned Hypervelocity and Reentry Vehicles, Part 1, pp. 313–328. Langley Research Center, April 11–14, 1960. Also Scientific Advisory Board, Boeing, Seattle, Washington, and IAS Symposium on Recovery of Space Vehicles, Los Angeles, California, August 31–September 1, 1960, (see N72-70963 06-99), 72N70983. 268. Walker, Harold J.; and Wolowicz, Chester H.: Theoretical Stability Derivatives for the X-15 Research Airplane at Supersonic and Hypersonic Speeds Including a Comparison With Wind-Tunnel Results. NASA TM X-287, August 1960, 87H25655, 65N24060, #. The hypersonic small-disturbance theory for lifting surfaces and the second-order shock-expansion method for bodies of revolution are employed in conjunction with the results from slender-body and linear theory to predict the longitudinal and lateral-directional stability and control derivatives for angles of attack from 0 degrees to 25 degrees and for Mach numbers extending well beyond the airplane design limit. The results are compared with available wind-tunnel data at Mach numbers from approximately 2 to 7 and with the limiting values given by Newtonian theory. Good agreement is 56 1960 Technical Publications 261. Anon.: Aerodynamic and Landing Measurements Obtained During the First Powered Flight of the North American X-15 Research Airplane. NASA TM X-269, January 1960, 62N72093. The first powered flight of the North American X-15 research airplane was performed on September 17, 1959. A Mach number of 2.1 and an altitude of 52,000 feet were attained. Static and dynamic maneuvers were performed to evaluate the characteristics of the airplane as subsonic and supersonic speeds. Data from these maneuvers as well as from the launch and landing phases are presented, discussed, and compared with predicted values. 262. Larson, Terry J.; and Washington, Harold P.: Summary of Rawinsonde Measurements of Temperatures, Pressure Heights, and Winds Above 50,000 Feet Along a Flight-Test Range in the Southwestern United States. NASA TN D-192, January 1960, 89N70968, 87H26504 Yearly, seasonal, diurnal, and 24-hour variations of temperatures, pressure heights, and winds in the 100-millibar to 2-millibar pressure range from rawinsonde data of four stations along a flight-test range in the southwestern United States are presented. This range, referred to as High Range, extends from Edwards, Calif., to Wendover, Utah, to accommodate flight operations of the X-15 and other highperformance craft. Comparisons of average temperatures and pressure heights are made with those of the U.S. Extension to the ICAO Standard Atmosphere. obtained for the most part, but several notable differences are found among the lateral- directional derivatives at high angles of attack. 269. Holleman, E. C.; and Sadoff, M.: Simulation Requirements for the Development of Advanced Manned Military Aircraft. NASA TM X-54672, Presented at the IAS National Meeting, San Diego, California, August 3, 1960, 75N72588. The present state of the art of the piloted flight simulator leaves no major deterrent to the mechanization of required simulators for the design of present or future manned military airplanes. The fixed-base simulator with adequate presentation and controls is completely satisfactory for the investigation of a wide range of airplane problems. However, there are some areas which require some form of motion stimulus. Other areas remain where simulator requirements are not yet resolved, but work is continuing to better define these simulator requirements. 270. Fischel, Jack: The X-15 Flight Research Program in Relation to the Development of Advance Military Aircraft. Presented at the IAS National Meeting, San Diego, California, August 1–3, 1960. 271. Bikle, Paul F.: Initial Remarks About the X-15 Flight Research Program. Presented at the 3rd Annual West Coast Meeting of American Astronautics Society, Seattle, Washington, August 9, 1960. 272. Day, Richard E.: Training Aspects of the X-15 Program. In The Training of Astronauts, Report of a Working Group Conference: Panel on Psychology, Armed Forces-NRC Committee on Bio-Astronautics, Woods Hole, Massachusetts, August 29–30, 1960, pp. 5–14, 62N12371. Various training aids in the development of the X-15 program are presented. Future flight data obtained in more critical control areas will afford the unique opportunity to assess the true value of these training aids for the X-15 and to establish training requirements for future vehicles. 273. Baker, Thomas F. Dyna-Soar I—Flight-Data Objectives. Presented at the IAS Symposium on Recovery of Space Vehicles, Los Angeles, California, August 31– September 1, 1960. 274. Nugent, Jack; and Powers, Bruce G.: Flight Tests of a Twin-Duct Induction System for a Mach Number Range of 0.78 to 2.07. NASA TM X-281, September 1960, 87H25628. Time histories of several airplane, engine, and inductionsystem parameters are presented for maneuvers made at altitude of 26,000 feet, 40,000 feet, and 55,000 feet over a 57 Mach number range of 0.78 to 2.07. A time history of a pushdown-turn maneuver at 40,000 feet is also presented. Comparisons of the data were made to show the effects of angle of attack. Mach number, altitude, and duct bypass area on the induction system parameters. 275. Weil, Joseph; and Matranga, Gene J.: Review of Techniques Applicable to the Recovery of Lifting Hypervelocity Vehicles. NASA TM X-334, September 1960, 87H26028. A general review of piloting problems concerned with the recovery phase of lifting hypervelocity vehicles is presented. A short discussion is offered pertinent to the maneuvering capabilities and piloting techniques applicable to the initial approach phase of gliders with low lift-drag ratios. The principal emphasis concerns factors affecting the final approach and landing operation of these gliders. The results of general flight studies as well as recent experience obtained in the approach and landing of the X-15 research airplane are reviewed. Finally, a definition of the limits of piloted flared landings is developed. 276. Tambor, Ronald: Flight Investigation of the Lift and Drag Characteristics of a Swept-Wing, Multijet, Transport-Type Airplane. NASA TN-D-30, (Corrected Copy), September 1960. The lift and drag characteristics of a Boeing KC-135 airplane were determined during maneuvering flight over the Mach number range from 0.70 to 0.85 for the airplane in the clean configuration at an altitude of 26,000 feet. Data were also obtained over the speed range of 130 knots to 160 knots at 9,000 feet for various flap deflections with gear down. 277. Stillwell, W. H.; and Larson, T. J.: Measurement of the Maximum Speed Attained by the X-15 Airplane Powered With Interim Rocket Engines. NASA TN D-615 September 1960, 62N71189, 87H27079. On August 4, 1960 a flight was made with the X-15 airplane to achieve the maximum speed possible with two interim engines. Presented are the details of the techniques utilized to determine the maximum Mach number (3.31 plus or minus 0.04) and speed (2,196 mph plus or minus 35) attained. 278. Wolowicz, C. H.; Drake, H. M.; Videan, E. N.; Morris, G. J.; and Stickle, J. W.: Simulator Investigation of Controls and Display Required for Terminal Phase of Coplanar, Orbital Rendezvous. NASA TN D-511, October 1960, 62N71085, 87H27353. A simulator study was made of presentations and control requirements for a manned astrovehicle employed for the interception of artificial satellites during the terminal phase of an orbital rendezvous. Two oscilloscope and one directvisual-observation presentation and three control modes were investigated. The study was considered in terms of a manned interceptor having a home berth at a manned space station which is in circular orbit 500 miles above the earth. Interceptions were restricted to coplanar conditions. 279. Stillwell, W. H.; and Larson, T. J.: Measurement of the Maximum Altitude Attained by the X-15 Airplane Powered With Interim Rocket Engines. NASA TN D-623, October 1960, 62N71197, 87H27097. On August 12, 1960, an X-15 flight was made to achieve essentially the maximum altitude expected to be possible with the interim rocket engines. Presented are the details of the techniques utilized to determine the maximum geometric altitude of this flight (136,500 ft plus or minus 600). 280. Beeler, De E.: The X-15 Research Program. AGARD Report 289, Tenth Annual General Assembly of AGARD, Istanbul, Turkey, October 3–8, 1960. Brief summary of the research program which led to the X-15 aircraft, and some of the early flight tests. Includes flight regimes, mission, research areas, shock-wave patterns at hypersonic speeds, simulator support, stability derivatives, control effectiveness derivatives, drag characteristics poweroff, heat transfer, wing temperature station, wing pressure distribution midspan station, wing loads, maximum speed and altitude, and damping. 281. White, R. M.; and Walker, Joseph A.: X-15 Program Status Report, Parts 1 and 2. Proceedings from SETP Annual Symposium, Vol. 5, No. 2, October 7, 1960. Contains a brief resume of four flights leading to, and including, the maximum speed flight of the X-15 with the interim engine, then discusses some of the results to date, and concludes with a summary of the intended flight program with the design XLR-99 engine. 282. Johnson, Clinton T.: Investigation of the Characteristics of 6-Foot Drogue-Stabilization Ribbon Parachutes at High Altitudes and Low Supersonic Speeds. NASA TM X-448, November 1960, 87H25617. Performance data are presented for two types of ribbon parachutes. The parachutes were forcibly deployed from an air-launched test vehicle at altitudes from 55,000 feet to 70,000 feet and at Mach numbers between 0.92 and 1.52. Opening shock, steady-state drag performance, and canopyporosity effects are evaluated with respect to Mach number and dynamic pressure. 283. Smith, Harriet J.: Experimental and Calculated Flow Fields Produced by Airplanes Flying at Supersonic Speeds. NASA TN D-621, November 1960, 62N71195, 87H27094. 58 Results are presented of a flight investigation conducted to survey the flow field generated by airplanes flying at supersonic speeds. The pressure signatures of an F-100, an F-104, and a B-58 airplane, representing widely varying configurations, at distances from 120 to 425 feet from the generating aircraft and at Mach numbers from 1.2 to 1.8 are shown. Calculations made by using Whitham’s method gave good agreement with experimental results. A procedure for calculating the F(y) function used in Whitham’s method is also given. 284. Andrews, W. H.; and Holleman, E. C.: Experience With a Three-Axis Side-Located Controller During a Static and Centrifuge Simulation of the Piloted Launch of a Manned Multistage Vehicle. NASA TN D-546, November 1960, 62N71120, 87H26912. The control problems associated with piloting multistaged vehicles to orbital conditions were investigated with static and dynamic simulators. A three-axis controller was used for primary control. Presented are design details of the controller, pilot opinions concerning its operation, and other data pertinent to the design and use of a controller of this type. 285. Saltzman, Edwin J.: Preliminary Full-Scale Power-Off Drag of the X-15 Airplane for Mach Numbers From 0.7 to 3.1. NASA TM X-430, December 1960, 87H25514, 62N72254. These drag data provide preliminary means of appraising present methods of extrapolating wind-tunnel drag results to full scale at Mach numbers up to 3. Estimated drag values based on wind-tunnel measurements are included. Freestream Reynolds numbers range from 13.9 x 10 (sup) 6 to 28 x 10 (sup) 6, based on the mean aerodynamic chord. 286. Holleman, E. C.: Utilization of the Pilot During Boost Phase of the Step 1 Mission. In its Joint Conference on Lifting Manned Hypervelocity and Reentry Vehicles, Part 2 1960, pp. 261–272, (see N72-71002 06-99), 72N71021. 287. Baker, T. F.; Russell, H. G.; and Schofield, B. L.: Dyna-Soar Step 1 Flight Test Program. In Joint Conference on Lifting Manned Hypervelocity and Reentry Vehicles, Part 2, 1960, pp. 311–324, (see N72-71002 06-99), 72N71025. 288. Truszynski, G. M.; and Lindfors, P. O.: Instrumentation and Communications Considerations. In its Joint Conference on Lifting Manned Hypervelocity and Reentry Vehicles, Part 2, 1960, pp. 325–334, (see N72-71002 06-99), 72N71026. 289. Holleman, Euclid C.; Armstrong, Neil A.; and Andrews, William H.: Utilization of the Pilot in the Launch and Injection of a Multistage Orbital Vehicle. Presented at the IAS 28th Annual Meeting, New York, New York, January 25–27, 1960, 87H30992. The Flight Research Center has conducted fixed-base and centrifuge simulator programs to investigate the capabilities of the pilot in providing control and guidance during launch. For the centrifuge program an attempt was made to minimize the adverse effects of acceleration on the pilot by designing a molded seat and three-axis controller that were functionally independent of acceleration. The effects of staging acceleration environment on the pilot’s performance were determined to 15g, and the control task associated with twoand four-stage vehicle launches were investigated. 290. Thompson, Milton O.: Piloting Aspect of Project ALSOR. For publication in SETP newsletter, 1960. X-15 Launch From a B-52 Airplane The ALSOR program is inclined toward providing current information in a reliable and efficient manner to support the X-15 program. 293. Holleman, Euclid C.; and Reisert, Donald: Controllability of the X-15 Research Airplane With Interim Engines During High-Altitude Flights. NASA TM X-514, March 1961, 87H25871. A peak geometric altitude of 136,500 feet with a minimum dynamic pressure of 10.6 lb/sq. ft. was attained with only the aerodynamic controls available to the pilot for controlling and stabilizing the airplane. Aerodynamic control was adequate throughout the flight, but at minimum dynamic pressure the airplane was lightly damped, which made precise control difficult. Because of the transient nature of the trajectory and the negligible load factors associated with the airplane oscillation, the pilot did not object to the poor dynamic characteristics of the airplane under these conditions and could satisfactory control the airplane along the trajectory. 294. McKay, J. M.; and Scott, B. J.: Landing-Gear Behavior During Touchdown and Runout for 17 Landings of the X-15 Research Airplane. NASA TM X-518, March 1961, 63N12563, #. Data are presented for the pretouchdown conditions, the impact period, and the runout phase of the landing for vertical velocities up to 9.5 feet per second and true ground speeds between 145 knots and 238 knots. The dynamic response of the airplane during the impact period is presented in the form of time histories of shock strut force and displacement, maingear and nose-gear drag forces, upper-mass acceleration, horizontal-tail setting, horizontal-tail load, airplane angle of attack, and pitching velocity. Also included is the variation of the coefficient of friction of the main-gear skid with ground speed during the runout of a typical landing. 295. Day, Richard E.: X-15 Simulation and the X-15 Flight Program. Presented to the National Academy of Sciences, Panel on Acceleration Stress, ARC, March 11, 1961. 59 E-4942 1961 Technical Publications 291. Reed, Robert D.; and Watts, Joe D.: Skin and Structural Temperatures Measured on the X-15 Airplane During a Flight to a Mach Number of 3.3. NASA TM X-468, January 1961, 87H25738. A survey of skin and structural temperatures was obtained on the X-15 airplane during a flight to a Mach number of 3.3. Fuselage, wing, horizontal-tail, and vertical-tail temperatures are presented to show temperature variations on the external surfaces and temperature differences between the skin and internal structure. The maximum temperature recorded was 440 °F on an unsupported skin area on the lower vertical tail. Temperature differences of 400 °F were recorded between the external skin and internal spar webs on the wing. Local external temperature differences caused by the heat-sink effect of the supporting structure were as great as 220 °F. Temperature indicating paint aided in identifying the location of areas of concentrated heating on the lower surface of the wing. 292. Matranga, Gene J.: Launch Characteristics of the X-15 Research Airplane as Determined in Flight. NASA TN D-723, February 1961, 62N71297, 87H27349. The first 16 air launches of the X-15 airplane demonstrated the feasibility of air launch from an asymmetric position under the wing of the B-52 carrier airplane. Use of the stability augmentation system markedly reduced the launch transients. Reasonable agreement exists between flight and predicted data. A general discussion of the role played by various simulators in the development of the X-15 program. 296. Banner, R. D.; and *Kinsler, M. R.: Status of X-15 Aerodynamic-Heating Studies. ARS Paper-1629-61. Presented at the ARS Missile and Space Vehicle Testing Conference, Los Angeles, California, March 13–16, 1961, 63N15312. One of the primary purposes of the X-15 program is to obtain full-scale aerodynamic-heating information that can be used to establish the adequacy of current theoretical methods and model tests. In conjunction with this purpose, a special flight was performed which maximized the heating rates and minimized the transient flight conditions. This flight reached a Mach number of 3.1. Results indicated a reasonable agreement between measured heat transfer data and simple theoretical predictions. Boundary-layer-transition data were obtained which pointed out a continuing problem of prediction that should probably be treated with conservatism until more detailed information is obtained. *North American Aviation, Downey, California. 297. Reisert, Donald; and Adkins, Elmor J.: Flight and Operational Experiences With Pilot-Operated Reaction Controls. ARS Paper-1674-61. Presented at the ARS Missile and Space Vehicle Testing Conference, Los Angeles, California, March 13–16, 1961. 298. Finch, Thomas W.: X-15 Flight Program. Presented at The ARS Missile and Space Vehicle Testing Conference, Los Angeles, California, March 13–16, 1961. This paper discusses a number of program phases that have been completed, including the contractor’s flight program and the Government research program utilizing an interim-engine configuration giving an indication of its experiences and a description of future program plans. 299. Taylor, Lawrence W., Jr.; and Smith, John W.: An Analytical Approach to the Design of an Automatic Discontinuous Control System. NASA TN D-630, April 1961, 62N71204, 87H27116. The design of a attitude-stabilization system for a vehicle experiencing negligible external moments is investigated analytically. A discontinuous control system employing a linear switching function and having a neutral zone and time delays is studied and equations are developed to generalize and optimize the system’s transient and limit-cycle performance. Example systems which can minimize power required, attitude error, and angular-velocity error within a specified period of operation are included 60 300. Kordes, Eldon E.; and Noll, Richard B.: Flight Experience of Panel Flutter. ARS Lifting Reentry Vehicles, Structures, Materials, and Design Conference, Palm Springs, California, April 4–6, 1961. Presents X-15 panel flutter data and compares with flutter boundary established by wind tunnel tests. 301. Weil, Joseph: Application of Analytical Techniques to Flight Evaluations in Critical Control Areas. AGARD Report No. 369, AGARD Specialist Meeting on Aircraft Stability and Control, Brussels, Belgium, April 10–14, 1961. Flight data can be dangerously misleading in the absence of careful interpretation. This report discusses test results pertinent to a variety of typical flight-control problem areas of the current generation of airplanes. The results presented were obtained from flight investigations of many research and operational aircraft at the NASA Flight Research Center over the past 10 years. The report considers basic stability problems such as pitch-up, roll coupling, and marginal directional stability. Development of augmentation systems and control system evaluations are also discussed in some detail. Throughout the report, the importance of coordinating flight and similar results are very much stressed and it is shown that in many areas even the most painstaking interpretation of flight data can lead to possible disaster if flight tests are not adequately supported by simulator studies using realistic stability and control derivatives. 302. Weil, Joseph; and Adkins, E. J.: Review of Selected X-15 Development and Operating Experiences. Presented at the ISA Aero-Space Instrumentation Symposium, Dallas, Texas, April 30–May 4, 1961. This paper reviews the flight-control simulation experiences of X-15 systems and components. The roll of simulators and auxiliary aircraft is discussed, with emphasis on pilot training and the development of flight procedures necessary to obtain the maximum research return on the investment. 303. Weil, Joseph; and Adkins, E. J.: Utilization of Aircraft as Systems and Flight Control Test Beds. Presented at the ISA Aero-Space Instrumentation Symposium, Dallas, Texas, May 4, 1961. 304. Taylor, Lawrence W., Jr.; and Day, Richard E.: Flight Controllability Limits and Related Human Transfer Functions as Determined From Simulator and Flight Tests. NASA TN D-746, May 1961, 87H27401. A simulator study and limited flight tests were performed to determine the levels of static stability and damping necessary for pilot control of the pitch, roll, and yaw attitudes of a vehicle for a the pilot to control the airplane at conditions that were otherwise uncontrollable. The influence on the controllability limits of the more important aerodynamic coefficients and other factors, such as learning and interruption of the pilot’s display, was also investigated. Information concerning human transfer functions applicable to marginally controllable tasks is presented which should aid in assessing the controllability of any specific configuration. 305. Matranga, Gene J.: Analysis of X-15 Landing Approach and Flare Characteristics Determined From the First 30 Flights. NASA TN D-1057, July 1961, 62N71631, 87H27661. This paper presents lift, drag, and angle-of-attack data for various approach and landing configurations. The conditions and problems encountered during the approach pattern and the flare to touchdown are discussed. The value of flight simulations of the approach and flare maneuvers is assessed. 306. Saltzman, Edwin J.: Preliminary Base Pressures Obtained From the X-15 Airplane at Mach Numbers From 1.1 to 3.2. NASA TN D-1056, August 1961, 62N71630, 87H27659. Base pressure measurements have been made on the fuselage, 10 degrees-wedge vertical fin, and side fairing of the X-15 airplane. Data are presented for Mach numbers between 1.1 and 3.2 for both powered and unpowered flight. Comparisons are made with data for small-scale-model tests, semiempirical estimates, and theory. 307. Jordan, G. H.; McLeod, N. J.; and Ryan, B. M.: Review of Flight Measurements of Sonic Booms and Effects of Shock Waves on Other Aircraft. Presented at 5th Annual Symposium of the Society of Experimental Test Pilots, Beverly Hills, California, September 29–30, 1961, 63N81390. 308. Andrews, William H.; Cooney, Thomas V.; and Fischel, Jack: Contributions from the X-15 Flight Test Program Relating to Design and Development of the Supersonic Transport. SETP 62-9686. Presented at 5th Annual Symposium of the Society of Experimental Test Pilots, Beverly Hills, California, September 29–30, 1961. Discussion of information and experience gained with the X-15 which is applicable to supersonic transport design, and of means by which future data may be provided. Extensive treatment is given to (1) structural heating, including temperature, temperature gradients, and temperature effects on structures; (2) structural dynamics, including a brief resume of flutter experience with the X-15; and (3) augmentation systems, including fundamental problems experienced and the reliability of the system. The air-datasensing systems employed in the X-15 and the related system-measurement accuracies obtained are covered. The 61 reliability of the subsystems and the maintenance of such items as the environmental control system and the hydraulic system are discussed. Future programs involving fundamental studies requiring specialized instrumentation not presently installed in the X-15 are considered. 309. Taylor, Lawrence W., Jr.: Analysis of a PilotAirplane Lateral Instability Experienced With the X-15 Airplane. NASA TN D-1059, November 1961, 62N71633, 87H27667. By using an experimentally developed human transfer function for the pilot and system-analysis methods, the pilotairplane lateral instability observed with the X-15 airplane is analyzed. The methods used adequately explain the lateralcontrol problem and can be used to predict the problem. The calculated area of lateral-control difficulty agreed with that determined on the X-15 piloted flight simulator and with flight data. 310. Beeler, De E.: Recent X-15 Flight Results. Flight Mechanics Panel, Paris, France, November 21–24, 1961. 311. Anon.: Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project. NASA TM X-57072, 1961, (see N71-75444), 71N75443. This document is a compilation of the papers presented at the Conference on the Progress of the X-l5 Project held at the NASA Flight Research Center, Edwards Air Force Base, California, November 20–21, 1961. This conference was held by the Research Airplane Committee of the U.S. Air Force, the U.S. Navy, and the National Aeronautics and Space Administration to report on the technical status of this research airplane. The papers were presented by members of the staffs of North American Aviation, Inc.; Aeronautical Systems Division, U.S. Air Force; Air Force Flight Test Center; and National Aeronautics and Space Administration. 312. Beeler, D. E.; and Toll, T. A.: Status of X-15 Research Program. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, Edwards AFB, California, November 20–21, 1961, pp. 1–10, 71N-75444. 313. Banner, R. D.; Kuhl, A. E.; and Quinn, R. D.: Preliminary Results of Aerodynamic Heating Studies on the X-15. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 11–28, (see N71-75443), 71N75445. (See also 341.) The results of the preliminary flight heat-transfer studies on the X-l5 airplane are presented, together with a discussion of the manner in which the data have been obtained, a comparison of measured and calculated turbulent heattransfer coefficients, a correlation of the model test results and the flight results for turbulent heat transfer; some information on boundary-layer transition, and a comparison of measured and calculated skin temperatures at several locations on the airplane. 314. Kordes, Eldon E.; Reed, Robert D.; and *Dawdy, Alpha L.: Structural Heating Experiences of the X-15. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 29–45, (see N71-75443), 71N75446. The expected structural temperatures and their effect on the development and design of the X-l5 airplane structure have been described in previous conferences, and Banner, Kuhl, and Quinn (paper no. 2) have discussed in detail the many factors affecting the heat input to the structure. The purpose of the present paper is to show the magnitude of structural temperatures measured during the flight program and to describe structural problems that have developed due to structural heating. *North American Aviation, Inc., Downey, California. 315. Jordan, G. H.; McLeod, N. J.; and *Guy, L. D.: Structural Dynamic Experiences of the X-15ResearchAirplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 47–59, (see N71-75443), 71N75447. This paper reviews the structural dynamics problems that influenced the design of the structure and discusses the experiences that have been encountered during the flight tests. *NASA Langley Research Center, Hampton, Virginia. 316. McKay, J. M.; and Kordes, E. E.: Landing Loads and Dynamics of the X-15 Airplane. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, 1961, pp. 61–71, (see N71-75443), 71N75448. (See also 342.) Because the landing-gear configuration represents a marked departure from previously used configurations, the present paper has been prepared to report on the landing loads experience of the X-15 A further purpose of this paper is to review the dynamics of landing and to present results of a recent theoretical study of the effects of various parameters on the landing loads. 317. Keener, E. R.; and Pembo, C.: Aerodynamic Forces on Components of the X-15. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, 1961, pp. 73–82, (see N71-75443), 71N75449. (See also 343.) An attempt has been made in the flight research program to verify some of the force measurements with both pressure and strain-gage measurements. This paper presents a summary of the flight force data obtained to date. The data are compared with the wind-tunnel results and with some of the more familiar theoretical methods and approximations. 318. *Hopkins, E. J.; **Fetterman, D. E., Jr.; and Saltzman, E. J.: A Comparison of Full-Scale X-15 Lift and Drag Characteristics with Wind-Tunnel Results and Theory. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 83–98, (see N71-75443), 71N75450. (See also 344.) Data on the lift and drag characteristics of the X-15 airplane obtained in flight are shown to be in agreement with windtunnel-model data for Mach numbers up to 5. Existing theoretical methods are indicated to be adequate for estimating the X-l5 minimum drag but underestimated the drag due to lift and overestimated the maximum lift-drag ratio. Two-dimensional theory is shown to be adequate for predicting the base pressures behind surfaces having very blunt trailing edges) such as those on the vertical tail of the X-l5. *NASA Ames Research Center, Moffett Field, California. **NASA Langley Research Center, Hampton, Virginia. 319. Walker, H. J.; and Wolowicz, C. H.: Stability and Control Derivative Characteristics of the X-15. ResearchAirplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 99–112, (see N71-75443), 71N75451. (See also 345.) The flight-determined derivative characteristics are compared with the predictions from wind-tunnel tests and theory for Mach numbers extending to 5.5 and angles of attack up to 17 degrees. With few exceptions, the predictions were found generally to be in good agreement with the flight data. Areas of deficient stability and control are briefly discussed. 320. White, R. M.; Robinson, G. H.; and Matranga, G. J.: Resume of X-15 Handling Qualities. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, 1961, pp. 113–130, (see N71-75443), 71N75452. (See also 346) The handling qualities of the X-15 research airplane have been assessed from pilot’s opinions, with verification in many cases by data acquired during flights. Areas of interest covered are the launch, climbout, ballistic, reentry, and landing phases of flights made to date. 321. Petersen, F. S.; Rediess, H. A.; and Weil, J.: Lateral Directional Control Characteristics. Research-Airplane- 62 Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 131–154, (see N71-75443), 71N75453. (See also 347.) The deterioration of lateral directional controllability with roll damper off and the pilot performing a lateral control task is explained. The problem area was defined by fixed-base and airborne simulators and verified by closed-loop analysis in which a human transfer function represents the pilot. A parameter which will predict the problem area for the X-15 airplane is developed. The means considered to alleviate the control problem in the X-15 airplane are also discussed. 322. Hoey, R. G.; and Day, R. E.: X-15 Mission Planning and Operational Procedures. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, 1961, pp. 155–169, (see N71-75443), 71N75454. The philosophy of the X-l5 flight-test program thus far has been to expand the flight envelope to the maximum speed and design altitude as rapidly as practical and simultaneously to obtain as much detailed research data on the hypersonic environment as possible. The envelope expansion program has been performed on an incremental performance basis; that is, each successive flight is designed to go to a slightly higher speed or altitude than the previous flight, thus permitting a reasonable extrapolation of flight-test data from one flight to the next and also building a backlog of pilot experience. The mission planning and operational procedures associated with the program are discussed in this paper. The effect on flight planning of systems reliability, stability limitations, and ranging considerations are also discussed. General piloting techniques and pilot training are mentioned. 323. Taylor, L. W., Jr.; and Merrick, G. B.: X-15 Stability Augmentation System. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, 1961, pp. 171–182, (see N71-75443), 71N75455. This paper describes the basic damper system currently installed in the X-l5, discusses some of the problems encountered during its development and flight testing, and reviews briefly the system reliability. 324. *Johannes, R. P.; Armstrong, N. A.; and Hays, T. C.: Development of X-15 Self-Adaptive Flight-Control System. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 183–194, (see N71-75443), 71N75456. In-house studies conducted at Wright-Patterson Air Force Base in l956 convinced the Flight Control Laboratory, Aeronautical Systems Division, of the theoretical feasibility of designing a self-adaptive flight-control system. As the name implies, such a system would automatically adapt itself in order to provide essentially constant damping and 63 frequency of the aircraft in combination with the control system as the vehicle encountered flight conditions of varying aerodynamic control-surface effectiveness. To this end a number of study contracts were awarded in l957 which soon led to flight-test programs testing adaptive concepts in F-94 airplanes by the Massachusetts Institute of Technology and the Minneapolis Honeywell Regulator Company. Minneapolis-Honeywell continued this effort with a company funded flight-test program for testing the system in an F-101A airplane. *Aeronautical Systems Division, U.S. Air Force. 325. *Mace, W. D.; and Ball, J. L.: Flight Characteristics of X-15 Hypersonic Flow-Direction Sensor. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 195–201, (see N71-75443), 71N75457. The purpose of this paper is to discuss the experience that has been obtained through the use of the nulling ball-nose flow direction sensor during flight testing of the airplane. *NASA Langley Research Center, Hampton, Virginia. 326. *Leiby, R. G.; Bellman, D. R.; and DeMar, N. E.: XLR99 Engine Operating Experience. Research-AirplaneCommittee Report on Conference on the Progress of the X-15 Project, 1961, pp. 215–226, (see N71-75443), 71N75459. XLR99-RM-1 rocket engine, which was developed specifically for the X-15 airplane, is the largest rocket engine designed from the outset for use in a manned vehicle to be completely controlled by the crew. In order to provide the desired safety and controllability required by the X-15 mission, many unique features were included in the design. Delays in the development of the engine required that the initial X-15 flights be made with an interim engine. However, the first flight with the XLR99 was made in November 1960, and the engine has been used in government flight operations since February 1961. Since the first flight, fifteen flights have been made with the XLR99. This paper summarizes the XLR99 operating experience during the flight program. *Air Force Flight Test Center, Edwards AFB, California. 327. *Rowen, B.; *Richardson, R. N.; and Layton, G. P., Jr.: Bioastronautics Support of the X-15 Program. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 255–264, (see N71-75443), 71N75461. The techniques of air-to-ground telemetry have been used in research aircraft testing since the start of the X-l program in l946. It became apparent during the development of the X-type research aircraft that personnel responsible for aerospace medical support of the pilot were not taking full advantage of the progress in telemetry systems to monitor for medical purposes the pilot and his environment during flight. One of the research objectives of the X-l5 program is to obtain the pilot’s physiological response to flight at increased speed and altitude. This objective is accomplished with the pilot wearing a full pressure suit; therefore, this garment and biomedical data acquisition equipment, techniques, and results are discussed in this paper. *Air Force Flight Test Center, Edwards AFB, California. 328. Love, J. E.; and Palmer, J. R.: Operational Reliability Experiences with the X-15 Aircraft. ResearchAirplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 277–287, (see N71-75443), 71N75463. It is the purpose of this paper to describe a comprehensive picture of X-l5 operational reliability. The curves and text presented are based on actual parts failure records, flight logs, and the daily repair work sheets. It is therefore not only a picture of the reliability with regard to safety in flight, but also in view of ground preparation time and cost. Repeated system and component failures have resulted in many costly delays. 329. Walker, J. A.: A Pilot’s Impression of the X-15 Program. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 303–312, (see N71-75443), 71N75465. It is the intent of this paper to be critical of the X-15 because of its deficiencies or problems. It should rather be kept in mind that many compromises had to be accepted in the design of the X-15 to get on the job, and rightly so, because there are some questions which still have not been resolved. 330. Bikle, P. F.; and *Pezda, E. F.: Future Plans for the X-15. Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project, 1961, pp. 329–333, (see N71-75443), 71N75467. This third X-l5 conference has given us an opportunity to review and evaluate, in considerable detail, the progress that has been achieved in the flight research program to date. Figures l and 2 have been selected as a summary of the areas thus far explored. Similar results have been discussed in detail in the papers presented. Although it is not possible, in any one or two figures, to show the desired information for all the varied areas of interest in the program, these plots of altitude and angle of attack against velocity do represent two of the many parameters of interest, and the shaded areas demonstrate roughly the progress that has been made. It appears that most of the work originally planned is nearly completed, with perhaps 50 percent of the aerodynamics, 64 structures, heating, and bioastronautics information already obtained. *Aeronautical Systems Division, U.S. Air Force. 1962 Technical Publications 331. Yancey, R. B.; Rediess, H. A.; and Robinson, G. H.: Aerodynamic-Derivative Characteristics of the X-15 Research Airplane as Determined From Flight Tests for Mach Numbers from 0.6 to 3.4. Wolowicz, C. H.: Appendix A–Approximate Equations for Determining NASA TN D-1060, C n ,C l , and ( C n – C n ) . β β r ˙ β January 1962, 62N10089, #, 87H27669. Lateral, directional, and longitudinal stability and control derivatives are determined from flight tests of the X-15 airplane with the low-power LR11 rocket engine. Approximate relationships are developed for determining the derivatives C sub n sub Beta, C sub l sub Beta, and (C sub n sub r - C sub n sub derivative of Beta) and for isolating the effects of stability augmentation. Wind-tunnel predictions are compared with the flight-determined derivatives. 332. Kordes, Eldon E.; and Noll, Richard B.: Theoretical Flutter Analysis of Flat Rectangular Panels in Uniform Coplanar Flow with Arbitrary Direction. NASA TN D-1156, January 1962, 62N10092, #, 87H27601. Numerical calculations show that small variations in flow direction have a marked effect on the flutter of simply supported rectangular panels. The results of the calculations also show that the critical flutter mode changes at small flow angles when the length-width ratio is less than 1. Flutter conditions for a given panel at different flow angles can be compared on a common basis by use of a dynamic-pressureparameter ratio referenced to flow conditions of an aligned panel. 333. Roman, James A.: Biomedical Monitoring InFlight. Lectures in Aerospace Medicine, School of Aerospace Medicine, Aerospace Medical Division, Brooks AFB, Texas, January 8–12, 1962, pp. 97–114, 62N14203. 334. Kordes, Eldon E.; and Noll, Richard B.: Flight Flutter Results for Flat Rectangular Panels. NASA TN D-1058, February 1962, 62N10043, #, 87H27664. Panel-flutter data obtained from several different aircraft during supersonic flight are presented and compared with a previously established flutter boundary based on results from wind-tunnel tests. The flight data were obtained for rectangular panels aligned with the flow and for rectangular panels swept at 52 degrees. Some results of a flutter analysis of swept, flat, rectangular panels are presented and used to compare the flight results with the flutter boundary for aligned panels. 335. Bellman, Donald R.; and Washington, Harold P.: Preliminary Performance Analysis of Air Launching Manned Orbital Vehicles. NASA TM X-636, H-229, February 1962, 72N71616, 87H26325. A preliminary performance analysis was made to determine the capability of large subsonic and supersonic bombers for air launching manned hypersonic and satellite vehicles. The bombers considered now exist or are being developed in the United States. Four booster configurations were used in the calculations, with a winged vehicle of the Dyna-Soar type as the payload. Comparisons were made on the basis of vacuum specific impulse, burnout velocity, ratio of payload weight to launch-package gross weight, and structural weight. 336. Taylor, Lawrence W., Jr.; and *Merrick, G. B.: X-15 Airplane Stability Augmentation System. NASA TN D-1157, March 1962, 62N10587, 87H27604, #. The basic damper system currently installed in the airplane is described. Some of the problems encountered during the development and flight testing of the system are discussed, and the reliability of the system is reviewed briefly. *North American Aviation, Inc., Inglewood, California. 337. Jordan, Gareth H.; McLeod, Norman J.; and Guy, Lawrence D.: Structural Dynamics Experiences of the X-15 Airplane. NASA TN D-1158, March 1962, 62N10586, 87H27606, #. The structural dynamic problems anticipated during the design of the X-15 airplane are reviewed briefly, and the actual flight experiences with the airplane are described. The noise environment, acoustic fatigue problems, and panelflutter experiences are discussed. Where these problems led to structural modifications, the modifications are described. 338. Hoey, Robert G.; and Day, Richard E.: Mission Planning and Operational Procedures for the X-15 Airplane. NASA TN D-1159, March 1962, 62N10585, 87H27608, #. Mission-planning methods and techniques used for the X-15 airplane envelope-expansion flight-test program are discussed. Use of the six-degree-of-freedom, ground-based simulator is indicated for prediction of performance, stability and controllability; development of piloting techniques and pilot training; evaluation of, and practice for, all possible emergency conditions; and energy management development. Other pilot-training devices and the role of the ground-monitoring station are also described. Predicted trajectory data and actual flight results are compared. The initial reasons and the final justifications for conducting the X-15 envelope expansion by performance increment are presented. 65 339. McLeod, Norman J.: Flight-Determined Aerodynamic-Noise Environment of an Airplane Nose Cone Up to a Mach Number of 2. NASA TN D-1160, March 1962, 62N10644, 87H27611, #. The aerodynamic-noise environment was measured at one point on the surface of a 24.5 degrees included-angle cone and at three internal positions. The data were obtained in flight for Mach numbers from 0.8 to 2, free-stream dynamic pressures from approximately 200 lb sq ft to 1,000 lb sq ft, and at altitudes of about 26,000 feet and 40,000 feet. The over-all noise levels and spectrum analysis of representative selected data are presented. Matranga, Gene J.; Dana, William H.; and Armstrong, Neil A.: Flight Simulated Off-the-Pad Escape and Landing Maneuvers for a Vertically Launched Hypersonic Glider. NASA TM X-637, March 1962, 66N33330, 87H26328, #. 340. A series of subsonic maneuvers was flown with an airplane having a maximum lift-drag ratio of 4.7. No particularly difficult piloting or maneuvering problems were encountered. A reduction of the pilot’s visibility from the cockpit did not appreciably impair his navigation capabilities, but did adversely affect his performance of the escape and landing maneuvers. 341. Banner, Richard D.; Kuhl, Albert E.; and Quinn, Robert D.: Preliminary Results of Aerodynamic Heating Studies on the X-15 Airplane. NASA TM X-638, March 1962, 66N29468, 87H26332, #. (See also 313.) The results of the preliminary flight heat-transfer studies on the X-15 airplane are presented, together with a discussion of the manner in which the data have been obtained, a comparison of measured and calculated turbulent heattransfer coefficients, a correlation of the model test results and the flight results for turbulent heat transfer, some information on boundary-layer transition, and a comparison of measured and calculated skin temperatures at several locations on the airplane. 342. McKay, James M.; and Kordes, Eldon E.: Landing Loads and Dynamics of the X-15 Airplane. NASA TM X-639, March 1962, 63N12564, 87H26336, #. (See also 313.) The loads, accelerations, and displacements of the X-15 airplane and landing-gear system measured during landing impact are discussed. The measured quantities are related to the initial touchdown conditions and are compared with data from a theoretical analysis to determine the effects of variations in such parameters as elevator position, skid coefficient of friction, main-gear location, and initial touchdown conditions beyond the range of the experimental data. 343. Keener, Earl R.; and Pembo, Chris: Aerodynamic Forces on Components of the X-15 Airplane. NASA TM X-712, March 1962, 65N23920, 87H25526, #. (See also 317.) Aerodynamic force data on the components of the X-15 airplane have been obtained by both pressure and strain-gage measurements in flights covering a Mach number range up to 6.04, altitudes up to about 217,000 feet, and angles of attack up to 15 degrees. Comparison of the flight data with windtunnel data shows generally good agreement for the flight conditions covered. 344. Hopkins, Edward J.; Fetterman, David E. Jr.; and Saltzman, Edwin J.: Comparison of Full-Scale Lift and Drag Characteristics of the X-15 Airplane With Wind-Tunnel Results and Theory. NASA TM X-713, March 1962, 65N23921, 87H25527. (See also 318.) Comparisons are made between the minimum drag characteristics of the full-scale X-15 airplane and windtunnel model data and theory extrapolated to flight Reynolds numbers for Mach numbers of 2.5 and 3.0. Similar comparisons are made for drag due to lift and maximum liftdrag ratio for Mach numbers up to about 5. Speed-brake drag and base-drag results are presented up to Mach numbers of 5.5 and 6, respectively. 345. Walker, Harold J.; and Wolowicz, Chester H.: Stability and Control Derivative Characteristics of the X-15 Airplane. NASA TM X-714, March 1962, 65N23922, 87H25529, #. (See also 319.) The flight-determined derivative characteristics are compared with the predictions from wind-tunnel tests and theory for Mach numbers extending to 5.5 and angles of attack up to 17 degrees. With few exceptions, the predictions were found generally to be in good agreement with the flight data. Areas of deficient stability and control are briefly discussed. 346. *White, Robert M.; Robinson, Glenn H.; and Matranga, Gene J.: Resume of Handling Qualities of the X-15 Airplane. NASA TM X-715, March 1962, 65N23923, 87H25530, #. (See also 320.) A summary of handling qualities is presented as assessed from pilot opinion and flight data. Segments of the flight profile which were evaluated include the launch, climbout, semiballistic flight, atmosphere entry, and landing. Longitudinal controllability is compared with results from current studies of reentry-type vehicles. *Air Force Flight Test Center, Edwards AFB, California, 347. Petersen, Forrest S.; Rediess, Herman A.; and Weil, Joseph: Lateral-Directional Control of the X-15 Airplane. 66 NASA TM X-726, March 1962, 65N23924, 87H25533, #. (See also 321.) The deterioration of lateral-directional controllability with roll damper off and the pilot performing a lateral-control task is discussed. The problem area was defined by fixed-base and airborne simulators and verified by closed-loop analysis in which a human transfer function represents the pilot. A parameter which predicts the problem area for the X-15 airplane is developed. The means considered to alleviate the control problem in the X-15 airplane are also discussed. 348. Taylor, Lawrence W., Jr.; Samuels, James L.; and Smith, John W.: Simulator Investigation of the Control Requirements of a Typical Hypersonic Glider. NASA TM X-635, H-226, March 1962, 72N71506, 87H26322. The handling qualities of a typical hypersonic glider were investigated with a flight simulator at Mach numbers of 0.26, 1.0, 3.5, 8, and 20 over an angle-of-attack range of 0° to 50°. Inasmuch as flight conditions influencing the control of the glider can be expected to change relatively slowly, a fivedegree-of-freedom mechanization was used. Pilots assessed the controllability of the glider without augmentation with fixed gain dampers, and with on adaptive control system. The investigation was limited to aerodynamic control. 349. Kordes, Eldon E.; Reed, Robert D.; and Dawdy, Alpha L.: Structural Heating Experiences on the X-15 Airplanes. NASA TM X-711. Prepared in cooperation with North American Aviation, Inc., Inglewood, California, March 1962, 71N75350, 87H25525. A survey of maximum structural temperatures measured on the X-15 airplane during speed flights up to a Mach number of 6 is presented. Structural problems caused by local hot spots and discontinuities are described. Structural modifications in the affected areas to eliminate the problems are discussed. 350. Walker, J. A.: The X-15 Program. Presented at The Institute of Aerospace Sciences meeting, St. Louis, Missouri, April 30–May 2, 1962, 62N12923, #. The high-temperature structural design approach utilized for the X-15 configuration has been successful; no major design deficiencies were encountered nor major modifications required. With but few exceptions, the local thermal problems encountered have not affected primary structural areas. In general, the aerodynamic derivatives extracted from flight-test data have confirmed the estimated derivatives obtained from wind-tunnel evaluations at hypersonic speeds. The aerodynamic flight control system and the simple stability augmentation system of the X-15 airplane have proved to be good technical designs. The airplane can be flown with satisfactory handling qualities through the range of dynamic pressures from about 1,500 lb/sq ft to below 100 lb/sq ft through the range of Mach numbers from 6.0 to subsonic landing conditions. Although only limited flight experience has been gained with the reaction control system, its basic design appears to be completely adequate. This type of system apparently provides an adequate means of attitude control for future space vehicles. Pilot transition from aerodynamic controls to reaction controls has been accomplished without problems. 351. Drake, H. M.: Crew Safety and Survival Aspects of the Lunar-Landing Mission. Presented at IAS Meeting on Man’s Progress in the Conquest of Space, St. Louis, Missouri, April 30–May 2, 1962, 62N12866, #. Some of the safety and survival aspects of the manned lunarlanding mission are examined. The conditions requiring abort to the earth, lunar orbit, and lunar surface are determined. Some of the possible design requirements to permit abort to lunar orbit or surface are indicated. Lunar orbital and surface survival kits are described, and the stationing of such kits in lunar orbit and at the intended landing site is proposed. 352. Weil, Joseph: Review of the X-15 Program. NASA TN D-1278, June 1962, 62N13289, 87H27936, #. The X-15 project is reviewed from its inception in 1954 through 1961. Some of the more important historical aspects of the program are noted, but major emphasis is placed on the significant research results. 353. Fichter, W. B.; and Kordes, E. E.: Response of Multiweb Beams to Static and Dynamic Loading. NASA TN D-1258, May 1962, 62N11650, #. 354. Drake, H. M.: Survey of FRC Recovery Research Presented to Meeting on Space Vehicle Landing and Recovery Research and Technology, (see N73-70937 04-99), July 11, 1962, 73N70943. 355. Horton, V. W.: Manned Paraglider Flight Tests. Presented to Meeting on Space Vehicle Landing and Recovery Research and Technology, (see N73-70937 04-99), July 11, 1962, 73N70944. 356. Armstrong, Neil A.; and Holleman, Euclid C.: A Review of In-Flight Simulation Pertinent to Piloted Space Vehicles. AGARD Report 403, 21st Flight Mechanics Panel Meeting, Paris, France, July 9–11, 1962. This paper shows how the environment of actual flight may be used to simulate many phase of manned space exploration. A number of simulations using conventional, modified, and specially built aircraft are discussed in relation to the portion of space flight to which they are generally applicable, that is, the launch, orbital, entry, or the landing approach phase. 357. Holleman, Euclid C.; and Armstrong, Neil A.: Pilot Utilization During Boost. Presented at the Inter-Center 67 Technical Conference on Control Guidance and Navigation Research for Manned Lunar Missions, Ames Research Center, Moffett Field, California, July 24–25, 1962, 63X14567. The capabilities of the pilot as the controller of an aircraft have been well documented, but relatively few investigators have considered the use of the pilot as the primary controller of a vertical launch vehicle. This role in the launch of a multistage vehicle has received increased interest recently. Figure l summarizes the studies in this area made to date. Of necessity, these programs utilized simulators to represent the launch vehicle. Some investigators used a human centrifuge to simulate the acceleration representative of these vehicles. 358. Taylor, Lawrence W., Jr.; and Adkins, Elmor J.: Recent X-15 Flight Test Experience With the MH-96 Adaptive Control System. Presented at the Inter-Center Technical Conference on Control Guidance and Navigation Research for Manned Lunar Missions, Ames Research Center, Moffett Field, California, July 24–25, 1962. 359. Matranga, Gene J.; and Bellman, Donald R.: Concept of a Free-Flying Lunar Landing and Take-Off Research Vehicle. Presented at the Inter-Center Technical Conference on Control Guidance and Navigation Research for Manned Lunar Missions, Ames Research Center, Moffett Field, California, July 24–25, 1962. 360. Layton, Garrison P.; and Thompson, Milton O.: Summary of Low-Speed Paraglider Flight Investigations. Part 1, Performance and Control Characteristics, Part 2, Flare and Landing. Presented at the Inter-Center Technical Conference on Control Guidance and Navigation Research for Manned Lunar Missions, Ames Research Center, July 24–25, 1962. 361. Kordes, Eldon E.: Experience With the X-15 Airplane in Relation to Problems of Reentry Vehicles. 3rd Congress of the International Council of the Aeronautical Sciences, Stockholm, Sweden, August 27–September 1, 1962. 62N12630. This paper discusses some of the results obtained from the flight program of the X-15 research airplane that have application to the design philosophy of future glide reentry vehicles. Experiences in the areas of panel flutter, landing dynamics, flight control systems, and aerodynamic and structural heating are described; and some of the problems that have developed are discussed briefly. A bibliography of papers published on the X-15 flight program is included. 362. Walker, J. A.: I Fly the X-15. National Geographic, Vol. 122, No. 3, September 1962, pp. 428–450. 363. Armstrong, N. A.; Walker, J. A.; Petersen, F. S.; and White, R. M.: The X-15 Flight Program. NASA Proceedings of the Second National Conference on the Peaceful Uses of Space, Seattle, Washington, May 8–10, 1962, November 1962, pp. 263–271, 63N11158. This paper reviews the philosophy of the X-15 airplane, describes its concept in operation, and prophesies its future. The X-15 was the first design to require rocket reaction control within its design envelope. The flight and landing of the X-15 are reviewed. Some of the follow-on uses of the X-15 include guidance-instrument experiments. precise determination of atmospheric density at extreme altitude, measurement of size and quantity of micrometeorites in near space, and determination of the intensity of ultraviolet and infrared rays in near space. 364. Veatch, Donald W.: X-20 Instrumentation Sensors. Paper for the X-20A (Dyna-Soar) Symposium, November 5–7, 1962. 365. Martin, J. A.: The Record-Setting Research Airplanes. Reproduced from Aerospace Engineering, Vol. 21, No. 12, December 1962, pp. 49–54, 63N13571. The first compilation of all available data on the unofficial records of the rocket airplanes is presented. The maximum Mach number, true velocity, and the altitude obtained by the X-1-1, D-558-II, X-1A, X-2, and X-15 airplanes are given in tabular form. Also, the physical characteristics of these airplanes are given and include wing spans, wing sweep, launch weight, and landing weight. 366. Tremant, R. A.: Operational Experiences and Characteristics of the X-15 Flight Control System. NASA TN D-1402, December 1962, 63N11123, #. X-15 flight and simulator experiences with the manual flight control and stability augmentation system for the period from December 1958 to January 1962 are presented. The flight data extend to a Mach number of 6.04 and an altitude of 217,000 feet, and the simulator data cover the design flight envelope. The characteristics of the manual flight control system and the stability augmentation system are discussed, in conjunction with pilot evaluation, operational problems, modifications, and reliability. Pertinent X-15 flight history is included. Five test vehicles were air-launched from an F-104A airplane to investigate some of the operational aspects and the practicability of using the energy input of the airplane as a first-stage booster for sounding rockets. A launch maneuver and launcher system were developed and matched to the airplane’s capabilities so that suitable repeatability of launch parameters was attained. 368. Larson, T. J.; and Webb, L. D.: Calibrations and Comparisons of Pressure-Airspeed-Altitude Systems of the X-15 Airplane From Subsonic to High Supersonic Speeds. NASA TN D-1724, February 1963, 63N12951, #. The X-15 flight calibration data to define static-pressure position errors are presented for two types of pressuresensing configurations: a standard NACA pitot-static tube attached to a nose boom, and two manifolded flush staticpressure ports on the ogive nose. The position-error calibrations are presented up to M = 3.31 for the standard nose boom installation and up to M = 4 for the flush static system. Presented also are stagnation-pressure errors sensed by a pitot probe ahead of the canopy. Methods used to determine the position errors are described. The nose-boom configuration is shown to be superior from the standpoint of position error and ease of calibration for the available data range. 369. Holleman, Euclid C.; and Wilson, Warren, S.: Flight-Simulator Requirements for High-Performance Aircraft Based on X-15 Experience. ASME Paper 63-AHGT-81, ASME Aviation and Space, Hydraulics, and Gas Turbine Conference and Products Show, Los Angeles, California, March 3–7, 1963, 63A17579. Review of the simulation experience acquired during the design and flight testing of the X-15 research airplane. Discussed are the problems encountered and the use of simulators in their solution, with particular reference to the X-15 fixed-base simulator. Simulator techniques which may be used in the supersonic-transport program, such as a variable-stability airborne simulator, are suggested. 370. Row, Perry V.; and Fischel, Jack: Operational Flight-Test Experience With the X-15 Airplane. AIAA Paper 63-075, AIAA Space Flight Testing Conference, Cocoa Beach, Florida, March 18–20, 1963, 63A15995. Review of the experience of the NASA Flight Research Center in coping with the problems of component and system checkout and operational flight procedures in an advanced flight research program. The operational evolution of the most troublesome and the most important systems on the North American X-15 is discussed, and operational and research data are presented. Procedures now being utilized and those that will be applied to newer systems are also discussed. The use of a flight simulator to check out several 68 1963 Technical Publications 367. Horton, V. W.; and Messing, W. E.: Some Operational Aspects of Using a High-Performance Airplane as a First-Stage Booster for Air-Launching Solid-Fuel Sounding Rockets. NASA TN D-1279, January 1963, 63N12192, #. flight systems and to practice normal and emergency operational procedures is assessed. Finally, the flight operational techniques evolved for ground monitoring of flight systems data and flight-trajectory information to provide pilot backup support are described. Evidence is introduced to show how these techniques have facilitated the rapid expansion of the flight envelope and have aided in achievement of research objectives. 371. Layton, G. P., Jr.; and Thompson, M. O.: Preliminary Flight Evaluation of Two Unpowered Manned Paragliders. NASA TN D-1826, April 1963, 63N14429, #. Towed and free-flight tests were made with unpowered, manned paragliders to study the performance, stability, and control characteristics of a typical paraglider. The paragliders used had maximum lift-drag ratios greater than 3.5 and wing loadings of approximately 4.0 lb/sq ft. The airspeed range was limited by the rearward center-of-pressure shift at angles of attack above and below trim angle of attack. Performance data obtained from flight tests are presented and compared with analytical results. Center-of-gravity shift, accomplished by tilting the wing relative to the fuselage, was used for control. This method of control was adequate for towed and free flight as well as for flare and landing. The pilot’s evaluation of the vehicle’s handling qualities, and a discussion of development problems are presented. 372. Walker, J. A.; and Weil, J.: The X-15 Program. Proceedings of AIAA 2nd Manned Space Flight Meeting, AD-400711, April 22–24, 1963, pp. 295–307, 63A19019, 63N23237. Review of the important operational problems encountered in the flights of the North American X-15 aircraft. The history of the project is considered, outlining the design development and the flight tests, including aerodynamic configurations, mode of operation, flight program, and performance. Summarized is the operational experience, such as the structural and thermostructural problems, and the rocket engines, auxiliary-power-unit, and the control and guidance systems experience. The piloting aspects of the X-15 mission are described, including the boost, entry, and landing techniques. 373. Noll, R. B.; and Halasey, R. L.: Theoretical Investigation of the Slideout Dynamics of a Vehicle Equipped With a Tricycle Skid-Landing-Gear System. NASA TN D-1828, May 1963, 63N16298, #. The equations-of-motion for the slideout of a vehicle equipped with a tricycle skid-type landing-gear system are presented and reduced to three degrees of freedom. A comparison of the results of numerical calculations for the three-degree-of-freedom slideout of the X-15 research 69 vehicle with flight-test results shows that the theoretical analysis of the slideout can adequately predict the slideout distance, the direction of lateral displacement, and the approximate lateral displacement. A numerical study of the slideout equations indicates that the velocity at which the aerodynamic influence on the vehicle becomes negligible can be predicted. 374. Ferguson, T. J.: Flight Research Center Instrumentation Program. NASA Goddard Space Flight Center Proceedings of the Optical Communications and Tracking Symposium, (see N68-84351), June 1963, 68N84355. 375. Row, Perry V.; and Fischel, Jack: X-15 Flight Experience. Astronautics and Aerospace Engineering, Vol. 1, June 1963, pp. 25–32, 63A17556. Survey of the operational aspects of the X-15 aircraft development program. Discussed are the major structural problems encountered, such as landing-gear overloading, panel flutter, side-fairing buckling, wing leading-edge skin buckling, windshield heat damage, and internal heat damage. It is shown how the X-15 was made operational for the simplest of tasks and was then built up to the maximum demands in discrete progressive steps. Accidents in the program, and their causes, are examined. 376. Matranga, G. J.; Washington, H. P.; Chenoweth, P. L.; and Young, W. R.: Handling Qualities and Trajectory Requirements for Terminal Lunar Landing, as Determined From Analog Simulation. NASA TN D-1921, August 1963, 63N19606, #. A six-degree-of-freedom analog study was performed to aid in defining handling qualities and trajectory potential for terminal lunar landing. Results showed that, for a maneuvering task in the pitch mode and a random-motioncorrection task in the roll and yaw modes, the pilots preferred rate or attitude command with control accelerations of about 10 deg/sec and reasonable artificial damping. Also, to consistently perform successful landings, the pilots generally used thrust-to-weight ratios throttled between a minimum value of 0.8 lunar g and maximum value of 1.8 lunar g. 377. Watts, J. D.; and Banas, R. P.: X-15 Structural Temperature Measurements and Calculations for Flights to Maximum Mach Numbers of Approximately 4, 5, and 6. NASA TM X-883, H-315, August 1963, 72N73396. Structural temperatures on the X-l5 airplane were measured continuously during three performance-envelope expansion flights to maximum Mach numbers of approximately 4, 5, and 6. Tabulations of temperature time histories, representing all surfaces and some wing internal structure, are presented for these flights. Methods of predicting surface temperature levels and gradients are described, and the resulting calculations are compared with measured temperatures. landings are described. Data from typical X-15 landings and from landings in which modified touchdown techniques were used are presented and compared. 381. Weil, Joseph: Piloted Flight Simulation at the NASA Flight Research Center. Presented at IEEE 10th Annual East Coast Conference on Aerospace and Navigation Electronics, Baltimore, Maryland, October 21–23, 1963. 382. Nugent, Jack: The X-15 Advanced Air-Breathing Engine Program. Presented at Bumblebee Composite Design Research Panel, November 1963. 383. Horton, V. W.; Layton, G. P., Jr.; and Thompson, M. O.: Exploratory Flight Tests of Advanced Piloted Spacecraft Concepts. NASA TM X-51360. Presented at the AIAA, AFFTC, and NASA FRC Testing of Manned Flight Systems Conference, Edwards, California, December 4–6, 1963, 65N89037. 384. Sisk, T. R.; and Andrews, W. H.: Utilization of Existing Aircraft in Support of Supersonic-Transport Research Programs. NASA TM X-51360. Presented at the AIAA, AFFTC, and NASA FRC Testing of Manned Flight Systems Conference, Edwards, California, December 4–6, 1963, pp. 67–76, 64N12881. The supersonic transport will not necessarily be derivative of a previous military airplane, as are the current family of jet transports. Therefore, full-scale test data and operational experience for this vehicle will be limited, and in some areas nonexistent. In an attempt to fill this void, the NASA Flight Research Center has initiated several programs utilizing existing high-performance aircraft to investigate some of the problems predicted in the supersonic-transport operational environment. This paper discusses three of these programs: a minimum-flight-speed investigation utilizing an F5D aircraft, and Air Traffic Control (ATC) compatibility program utilizing an A-5A aircraft, and specific vehicle research on the XB-70 aircraft. 385. Rediess, H. A.; and Deets, D. A.: An Advanced Method for Airborne Simulation. NASA TM X-51360. Presented at the AIAA, AFFTC, and NASA FRC Testing of Manned Flight Systems Conference, Edwards, California, December 4–6, 1963, pp. 33–39, 64N12880. (See also 406.) The NASA Flight Research Center has conducted and sponsored studies leading to the design and developments of a general-purpose airborne simulator (GPAS) to support the supersonic-transport program and to perform general research. This paper presents some of the results of these studies. The response feedback, and the model-control concepts for an airborne simulator are discussed and evaluated. The model-following performance of the system designed for the GPAS, and, other results believed generally applicable are also presented. 70 EC88-0180-1 X-15 Airplane 378. Videan, Edward N.; Banner, Richard D.; and Smith, John P.: The Application of Analog and Digital Computer Techniques in the X-15 Flight Research Program. Presented at the International Symposium on Analog and Digital Techniques Applied to Aeronautics, Liege, Belgium, September 9–12, 1963. This paper is limited, however, to the particular area of computer application to research flight planning and system implementation. Two systems are described, one analog and one digital, which support the flight planning and system implementation. Today, because of changing flight test requirements, a relatively sophisticated and complete simulation system is considered necessary to carry on a flight research program. 379. Thompson, Milton O.: Preliminary Results of the Lifting-Body Flight Program. NASA TM X-56005. Presented at the 7th Annual SETP Symposium, Beverly Hills, California, September 27–28, 1963. This paper covers the design, construction, preflight, and initial flight testing of the NASA Flight Research Center’s M-2 lifting-body vehicle. The paper also discusses the concept of lifting-body utilization and the reasons for construction of a lightweight vehicle. 380. Noll, R. B.; Jarvis, C. R.; Pembo, C.; Lock, W. P.; and Scott, B. J.: Aerodynamic and Control-System Contributions to the X-15 Airplane Landing-Gear Loads. NASA TN D-2090, October 1963, 63N22117, #. The effects of the X-15 manual flight control and stability augmentation systems on the horizontal-tail load, and the effect of wing-flap position on the wing load during touchdown are investigated. Methods for significantly reducing the maximum total load on the main gear during 386. Larson, T. J.; and Montoya, E. J.: Stratosphere and Mesosphere Density-Height Profiles Obtained With the X-15 Airplane. NASA TM X-51734, 1963, 65N33708, #. Density-height profiles in the stratosphere and mesosphere were obtained from impact-pressure, velocity,. and altitude measurements made on six X-l5 research airplane flights. A form of the Rayleigh pilot formula was used for density computations. Because of pressure-instrumentation limitations and pressure lag, the maximum altitude for reasonably accurate density determination was considered to be about 65 km. Good agreement was obtained between temperatures calculated from faired density-height profiles of two X-l5 flights and temperatures measured by rocketsondes launched near the times of flight from the Pacific Missile Range, Point Mugu, California. ranging for landing; however, the pilots indicated that fasteracting speed brakes would allow more flexible operation. 389. Holleman, E. C.; and Adkins, E. J.: Contributions of the X-15 Program to Lifting Entry Technology. AIAA Paper 64-17, January 1964, 64N15273. (See also 388.) 390. Drake, H. M.: Aerodynamic Testing Using Special Aircraft. NASA TM X-51605. Presented at AIAA Aerodynamic Testing Conference, Washington, D. C., March 9–10, 1964, 65N88557, 65N35225, #. (See also 391.) In this paper some of the recent applications of special aircraft to aerodynamic testing are reviewed and something of the complementary relationship such testing bears to theory and to research in ground facilities is indicated. Some of the primary reasons for flight research are to verify theory, ground facilities, and design. Encounter new or overlooked problems. Investigate flight in the true environment. Establish crew-vehicle integration and requirements. Study the atmosphere, earth, and space. 391. Drake, H. M.: Aerodynamic Testing Using Special Aircraft. American Institute of Aeronautics and Astronautics Aerodynamics Testing Conference, March 10, 1964, pp. 78– 188, 64N17019. (See also 390.) 392. Quinn, R. D.; and Kuhl, A. E.: Comparison of Flight-Measured and Calculated Turbulent Heat Transfer on the X-15 Airplane at Mach Numbers From 2.5 to 6.0 at Low Angles of Attack. NASA TM X-939, H-332, March 1964, 72N73498. Turbulent heat-transfer data obtained on the X-l5 airplane for a flight to a Mach number of 6.0 are presented and compared with calculated values. Calculated boundary-layer thicknesses and Mach number profiles in the shear layer are also presented. Comparisons between measured and calculated heat-transfer coefficients show that the calculated heat-transfer coefficients are from 30 to 60 percent higher than the measured values when bluntness effects are included in estimates of the local Mach number at the edge of the boundary layer. 393. Fischel, Jack; and Toll, Thomas A.: The X-15 Project: Results and New Research. Astronautics & Aeronautics, March 1964, pp. 20–28. 394. Chenoweth, P. L.; and Dana, W. H.: Flight Evaluation of Wide-Angle Overlapping Monoculars for Providing Pilot’s Field of Vision. NASA TN D-2265, April 1964, 64N17753, #. A qualitative evaluation was made of the effectiveness of wide-angle, overlapping monoculars as the sole source of outside visual reference during takeoffs, aerial maneuvers, 71 1964 Technical Publications 387. Pyle, J. S.: Comparison of Flight Pressure Measurements With Wind-Tunnel Data and Theory for the Forward Fuselage of the X-15 Airplane at Mach Numbers From 0.8 to 6.0. NASA TN D-2241, January 1964, 64N12961, #. The results of flight pressure measurements on the forward fuselage of the X-15 airplane are presented for angles of attack from O degrees to 15 degrees and Mach numbers from 0.8 to 6.0. Comparisons of flight and wind-tunnel data showed good agreement, and theoretical calculations predicted flight pressure measurements reasonably well. 388. Holleman, E. C.; and Adkins, E. J.: Contributions of the X-15 Program to Lifting Entry Technology. NASA TM X-51359. Presented at the AIAA Aerospace Sciences Meeting, New York, January 20–24, 1964, 65N89079. (See also 389.) Entries from altitudes greater than 350,000 ft with the X-15 airplane have provided piloting experience and verification of predicted control characteristics and operational techniques. The airplane re-enters as a glider and duplicates several phases in the recovery of higher-performance vehicles, for example, transition from near-zero dynamic pressure to aerodynamic flight, and the terminal-area ranging and landing. During entries, reaction controls have been used to surprisingly high dynamic pressures. Rate command control provided satisfactory control, and hold modes were appreciated by the pilots for secondary control modes. With conservatively planned flights, the pilots have had no problem controlling range to base with contact navigation. Landmarks have been observed from above 300,000 ft and 160 miles range. The approach and landing of the low-liftdrag-ratio X-15 airplane has become routine, with relatively small dispersion in touchdown and slideout distance. The speed brakes have been an important control for regulation of visual navigation, and approaches and landings in a light observation aircraft. The evaluation was made during the day and at night and in air conditions which varied from no turbulence to severe turbulence. 395. Holleman, E. C.: Piloting Performance During the Boost of the X-15 Airplane to High Altitude. NASA TN D-2289, April 1964, 64N19002, #. 396. Gray, W. E., Jr.: NASA Flight Research Center Handling-Qualities Program on General-Aviation Aircraft. NASA TM X-56004, April 21, 1964, 65N35235, #. 397. Jarvis, C. R.; and Adkins, E. J.: Operational Experience With X-15 Reaction Controls. NASA TM X-56002. Presented at the SAE-ASME Symposium on Position, Attitude and Thrust Vector Control, April 21, 1964, 64N20683, #. The four reaction-control-system configurations investigated during the X-15 program include a proportional acceleration command system, on-off proportional rate command and attitude hold systems, and a rate-sensing on-off stability augmentation system. Each of the systems is described briefly, and development problems encountered in hardware design, component compatibility, and systems integration are discussed. The practical aspects of system design and operation are emphasized. Flight experience with each system is also discussed. Flight data showing the results of open-loop and closed-loop control during critical X-15 reentry maneuvers are presented. 398. Sanderson, K. C.: The X-15 Flight Test Instrumentation. NASA TM X-56000. Flight Test Instrumentation, Vol. 3, pp. 267–290, proceedings of the Third International Symposium, (see TL 671.7 I54 V. 3), April 21, 1964, 1964, 64N19899, #. The basic instrumentation philosophy for the X-15 program was dictated primarily by two factors. First, if the X-15 were to successfully fulfill its mission of providing timely research data, it had to be built and instrumented quickly. Second, the instrumentation had to be accurate and reliable. The philosophy adopted was as follows: Onboard recording would be used, selected parameters would be telemetered and displayed to ground monitors in real time, continuous ground radar tracking provided instrumentation system would have to be flexible, maximum use of off-the-shelf instrumentation components and systems. 399. Wall, D. E.: A Study of Hypersonic Aircraft. NASA TM X-56001, April 24, 1964, 64N20549, #. (See also 400.) A study is being made at the NASA Flight Research Center to determine the gross characteristics of future hypersonic aircraft, without the refinement of configuration 72 optimization. The characteristics defined by this study are to be used as a guide in assessing the need for future hypersonic flight research. 400. Wall, D. E.: A Study of Hypersonic Aircraft. NASA TM X-51641, 1964, 65N35263, #. (See also 399.) 401. Taylor, L. W., Jr.; and Adkins, E. J.: Adaptive Flight Control Systems—Pro and Con. NASA TM X-56008. Presented at the AIAA Specialists Meeting, Los Angeles, California, April 28, 1964, 64N27261, #. In light of difficulties posed by the X-15, the adaptive flight control system was developed and has been most successful. Although several problems were encountered during the development of the MH-96 adaptive system, and emphasis on them in this paper tends to paint a dark picture, these problems were solved on the ground before the first flight, except for some insignificant details which affected only the periphery functions of the MH-96 even during the early flights. There is a saying that “a bird in hand is worth two in the bush.” For adaptive flight control system concepts, we would put the ratio at about 10. An adaptive control system which has been successfully demonstrated in the X-15 is worth about 10 proposed new adaptive concepts which have not been exposed to the idiosyncrasies of control-system hardware. 402. Kordes, E. E.; and Tanner, C. S.: Preliminary Results of Boundary-Layer Noise Measured on the X-15 Airplane. NASA TM X-56003, May 1, 1964, 65N35284, #. In order to provide detailed information on boundary-layer noise over a wide range of controlled flight conditions, the NASA Flight Research Center is conducting a boundarylayer-noise research program with the X-l5 airplane. This paper describes the program and presents some of the preliminary results. 403. Thompson, M. O.: Aerospace Medical and Bioengineering Considerations in Lifting-Body and Research-Aircraft Operations. NASA TM X-56005. Presented at the 35th Aerospace Medical Association Annual Meeting, Miami Beach, Florida, May 11–14, 1964, 64N22440, #. The lifting-body vehicle we have flown at the Flight Research Center is the M-2 rather than the M-l; thus, it is this vehicle I shall discuss. For those who may not be familiar with the M-2 or the lifting-body concept, I shall describe it briefly. As the name implies, a lifting body is a vehicle with a body shape, rather than wings, which generates lift at an angle or attack. The only irregularities or protuberances in the body shape are the surfaces required for aerodynamic control. Figure l compares the advantages or the three configurations having reentry capability, that is, the ballistic or semiballistic, the lifting body, and the winged vehicle. The energy footprints or the vehicles, or landing areas available to each, can be estimated. For operational usage, a lifting reentry vehicle appears to be highly desirable because of its versatility for reentry from a number of orbit planes or the capability for recovery at a number of landing sites within the United States. 404. Maher, J. F., Jr.; Ottinger, C. W.; and Capasso, V. N., Jr.: YLR99-RM-1 Rocket Engine Operating Experience in the X-15 Aircraft. NASA TN D-2391, July 1964, 64N25810, #. This paper describes the unique operating experience obtained during the first 50 government flights with the YLR99 engine installed in the X-15 aircraft, with emphasis on problem areas of the engine and their effects on the X-15 program. 405. Powers, B. G.; and Matheny, N. W.: Flight Evaluation of Three Techniques of Demonstrating the Minimum Flying Speed of a Delta-Wing Airplane. NASA TN D-2337, July 1964, 64N24966, #. A flight test program was conducted with an F5D airplane to evaluate three techniques for demonstrating the minimum flying speed of a delta-wing aircraft: The Civil Air Regulations stall-speed demonstration, the 1 g demonstration, and the constant-rate-of-climb demonstration. The Civil Air Regulation stall-speed demonstration currently used for civil transport aircraft was found to be inadequate for demonstrating the minimum speed of a delta-wing airplane, because this type of airplane does not have a well-defined stall point near maximum lift coefficient. The 1 g minimum speed, which is based on maintaining a constant 1 g normal acceleration, was difficult to determine precisely, especially when buffeting was present. The constant-rate-of-climb minimum-speed maneuvers, which are based on the ability to maintain a constant rate of climb, were reasonably easy to perform and were unaffected by the aircraft buffet characteristics. The level-flight minimum speed obtained from the constant-rate-of-climb techniques was found to be the most rational minimum speed for a delta-wing aircraft. The applicability of these techniques to other types of aircraft was shown in limited tests on a sweptwing airplane. 406. Rediess, H. A.; and Deets, D. A.: An Advanced Method for Airborne Simulation. NASA RP 337. Reprinted from J. Aircraft, Vol. 1, No. 4, July–August 1964, pp. 185–190. Presented at the AIAA, AFFTC, and NASA FRC Testing of Manned Flight Systems Conference, Edwards AFB, California, December 4–6, 1963, 64N31214. (See also 385.) In a general discussion of airborne simulation, it is observed that the motion of a specific aircraft cannot be matched completely with an airborne simulator, except at certain specific conditions, if the number of independent control devices for angular and linear motion is less than the number 73 of corresponding degrees of freedom to be stimulated. However, airborne simulators can be valuable research and pilot-training tools through proper choice of the motion parameters to be matched and by tailoring the program to the particular simulator used. 407. Fischel, J.; and Webb, L. D.: Flight-Informational Sensors, Display, and Space Control of the X-15 Airplane for Atmospheric and Near-Space Flight Missions. NASA TN D-2407, August 1964, 64N26629, #. This paper presents pertinent information obtained during the X-15 program and discusses its use by the pilot in performing a variety of atmospheric and near-space flight missions. 408. Saltzman, Edwin J.: Base Pressure Coefficients Obtained From the X-15 Airplane for Mach Numbers Up to 6. NASA TN D-2420, August 1964, 64N27122, #. Base pressure measurements were made on the vertical fin, side fairing, fuselage, and wing trailing edge of the X-15. Data are presented between Mach numbers of 0.8 and 6. Power-off and power-on data are included and compared with wind-tunnel measurements and theory. 409. Hughes, D. L.; Powers, B. G.; and Dana, W. H.: Flight Evaluation of Some Effects of the Present Air Traffic Control System on Operation of a Simulated Supersonic Transport. NASA TN D-2219, November 1964, 64N33082, #. An exploratory flight program was conducted to investigate the effect of the present Air Traffic Control system on the operation of a simulated supersonic transport in the Los Angeles terminal area. The climb and descent portions of a representative supersonic transport flight profile were flown with an A-5A airplane. In addition, en route problems were explored within the speed and altitude data were obtained, as well as flight-crew opinions and ground-personnel comments. 410. Yancey, R. B.: Flight Measurements of Stability and Control Derivatives of the X-15 Research Airplane to a Mach Number of 6.02 and an Angle of Attack of 25 Degrees NASA TN D-2532, November 1964, 65N10638, #. Flight tests of the X-15 airplane provided data from which longitudinal, lateral, and directional stability and control derivations were determined over a Mach number range from 0.60 to 6.02 and over an angle-of-attack range from -2.7 degrees to 25 degrees. The data were obtained with the lower rudder on and off, speed brakes open and closed, and power on and off. The longitudinal derivatives show the expected trends of increasing levels through the transonic region and diminishing levels as the Mach number increases in the supersonic region. A high level of longitudinal stability is indicated by the flight data. 411. Holleman, E. C.; and Adkins, E. J.: Contributions of the X-15 Program to Lifting Entry Technology. J. Aircraft, Vol. 1, No. 6, November–December 1964. Entries from altitudes greater than 350,000 ft with the X-15 airplane have provided piloting experience and verification of predicted control characteristics and operational techniques. The airplane re-enters as a glider and duplicates several phases in the recovery of higher-performance vehicles, for example, transition from near-zero dynamic pressure to aerodynamic flight, and the terminal-area ranging and landing. During entries, reaction controls have been used to surprisingly high dynamic pressures. Rate command control provided satisfactory control, and hold modes were appreciated by the pilots for secondary control modes. With conservatively planned flights, the pilots have had no problem controlling range to base with contact navigation. Landmarks have been observed from above 300,000 ft and 160 miles range. The approach and landing of the low-liftdrag-ratio X-15 airplane has become routine, with relatively small dispersion in touchdown and slideout distance. The speed brakes have been an important control for regulation of ranging for landing; however, the pilots indicated that fasteracting speed brakes would allow more flexible operation. 412. Thompson, M. O.: General Review of Piloting Problems Encountered During Simulation and Flights of the X-15. NASA TM X-56884, Society of Experimental Test Pilots Ninth Annual Report. Presented at the SETP Symposium, Beverly Hills, California, 1964, 66N83857. 413. Fischel, J.; and Toll, T. A.: The X-15 Project— Results and New Research. NASA RP 186, 1964, 64N22066. 414. Fischel, J.; and Toll, T. A.: The X-15 Research Aircraft—Research Accomplished and Planned. NASA TM X-51485, 1964, 65N89070. 415. Winglade, R. L.: Current Research on Advanced Cockpit Display Systems. NASA TM X-56010, 1964, 65N20814, #. Current cockpit-display philosophy is discussed in terms of the pilot’s informational requirements. Pilots scan patterns obtained through the use of an eye-position camera and a ground-based simulator are depicted for a conventional display system and for two advanced concepts. Preliminary results of some flight-test and ground-simulation evaluations of advanced concepts, such as totally integrated displays and indirect pilot viewing systems, are discussed. 416. Montoya, Earl J.; and Larson, Terry J.: Stratosphere and Mesosphere Densities Measured With the X-15 Airplane. NASA RP 499, NASA TM X-56009, 1964. (See also Geophysical Research, Vol. 69, No. 4, pp. 5123–5130, 1964.) 74 Density-height profiles in the stratosphere and mesosphere were obtained from measurements of impact pressure, velocity, and altitude on six X-15 research airplane flights. A form of the Rayleigh pilot formula was used for density computations. Because of pressure-instrumentation limitations and pressure lag, the maximum altitude for reasonably accurate density determination was considered to be about 65 km. Temperatures calculated from faired densityheight profiles of two X-l5 flights agreed well with temperatures measured by rocketsondes launched near the times of flight from the Pacific Missile Range, Point Mugu, California. 1965 Technical Publications 417. Pyle, J. S.: Flight-Measured Wing Surface Pressures and Loads for the X-15 Airplane at Mach Numbers From 1.2 to 6.0. NASA TN D-2602, January 1965, 65N14854, #. 418. Roman, James: Long-Range Program to Develop Medical Monitoring In Flight—The Flight Research Program-I. Aerospace Medicine, Vol. 36, No. 6, June 1965. NASA’s Flight Research Center is conducting a long-range program designed to advance the state of the art in biomedical monitoring. Better knowledge of the physiological parameters used in monitoring the crew is one of major aims of the program. An instrumentation-development phase and a phase involving development of computer techniques for handling medical flight data both contribute to the overall program. The physiological-parameters-research phase and the instrumentation-development phase have yielded significant results after one year of operation. 419. Jarvis, C. R.; and Lock, W. P.: Operational Experience With the X-15 Reaction Control and Reaction Augmentation Systems. NASA TN D-2864, June 1965, 65N25725, #. This paper describes the X-15 reaction control system and discusses the system characteristics, operational experiences, and development problems. Data are presented from X-15 high-altitude flights during which both the manual control and reaction augmentation systems were operated. 420. Cary, J. P.; and Keener, E. R.: Flight Evaluation of the X-15 Ball-Nose Flow-Direction Sensor as an Air-Data System. NASA TN D-2923, July 1965, 65N27945, #. This paper assesses the suitability of the modified ball-nose system for obtaining Mach number and pressure altitude from pressure measurements at Mach numbers up to 5.3, altitudes up to 130,000 feet, and Reynolds number from 0.1 to 1.6 × 10 to the 6th per foot. The results are compared with experimental and theoretical results for spheres. 421. Pyle, J. S.: Flight Pressure Distributions on the Vertical Stabilizers and Speed Brakes of the X-15 Airplane at Mach Numbers From 1 to 6. NASA TN D-3048, September 1965, 65N34437, #. This paper, the third in series on the X-15 surface-pressure distributions, presents flight-measured pressure distributions for the upper and lower vertical stabilizers with the speed brakes opened and closed. Data are shown for Mach numbers from 1 to 6 and angles of attack from 0 degree to 15 degree. Comparisons are made with wind-tunnel data and theory. 422. Smith, H. J.: Evaluation of the LateralDirectional Stability and Control Characteristics of the Lightweight M2-F1 Lifting Body at Low Speeds. NASA TN D-3022, September 1965, 65N33839, #. This paper summarizes the lateral-directional stability and control characteristics investigated during the flight tests and compares some wind-tunnel data with the flight values. Performance data from the tests are reported. Research Vehicle. NASA TN D-3023, September 1965, 65N33549, #. This paper represents the significant technical details and research capabilities of a free-flight lunar-landing simulator as they existed at the time of the initial flights of the vehicle. The lunar-landing research vehicle (LLRV) consists of a pyramid-shaped structural frame with four truss-type legs. A pilot’s platform extends forward between two legs, and an electronics platform is similarly located, extending rearward. A jet engine is mounted vertically in a gimbal ring at the center of the vehicle. The LLRV is instrumented for research purposes. The data obtained are converted to digital form transmitted to a ground tape recorder by means of an 80-channel pulse-code-modulation type (PCM) telemetry system. Each channel can be read every 0.005 second, if desired. ECN-535 Lunar Landing Research Vehicle (LLRV) EC-64-404 M2-F1 Lifting Body Vehicle 423. Horton, V. W.; Eldredge, R. C.; and Klein, R. E.: Flight-Determined Low-Speed Lift and Drag Characteristics of the Lightweight M2-F1 Lifting Body. NASA TN D-3021, September 1965, 65N33357, #. The low-speed lift and drag characteristics of a manned, lightweight M-2 lifting-body vehicle were determined in unpowered free-flight tests at angles of attack from 0 degrees to 22 degrees (0.38 radian) and at calibrated airspeeds from 61 knots to 113 knots (31.38 to 58.13 meters/second). Flight data are compared with results from full-scale wind-tunnel tests of the same vehicle. 424. Bellman, D. R.; and Matranga, G. J.: Design and Operational Characteristics of a Lunar-Landing 75 425. Banas, R. P.: Comparison of Measured and Calculated Turbulent Heat Transfer in a Uniform and Nonuniform Flow Field on the X-15 Upper Vertical Fin at Mach Numbers of 4.2 and 5.3. NASA TM X-1136, H-382, September 1965, 72N73703. Turbulent heat-transfer coefficients and measured local static pressures were obtained in flight on the X-15 upper vertical fin with both a sharp and a blunt leading edge. The data are compared with calculated values. Calculated and measured Mach number profiles in the shear layer are also presented. Heat-transfer coefficients were obtained from measured skin temperatures at free-stream Mach numbers of approximately 4.2 and 5.3 and free-stream Reynolds numbers between 1.8 x l0(6) and 2.5 x l0(6) per foot. Comparisons of measured and calculated heat-transfer coefficients obtained in both a uniform flow field and a nonuniform flow field show that the heat-transfer coefficients calculated by Eckert’s referencetemperature method were from 32 percent to 57 percent higher than the measured values. 426. Anon.: Progress of the X-15 Research Airplane Program. NASA-SP-90, USAF, USN, and NASA Conference on Progress of the X-15 Research Airplane Program, Edwards AFB, California October 7, 1965, 73N71303. 427. Love, J. E.; and Fischel, J.: Status of X-15 Program. NASA SP-90, (see N73-71303 05-99), 1965, pp. 1–15, 73N71304. This paper briefly reviews the significant activities and present status of the project in order to aid you in properly relating the information presented during this conference to the total X-15 program. A comprehensive bibliography of information related to the X-15 program is included at the end of this paper. 428. Banner, R. D.; and Kuhl, A. E.: A Summary of X-15 Heat Transfer and Skin Friction Measurements. NASA SP-90, (see N73-71303 05-99), 1965, pp. 17–26, 73N71305. (See also 449.) 429. Lewis, T. L.; and McLeod, N. J.: Flight Measurements of Boundary Layer Noise on the X-15. NASA SP-90, (see N73-71303 05-99), 1965, pp. 27–33, 73N71306. (See also 451.) Boundary-layer-noise data measured in flight over a Mach number range from 1.0 to 5.4 and at altitudes from 45,000 feet to 105,000 feet are presented. The data were obtained at four locations on the X-l5 (selected to provide varied boundary-layer conditions). The highest recorded noise level was 150 decibels. Boundary-layer parameters were measured at one location and are used to present the noise data in a nondimensional form for comparison with data from flat-plate wind-tunnel studies by other experimenters. 430. McKay, J. M.; and Noll, R. B.: A Summary of the X-15 Landing Loads. NASA SP-90, (see N73-71303 05-99), 1965, pp. 35–43, 73N71307. (See also 447.) The purpose of this paper is to review the present status of the X-15 landing-gear loads, to discuss the parameters which affect these loads, and to show additional modifications that might be made to improve the landing gear system. 431. Taylor, L. W., Jr.; Robinson, G. H.; and Iliff, K. W.: A Review of Lateral Directional Handling Qualities Criteria as Applied to the X-15. NASA SP-90, (see N73-71303 05-99), 1965, pp. 45–60,73N71308. The lateral-directional handling qualities of the X-l5 have been extensively surveyed in terms of pilot ratings and vehicle response characteristics throughout the operational envelope of the airplane. Results are reviewed for two 76 vertical-tail configurations as well as for dampers on and off, and significant problem areas are discussed in relation to the basic stability and control parameters and the influence of the pilot’s control. These results are used to assess the validity and limitations of some of the lateral-directional handlingqualities design criteria currently in use. Finally, a new and more generally applicable criterion recently proposed by the NASA Flight Research Center is described and similarly assessed against a broad range of test conditions available with the X-l5 vehicles. 432. Holleman, E. C.: Control Experiences of the X-15 Pertinent to Lifting Entry. NASA SP-90, (see N73-71303 05-99), 1965, pp. 61–73, 73N71309. (See also 448.) The purpose of this paper is to discuss the flight experiences obtained in recovering the X-15 airplanes from high altitude with conventional and adaptive controls, and to place these experiences in proper perspective relative to future lifting entry programs. 433. Burke, M. E.; and Basso, R. J.: Resume of X-15 Experience Related to Flight Guidance Research. NASA SP-90, (see N73-71303 05-99), 1965, pp. 75–84, 73N71310. The purpose of this paper is twofold. The first is to present a resume of the experience gained to date in using these two systems, and the second is to discuss a planned guidance research program that will be implemented in the near future on the X-15. 434. Adkins, E. J.; and Armstrong, J. G.: Development and Status of the X-15-2 Airplane. NASA SP-90, (see N73-71303 05-99), 1965, pp. 103–115, 73N71313. The original X-15-2 airplane has been extensively modified to provide a Mach 8 configuration. The modifications included jettisonable tanks for additional propellants which would provide the increased performance and consequently would provide a realistic environment for the development and evaluation of a hypersonic air-breathing propulsion system. This paper summarizes the development and initial evaluation of the modified airplane. 435. Watts, J. D.; Cary, J. P.; and *Dow, M. B.: Advanced X-15-2 Thermal Protection System. NASA SP-90, (see N73-71303 05-99), 1965, pp. 117–125, 73N71314. The use of silicone-based elastomeric ablative material for the advanced X-l5-2 thermal protection system is discussed and results of candidate ablator evaluation tests in arc facilities and on X-15 flights at Mach 5 are presented. *NASA Langley Research Center, Hampton, Virginia. 436. Bikle, P. F.; and *McCollom, J. S.: X-15 Research Accomplishments and Future Plans. NASA SP-90, (see N73-71303 05-99), 1965, pp. 133–139, 73N71316. The purpose of this paper is twofold: first to review the overall achievements of the X-15 project with proper emphasis on the highlight of the papers presented and second, to indicate future X-15 plans, both the definitely planned and approved programs and several proposals that are not presently approved but are believed to offer the potential of an excellent return on investment. *U.S. Air Force, Aeronautical Systems Division. 437. Matranga, G. J.; and Walker, J. A.: An Investigation of Terminal Lunar Landing With the Lunar Landing Research Vehicle. NASA TM X-74475. Presented at AIAA Manned Space Flight Meeting, St. Louis, Missouri, October 11–13,1965, 77N74066. 438. Wolowicz, C. H.; and Gossett, T. D.: Operational and Performance Characteristics of the X-15 Spherical, Hypersonic Flow-Direction Sensor. NASA TN D-3070, November 1965, 66N10603, #. The basic design concepts, operational experiences (malfunctions, system characteristics, and system improvements), and flight-data measurements of the sensor are discussed and analyzed. The accuracy of the sensor in measuring angle of attack and angle of sideslip is assessed on the basis of an analysis of flight data and comparisons of these data with X-15 flight data determined from vane-type nose-boom installations and X-15 wind-tunnel data. Some practical limitations in the use of the sensor for extreme altitude applications are also considered. 439. Love, J. E.; and Young, W. R.: Component Performance and Flight Operations of the X-15 Research Airplane Program. NASA TM X-74527. Presented at the Annual Symposium on Reliability, San Francisco, California, January 25–27, 1966, November 1965, 77N74609. This paper discusses and analyzes the system and component failures that have occurred during the X-l5 program. Component performance is expressed in terms of its effect upon the entire operation, that is, as a failure rate per flight. Three representative systems are discussed: the engine system, the auxiliary power system, and the propellant system. Failures of shelf-stock components prior to their installation on the flight vehicles are also examined. 440. McTigue, J. G.; and Thompson, M. O.: Lifting-Body Research Vehicles in a Low-Speed Flight Test Program. NASA TM X-57412. Presented at ASSET/ 77 Advanced Lifting Reentry Technological Symposium, Miami, Florida, 14–16 December 1965, December 1965, 76N70924. The lightweight M-2 flight test program has demonstrated the capability of a pilot to control lightweight lifting body during approach, flare, and landing. Further investigation is needed, however. Areas that are important, and that being investigated, include the use of optical landing systems, night and instrument capability, and thrust-augmented flare. A serious effort is required to reduce the complexity of the aerodynamic control system to prevent the lifting reentry vehicle from being seriously compromised in weight. 441. Stillwell, Wendell H.: X-15 Research Results With a Selected Bibliography, NASA-SP-60, 1965, 65N20162, #. Contents include X-15 aircraft development concept, flight research, aerodynamic characteristics of supersonichypersonic flight, hypersonic structure, flying laboratory, and bibliography. 442. Sisk, T. R.; Irwin, K. S.; and McKay, J. M.: Review of the XB-70 Flight Program. NASA SP-83, NASA Conference on Aircraft Operating Problems, May 10–12, 1965, 1965, 65N31120. Although the major NASA research effort is directed toward XB-70-2, which will not enter its flight program until the summer of l965, a limited amount of information is available from the early flights of the XB-70-l airplane. Initial take-off and landing performance data have generally substantiated predictions and indicate no unforeseen problems for this class of vehicle. Vertical velocities at impact are of the same order of magnitude as those being experienced by present-day subsonic jets. The XB-70 distances from brake release to liftoff graphically illustrate the advantage of the increased thrust-weight ratio of the supersonic cruise vehicle. The landing loads are well within the design limits up to the highest vertical velocities encountered to date, and recorded data show the response at the pilot station to be somewhat greater than that recorded at the center of gravity. Persistent shaking has been encountered in flight at subsonic speed. The cause of the excitation is not known at present but the oscillation does not appear to be conventional buffeting. The oscillation occurrence drops off appreciably at supersonic speeds and can be correlated with atmospheric turbulence. The stability and control characteristics at subsonic speeds appear satisfactory with stability augmentation on and off. A longitudinal trim discrepancy from predictions has been noted in the transonic region which appears to be decreasing with increasing supersonic speed. The supersonic handling qualities are considered adequate with stability augmentation off; however, sensitive lateral control has resulted in small pilot-induced oscillations. Conference at JPL, Jet Propulsion Laboratory, Pasadena, California, November 8–9, 1965. 445. Tanner, C. S.; and McLeod, N. J.: Preliminary Measurements of Take-Off and Landing Noise From a New Instrumented Range. 1965, 65N31110. This paper describes the NASA noise-survey instrumentation system presently in use at Edwards Air Force Base, California, and presents preliminary noise data from an F-104 airplane. Also presented are noise measurements of the XB-70 and 707-131B airplanes obtained with essentially the same equipment at another location. The difference between measured noise levels for the XB-70 and 707 is illustrated and comparisons of perceived noise levels are made. The adequacy of noise predictions is discussed briefly. 1966 Technical Publications EC-16695 XB-70A Airplane 443. Andrews, W. H.; Butchart, S. P.; Sisk, T. R.; and Hughes, D. L.: Flight Tests Related to Jet-Transport Upset and Turbulent-Air Penetration. NASA SP-83, NASA Conference on Aircraft Operating Problems, May 10–12, 1965, 1965, 65N31114. A flight program, utilizing a Convair 880 and a Boeing 720 airplane, was conducted in conjunction with wind-tunnel and simulator programs to study problems related to jet-transport upsets and operation in a turbulent environment. During the handling-qualities portion of the program the basic static stability of the airplanes was considered to be satisfactory and the lateral-directional damping was considered to be marginal without damper augmentation. An evaluation of the longitudinal control system indicated that this system can become marginal in effectiveness in the high Mach number and high dynamic-pressure range of the flight envelope. From the upset and recovery phase of the program it was apparent that retrimming the stabilizer and spoiler deployment were valuable tools in effecting a positive recovery; however, if these devices are to be used safely, it appears that a suitable g-meter should be provided in the cockpit because the high control forces in recovery tend to reduce the pilot’s sensitivity to the actual acceleration loads. During the turbulence penetrations the pilot noted that the measured vibrations of 4 to 6 cps in the cockpit considerably disrupted their normal scan pattern and suggested that an improvement should be made in the seat cushion and restraint system. Also it was observed that the indicator needles on the flight instruments were quite stable in the turbulent environment. 444. Beeler, De E.: NASA Flight Research Center Technical Programs. NASA-Western University 78 446. Jenkins, J. M.; and Sefic, W. J.: Experimental Investigation of Thermal-Buckling Characteristics of Flanged, Thin-Shell Leading Edges. NASA TN D-3243, January 1966, 66N15493, #. The thermal-buckling behavior of a wide range of flanged, thin-shell leading-edge specimens was investigated. Specimens of varying geometry were subjected to temperature-rise up to 50 degrees F per sec (27.7 degrees per sec) and maximum heating rates up to 19.6 Btu/ft 2-sec (222.4 kW/m2). The specimens investigated were constructed of 2024-T3 aluminum, SAE 4130 steel, or Inconel X-750. Regions of stable structural behavior were established on the basis of leading-edge dimensional and thermal-load parameters. Two types buckling were observed in the flanges of most of the specimens. The results of the experiments provide thermal-buckling information from which a variety of flanged, thin-shell leading-edge geometries may be selected that are free of unstable structural behavior while under the influence of severe thermal loadings. 447. McKay, J. M.; and Noll, R. B.: A Summary of the X-15 Landing Loads. NASA TN D-3263, February 1966, 66N15644, #. (See also 430.) The dynamic response of the X-15 airplane at touchdown is reviewed briefly to show the unusual landing characteristics resulting from the airplane configuration. The effect of sinking speed is discussed, as well as the influence of the horizontal-stabilizer load, wing, lift, and increased landing weight on the landing characteristics. Consideration is given to some factors providing solutions to these problems, such as cutout of the stability augmentation damper at gear contact, pilot manipulation of the stabilizer, the use of a stick pusher at touchdown, and a proposed third skid installed in the unjettisoned portion of the lower ventral fin. Studies to determine the effect on the main-landing-gear loads of relocating the X-15 nose gear are discussed. 448. Holleman, E. C.: Control Experiences of the X-15 Pertinent to Lifting Entry. NASA TN D-3262, February 1966, 66N15643, #. (See also 432.) In the program to expand the flight envelope of the X-15 airplane, flights to and entries from altitudes up to 350,000 feet have been accomplished. During these entries, flight-control experience was obtained with four different control-system configurations having varying degrees of complexity. The high steady acceleration and rapidly changing aerodynamic environment did not affect the pilot’s capability to control the entry. All the control systems evaluated were judged by the pilots to be satisfactory for the control of the X-15 entry from the design altitude. Entries have been made that presented more severe control problems than predicted for entries of advanced vehicles at higher velocities. 449. Banner, R. D.; and Kuhl, A. E.: A Summary of X-15 Heat Transfer and Skin Friction Measurements. NASA TM X-1210, Second Annual NASA-University Conference on Manual Control, M.I.T., Cambridge, Massachusetts, February 28–March 2, 1966, February 1966, 71N72665. (See also 428.) Measured local Mach numbers and heat transfer obtained on the lower surface of the X-l5 wing and bottom centerline of the fuselage at angles of attack up to 18° and on the vertical fin with both a sharp and a blunt leading edge are summarized and compared with calculations using Eckert’s referencetemperature method. Direct measurements of skin friction on the surface of the sharp-leading-edge vertical fin are also presented. It is shown that both the heat-transfer and skinfriction data can be predicted by neglecting the effect of wall temperature in the calculation of the reference temperature by Eckert’s method. Uncertainties in level and trend of Reynolds analogy factor with Mach number are discussed, and a planned flight investigation is described. 450. Smith, Harriet J.: Human Describing Functions Measured in Flight and on Simulators. NASA SP-128, Second Annual NASA-University Conference on Manual Control, M.I.T., Cambridge, Massachusetts, February 28– March 2, 1966. (See also 476, 507.) Comparisons have been made between human describing functions measured in flight and on the ground using two different types of ground simulators. A T-33 variablestability airplane was used for the in-flight measurements. The ground tests were conducted in the T-33 airplane on the ground with simulated instrument flight and also on a general-purpose analog computer in conjunction with a contact analog display. For this study a multiple-degree-offreedom controlled element was used in a single-loop compensatory tracking task. The input disturbance in each case consisted of the sum of 10 sine waves with a cutoff frequency of 1.5 radians per second. The results of this investigation indicate no significant difference between the average describing functions measured in flight and those measured in a fixed-base simulator. However, the variance was found to be considerably higher in the flight data. The system open-loop describing functions measured in the fixedbase simulator agreed well with the results of an investigation by McRuer in which the tracking task was similar, although the controlled-element dynamics were different. The average linear coherence was also close to the values found in this same investigation. Contrary to the results of previous investigations, the linear-correlation functions ρ were always equal to 1. 451. Lewis, T. L.; and McLeod, N. J.: Flight Measurements of Boundary-Layer Noise on the X-15. NASA TN D-3364, March 1966, 66N19602, #. (See also 429.) This paper was included in a classified report entitled “Fourth Conference on Progress of the X-15 Research Airplane Program,” Flight Research Center, Oct. 7, 1965. NASA SP-90, 1965 [see chronological numbers 425 through 435]. An appendix has been added to describe the instrumentation, and its frequency response, that was used in obtaining the data. 452. Saltzman, E. J.; and Garringer, D. J.: Summary of Full-Scale Lift and Drag Characteristics of the X-15 Airplane. NASA TN D-3343, March 1966, 66N19345, #. Full-scale power-off flight lift and drag characteristics of the X-15 airplane are summarized for Mach numbers from 0.65 to 6.0 and for free-stream Reynolds numbers from 0.2 × 106 to 2.8 × 106 per foot. Comparisons are made between flight results and the wind-tunnel data that most nearly simulate the full-scale flight conditions. The apparent effect of a sting support on the base pressure of an X-15 wind-tunnel model was propagated onto the vertical-fin base at least one sting diameter above and about one-half sting diameter forward of the sting-model intercept at Mach numbers between 2.5 and 3.5. For the X-15, the effect amounts to from 8 to 15 percent of the base drag between these Mach numbers. For some future vehicles and missions, proper accounting of this interference effect may be necessary to adequately predict the full-scale transonic and supersonic performance. Specifically conducted wind-tunnel-model drag studies, when extrapolated to full-scale Reynolds numbers by the T’ (reference temperature) method, accurately predicted the full-scale zero-lift drag minus base drag of the X-15 at Mach numbers of 2.5 and 3.0. 79 453. Wilson, R. J.: Drag and Wear Characteristics of Various Skid Materials on Dissimilar Lakebed Surfaces During the Slideout of the X-15 Airplane. NASA TN D-3331, March 1966, 66N18172, #. An investigation was made to determine the coefficients of friction and the wear characteristics for X-15 landing gear skids of various materials. Data are presented for skids made of 4130 steel, with and without cermet coating, and Inconel X for several lakebed-surface conditions. The mean coefficient of friction on a dry-hard surface was found to be 0.30 for 4130 steel skids, 0.36 for 4130 steel skids with cermet coating, and 0.35 for Inconel X surface was 0.46; for Inconel skids on a damp surface the mean value was 0.25. Flight data are compared with experimental ground-tow test data on natural and simulated lakebed surfaces. Also included is the variation of skid wear with slideout distance. 454. Wolowicz, C. H.: Analysis of an Emergency Deceleration and Descent of the XB-70-1 Airplane Due to Engine Damage Resulting From Structural Failure. NASA TM X-1195, March 1966, 66N21099, #. An emergency on flight 12 of the XB-70-1 airplane at a Mach number of 2.6 and a pressure altitude of 63,000 feet provided unusual operational, handling qualities, and stability and control data of interest to the supersonic-transport designer. Failure of the wing apex, its ingestion into the right inlet duct, and subsequent damage to the engines produced a steadily deteriorating propulsion situation, which led to resonant vibrations in the relatively flexible fuselage and subsequent stability and control problems in attempting to deal with the vibrations. The results of an analysis of this emergency may be useful in developing adequate operational margins and procedures in the design of the supersonic transport. 455. Barber, Marvin R.; Haise, Fred W.; and Jones, Charles K.: An Evaluation of General Aviation Aircraft Flying Qualities. SAE-Paper-660219, Business Aircraft Conference, Wichita, Kansas, March 30–April 1, 1966. 456. Holleman, E. C.: Summary of High-Altitude and Entry Flight Control Experience With the X-15 Airplane. NASA TN D-3386, April 1966, 66N21041, #. This paper summarizes the high-altitude X-15 flight experience, which culminated in a flight to an altitude of 354,200 feet. Discussed are the basis stability, control, and handling characteristics of the airplane, the cockpit displays, and the operational techniques that enabled it to be successfully flown to and recovered from high altitudes without special piloting aids other than stability augmentation. Flight experience to moderately high altitude with the airplane equipped with interim rocket engines is discussed. 457. Taylor, L. W., Jr.; and Iliff, K. W.: Recent Research Directed Toward the Prediction of LateralDirectional Handling Qualities. NASA TM X-59621, AGARD paper R-531 presented at AGARD 28th Meeting of the Flight Mechanics Panel, Paris, France, May 10–11, 1966, May 1966, 67N23242, #. A survey of lateral-directional handling qualities has been made for the purpose of developing; a technique for predicting pilot ratings. This survey was made by obtaining pilot ratings of lateral control on a fixed-base simulator in conjunction with a color contact analog display. The effect of five lateral-directional handling qualities parameters were studied by systematically varying them over a wide range. Forty-five charts comprise the results of this survey. However these have been condensed into three charts to provide a rapid means for hand computing the pilot ratings. For more accurate predictions a digital computer program was written which incorporated the data from all 45 charts. 458. Berry, D. T.; and Deets, D. A.: Design, Development, and Utilization of a General Purpose Airborne Simulator. NASA TM X-74543, AGARD Paper 529. Presented at AGARD 28th Flight Mechanics Panel, Paris, France, May 10–11 1966. May 1966, 77N74646. 459. Patten, C. W.; Ramme, F. B.; and Roman, J. A.: Dry Electrodes for Physiological Monitoring. NASA TN D-3414, May 1966, 66N25548, #. A method for very rapid application of electrocardiogram electrodes by spraying a conductive mixture is described. The electrodes are also suitable for electroencephalograms. All required equipment and the application procedure are described in detail. The finished electrode is dry and is less than 0.01-inch thick. Electrical and operational factors are not considered. 460. Sadoff, Melvin; Bray, Richard S.; and Andrews, William H.: Summary of NASA Research on Jet Transport Control Problems in Severe Turbulence. Journal of Aircraft, Vol. 3, No. 3, May–June 1966. 461. Andrews, W. H.: Summary of Preliminary Data Derived From the XB-70 Airplanes. NASA TM X-1240, Washington, NASA, June 1966, 66N28013, #. Preliminary data obtained during the initial flight-envelope expansion of the XB-70 airplanes are presented in the areas 80 of stability and control, general performance, propulsionsystem inlet operation, structural thermal response, internal noise, runway noise, and sonic boom. manner in which rawinsonde and rocketsonde data are applied. Flow-angularity effects were avoided by measuring stagnation pressures on a spherical flow-direction sensor, which maintains continuous alignment of the pilot-pressure port with the local flow vector. The quality of the data was further enhanced by applying semiempirical lag corrections to the pressure measurements. The measured density and derived temperature data from the X-15 agree well with rawinsonde data at low altitudes and with Arcas rocketsonde data at higher altitudes. 464. Jenkins, J. M.: A Pretensioning Concept for Relief of Critical Leading-Edge Thermal Stress. NASA TN D-3507, July 1966, 66N30079, #. This paper introduces an analytical concept designed to relieve problems arising from the chordwise temperature gradients by reducing the magnitude of critical compressive stress in a leading edge. The reduction is accomplished by adding an internal column that applies an internal load to the ends of the leading edge. Equations that define the behavior of a pretensioned leading edge are developed and applied to a mathematical model to demonstrate the mechanics of using the concept. 980068 XB-70A Airplane, Three-View Drawing 462. Palitz, M.: Measured and Calculated Flow Conditions on the Forward Fuselage of the X-15 Airplane and Model at Mach Numbers From 3.0 to 8.0. NASA TN D-3447, June 1966, 66N26849, #. Early analyses of X-15 flight heat-transfer data were based on calculated values of the local-flow conditions. The resultant differences between measured and predicted heat transfer were thought to be partially due to an incomplete knowledge of the local fluid properties. Subsequently, a flight investigation was made to determine the extent and character of the local flow on the X-15 airplane in order to aid in the interpretation of the measured heat-transfer data. The results of the flow-field investigation on the forebody of the X-15 are presented and analyzed in this paper. 463. Larson, Terry J.; and Covington, Alan: A Technique for Measuring Mesospheric Densities With the X-15 Research Airplane. Presented at the Fourth Aerospace Sciences Meeting, Los Angeles, California, June 27–29, 1966. Atmospheric-density measurements for altitudes between 30 kilometers and 74 kilometers were obtained during flights with the X-15 research airplane in the southwestern United States. The pilot-pressure method used to derive the densities is discussed in terms of its applicability to X-15 trajectory and instrumentation characteristics. The use of radar tracking to derive X-15 velocity and altitude is described, as well as the 81 465. Powers, B. G.: A Parametric Study of Factors Influencing the Deep-Stall Pitch-Up Characteristics of T-Tail Transport Aircraft. NASA TN D-3370, August 1966, 66N32326, #. This paper presents the results of the program, in which the transport-type aircraft were investigated. A series of stall maneuvers was made with deceleration rates into the stall of 1, 3, and 5 knots per second, with recovery initiated over a range of angle attack. The relative effects of the shape of the pitching-moment curves in the deep-stall region as well as in the initial-stall region were determined in terms of angle-ofattack overshoot and altitude losses during recovery. 466. Roman, James: Flight Research Program-III High Impedance Electrode Techniques. Aerospace Medicine, Vol. 37, No. 8, August 1966. This paper describes electrode techniques designed for largescale flight physiological data collection on a routine basis. Large-scale data collection requires both smaller demands on crew time and less interference with crew comfort than could be achieved by former methods. The resistive components of electrode impedance appears to be related primarily to the extent of skin preparation. For any one method of skin preparation, both resistance and capacitance appear to be primarily a function of electrode area. Motion artifacts are not caused by changes in electrode impedance. Dry electrodes showing a resistive component in excess of 50,000 ohms can be used to obtain tracings of quality comparable, and in some cases superior to those obtained with larger wet electrodes. 467. Quinn, R. D.; and Palitz, M.: Comparison of Measured and Calculated Turbulent Heat Transfer on the X-15 Airplane at Angles of Attack Up to 19.0 Degrees. NASA TM X-1291, September 1966, 73N70599. 468. Barber, Marvin R.; and Haise, Fred W., Jr.: Handling Qualities Evaluation of Seven General Aviation Aircraft. Symposium Proceedings, Society of Experimental Test Pilots, Vol. 8, No. 2, September 23–24, 1966. 469. Mallick, Donald L.; Kluever, Emil E.; and Matranga, Gene J.: Flight Results With a NonAerodynamic, Variable Stability, Flying Platform. Symposium Proceedings, Society of Experimental Test Pilots, Vol. 8, No. 2, September 23–24, 1966. 470. Thompson, Milton O.; Peterson, Bruce A.; and *Gentry, J. R.: Manned Lifting-Body Flight Testing. NASA TM X-59042. Presented at the SETP 10th Symposium, Los Angeles, California, September 1966. *Air Force Flight Test Center, Edwards, California. This paper describes the Joint NASA-Air Force lifting-body flight test program and the three research vehicles, the M2-F2, the HL-10 and the SV-5P. These three vehicles are representative of manned maneuverable reentry spacecraft capable of horizontal landing. 471. Jarvis, C. R.: Operational Experience With the Electronic Flight Control Systems of a Lunar-Landing Research Vehicle. NASA TN D-3689, October 1966, 66N39536, #. Two research vehicles were delivered to the NASA Flight Research Center in the Spring of 1964. After delivery, several months were devoted to checking systems and installing research instrumentation. During this period, many problems were encountered which required extensive modifications to the vehicle and its systems. Subsequent development flight testing disclosed additional problems and resulted in further modifications. This paper discusses the nature of these problems and the performance of the flight control systems during the early flights. 472. Layton, G. P., Jr.; and Dana, W. H.: Flight Tests of a Wide-Angle, Indirect Optical Viewing System in a High-Performance Jet Aircraft. NASA TN D-3690, October 1966, 66N38800, #. A wide-angle, indirect optical viewing system was qualitatively evaluated in an F-104B aircraft as a means of providing visual reference to the pilot. Safe and acceptable performance using the indirect viewing system was demonstrated for all phases of daytime visual flight. Landings were performed in both the conventional and low lift-dragratio configurations. When the horizon was in the field of view, aircraft attitude sensing with the optics was satisfactory 82 about all axes except pitch attitude in climbing flight. This degraded pitch-attitude sensing was due to the poor resolution at the bottom of the field and the lack of view to the sides. A night flight was also performed. The system, in its present form, was considered unacceptable for this use because of large light losses and degraded resolution. It was evident in the study that additional view directly to the side is required for performing circling approaches. 473. Love, J. E.; and Young, W. R.: Survey of Operation and Cost Experience of the X-15 Airplane as a Reusable Space Vehicle. NASA TN D-3732, November 1966, 67N11328, #. The X-15 airplane has been flown more than 150 times in an environment similar to that anticipated for many of the reusable space vehicles being studied. Data are presented on X-15 development and operational costs, turnaround time, and refurbishment cycles, based upon actual operation of the aircraft. For example, 27 flights were accomplished in 1964 at a total cost of $16,268,000, or an average cost of more than $602,000 per flight. It is believed that information from the X-15 program will be helpful in feasibility studies of the reusable-vehicle concept, inasmuch as the X-15 operation is more directly comparable than any other operational program to the reusable systems being considered. 474. Barber, M. R.; Jones, C. K.; Sisk, T. R.; and Haise, F. W.: An Evaluation of the Handling Qualities of Seven General-Aviation Aircraft. NASA TN D-3726, November 1966, 66N39905, #. A review of existing criteria indicated that the criteria have not kept pace with aircraft development in the areas of dutch roll, adverse yaw, effective dihedral, and allowable trim changes with gear, flaps, and power. This study indicated that criteria should be specified for control-system friction and control-surface float. Furthermore, this program suggests a method of quantitatively evaluating the handling qualities of aircraft by the use of a pilot-workload factor. 475. Roman, James; and *Brigden, Wayne H.: Flight Research Program: V. Mass Spectrometer in Medical Monitoring. Aerospace Medicine, Vol. 37, No. 12, December 1966. Mass spectrometers, traditionally large and complicated instruments, have been miniaturized and greatly simplified for the National Space Program. This recent development opens new areas to medicine and to space medicine in particular. The principles of operation of mass spectrometers will soon be important to those engaged in physiological research or in medical monitoring. They are discussed in this paper. A summary of flight test data obtained with a small mass spectrometer in a jet aircraft is presented. *Volt Technical Corporation, NASA Field Team, Edwards, California. 476. Smith, H. J.: Human Describing Functions Measured in Flight and on Simulators. Manual Control, (see N67-15859), 1966, pp. 279–290, 67N15871. (See also 450, 507.) 477. Smith, R. H.; and Schweikhard, W. G.: Initial Flight Experience With the XB-70 Air-Induction System. NASA SP-124, (see N75-71754 05-98), 1966, pp. 185–194, 75N71767. The preliminary results and developmental problems from flight tests of the XB-70 air-induction system are briefly reviewed. The system is generally satisfactory, is adequately matched to the engine flow requirements, and can be controlled for the various flight ranges. Inlet unstarts at cruise Mach number constitute a new problem for high supersonic aircraft seriously affecting the dynamics of the inlet and airframe. 478. *Rolls, L. S.; *Snyder, C. T.; and Schweikhard, W. G.: Flight Studies of Ground Effects on Airplanes With Low-Aspect-Ratio Wings. NASA SP-124, (see N75-71754 05-98), 1966, pp. 285–295, 75N71774. The ground effects on two aircraft with low-aspect-ratio delta wings, the F5D-l and the XB-70A, were measured in flight tests. In a companion program, both small and full-scale models and several wind tunnels were used to document the ground effects for the F5D-l. These flight tests indicated ground effects were not a problem in landing either of these vehicles. The limited wind-tunnel program indicated that scale effects were not of first-order importance in defining ground effects, and that wind-tunnel tests provide reasonable agreement with the values in flight. A simulation study, using a fixed-cockpit projection-type simulator, performed in conjunction with these studies indicated levels of moment and lift changes which would be unsatisfactory from the pilot’s viewpoint; however, some possible alleviating features were noted. *Ames Research Center, Moffett Field, California. 479. Wolowicz, Chester H.: Considerations in the Determination of Stability and Control Derivatives and Dynamic Characteristics From Flight Data. AGARD Report 549, Part I, 1966. This report is the handbook on the determination of stability and control characteristics from flight test. It describes several axis systems, axis transformations, the equations of motions and their limitations, techniques used to determine the mass characteristics of the airplane, the installation and behavior of flight instrumentation, flight test techniques, and the theory and limitations of techniques used to determine the stability and control characteristics from flight data. This report brings all the factors together in the determination of stability and control and provide a ready reference of 83 pertinent information. It is a greatly expanded version of AGARD 224, Stability-Derivative Determination From Flight Data, by Chester H. Wolowicz and Euclid C. Holleman. 1967 Technical Publications 480. Wilson, R. J.: Statistical Analysis of Landing Contact Conditions of the X-15 Airplane. NASA TN D-3801, January 1967, 67N14935, #. The landing contact conditions and slideout distances for 135 landings of the X-l5 research airplane are discussed. The conditions are similar to those that might he experienced by future lifting-body reentry vehicles or other flight vehicles with low lift-drag ratios. Results are presented in the form of histograms for frequency distributions, and Pearson Type III probability curves for the landing contact conditions of vertical velocity, calibrated airspeed, true ground speed, rolling velocity, roll angle, distance from intended touchdown point, and slideout distance. 481. Garringer, D. J.; and Saltzman, E. J.: Flight Demonstration of a Skin-Friction Gage to a Local Mach Number of 4.9. NASA TN D-3830, February 1967, February 1967, 67N17173. A small, commercially available skin-friction gage was flight tested on the X-15 airplane. The Reynolds number range investigated extended from 3.8 × 106 to 10 × 106, and local Mach numbers ranged from 0.7 to 4.9. The ratio of wall-torecovery temperature varied from about 0.4 to 1.4. The gage, its cooling system, and the supporting instrumentation performed well. Turbulent skin-friction values measured in flight for a wide range of wall-to-recovery temperature ratios are similar in level to adiabatic flat-plate and wind tunnel results for corresponding Mach numbers and Reynolds numbers. Thus, for the present tests the influence of wall-torecovery temperature ratio appears to be less than estimated by turbulent theory. 482. Jarvis, Calvin R.: Fly-By-Wire Control System Experience With a Free-Flight Lunar-Landing Research Vehicle. AIAA Paper 67-273, AIAA Flight Test, Simulation, and Support Conference, Cocoa Beach, Florida, February 6–8, 1967. 483. Matranga, Gene J.; Mallick, Donald L.; and Kluever, Emil E.: An Assessment of Ground and Flight Simulators for the Examination of Manned Lunar Landing. AIAA Paper 67-238, AIAA Flight Test, Simulation, and Support Conference, Cocoa Beach, Florida, February 6–8, 1967. 484. Roman, James, Perry, John J.; Carpenter, Lewis R.; and *Awni, Shiban: Flight Research Program: VI. Heart Rate and Landing Error in Restricted Field of View Landing. Aerospace Medicine, Vol. 38, No. 3, February 1967. Two pilots were instrumented for electrocardiogram in a T-33 jet aircraft in the course of eleven flights in which pilot horizontal field of view was varied from 360 degrees to 5.7 degrees. Landing error was recorded in terms of distance from the desired touchdown point. A high degree of correlation was found to exist between heart rate and landing error. There was no significant correlation between heart rate and field of view, nor was there significant correlation between field of view and landing error for the fields of view tested. At the 5.7 degree field of view the monocular fields of view did not overlap, so that only one eye could be used. Landing error did not increase significantly when only one eye was used. This finding has implications with respect to aeromedical standards. *Computing and Software, Incorporated, Panorama City, California. 485. Roman, James; *Older, Harry; and **Jones, Walton L.: Flight Research Program: VII. Medical Monitoring of Navy Carrier Pilots in Combat. Aerospace Medicine, Vol. 38, No. 2, February 1967. The feasibility of medical monitoring in combat was demonstrated by instrumenting ten dive-bombing missions from a Navy attack aircraft carrier operating in the Gulf of Tonkin. Nine missions suitable for data analysis were obtained. The results were remarkable primarily for the low heart rates seen on these opposed missions. The overall heart rate for 18 hours of data was 87.6 beats per minute. The heart rates at launch and recovery were substantially higher than the bombing heart rates, in spite of the significant normal acceleration experienced during the bomb runs. The difference between launch or recovery, and bombing was statistically highly significant. Comparisons between the first and the second combat missions of the day for the same pilots on the same day showed heart rate to be substantially lower on the second mission. The difference was statistically significant. The pilots were of an unusually high experience level, and the data presented could not be considered representative for a pilot group of average combat experience, or average carrier operations experience. *Consultant to NASA, Washington, D. C. **NASA Headquarters, Washington, D. C. 486. Gord, P. R.: Measured and Calculated Structural Temperature Data From Two X-15 Airplane Flights With 84 Extreme Aerodynamic Heating Conditions. NASA TM X-1358, H-442, March 1967, 74N71363. This paper presents structural temperature data from two flights of the X-15 airplane in which extreme aerodynamicheating conditions were experienced. These flights, shown on the X-15 flight envelope represent the extremes of maximum dynamic pressure and maximum altitude achieved during X-15 flights to date. The temperature data recorded are presented in tabulated form for 103 locations of the airplane. 487. Taylor, L. W., Jr.: A Comparison of Human Response Modeling in the Time and Frequency Domains. NASA TM X-59750. Presented at the USC and NASA Conference on Manual Control, Los Angeles, California, March 1–3 1967, (previously announced as N68-25276), 1967, 68N37735. (See also 511.) Frequency and time domain methods of analyzing human control response while performing compensatory tracking tasks are reviewed. Sample linear model results using these methods are compared and discussed. The inherent requirement of constraining the freedom of the form of the pilot models is also discussed. The constraint in the frequency domain consists of smoothing with respect to frequency; whereas, the constraint for the time domain model is more natural and meaningful in that it consists simply of limiting the memory of the pilot model. The linear models determined by both methods were almost identical. 488. Jarvis, C. R.: Flight-Test Evaluation of an On-Off Rate Command Attitude Control System of a Manned Lunar-Landing Research Vehicle. NASA TN D-3903, April 1967, 67N23293, #. This paper deals specifically with the evaluation of the capability of an on-off rate command attitude control system to provide satisfactory control for maneuvering a vehicle in a lunar-gravity environment. Control boundaries, based on pilot ratings, are established from fixed-base simulator studies that define satisfactory valued of rate dead band, controller sensitivity, and angular acceleration. These boundaries are then compared to flight results obtained with the LLRV. Results are presented for both Earth and lunaroriented operation. 489. Adkins, Elmor J.: X-15 Research Program Accomplishments and Plans. Presented at the AIAA Symposium of Hypersonic Flight, Los Angeles, California, April 27, 1967. 490. Saltzman, Edwin J.; and Hintz, John: Flight Evaluation of Splitter-Plate Effectiveness in Reducing Base Drag at Mach Numbers From 0.65 to 0.90. NASA TM X-1376, May 1967, 67N26558, #. An experiment has been conducted to determine the effectiveness of a splitter plate in reducing base drag at subsonic speeds. The test configuration was a “fin-like” shape mounted on the under belly of a F-104 which may be representative of blunt-trailing-edge stabilizing surfaces of future hypersonic aircraft or reentry vehicles. The test chord Reynolds numbers, up to 36.2 × 106, are believed to be representative of chord Reynolds numbers for the terminal, subsonic phase of a lifting-body reentry. The splitter plate which extended into the wake a distance of 1 base width, reduced the negative base pressure coefficients between 30 percent and 40 percent. This increment in base pressure coefficient was as large as obtained on a two-dimensional wind- tunnel model at the higher comparable Mach numbers and about 12 percent lower at the lower comparable Mach numbers, even though the flight results represented higher Reynolds numbers and contained outboard end (threedimensional) effects. 491. Szalai, K. J.: The Influence of Response Feedback Loops on the Lateral-Directional Dynamics of a VariableStability Transport Aircraft. NASA TN D-3966, Washington, NASA, May 1967, Refs, May 1967, 67N26545. Several response feedback loops are analyzed to determine their effects on the lateral-directional dynamics of a variablestability transport aircraft. The response feedback system feeds back response variables such as sideslip angle or roll rate as rudder or aileron commands, or both, thus altering the various transfer functions which describe the dynamic characteristics of the aircraft. The range of the feedback gain for which approximate expressions are valid describing the effect of a particular loop is noted. The root-locus method is used to show the dutch roll, roll, and spiral modes are influenced as a function of feedback gain. Expressions are developed which directly relate feedback gains to some response parameter, such as dutch roll frequency, damping ratio, or roll and spiral mode time constants. The expansion to multiloop systems is discussed, along with limitations of the response feedback system from the standpoint of operating a variable-stability aircraft. 492. Lytton, L. E.: Evaluation of a Vertical-Scale, Fixed-Index Instrument Display Panel for the X-15 Airplane. NASA TN D 3967, May 1967, 67N25037,#. A comparative evaluation was performed on an analog simulator to compare pilot performance when using the operational X-15 instrument panel and a panel incorporating vertical-scale, fixed-index flight instruments. The purpose of the evaluation was to provide experiential evidence to complement pilot opinion concerning the acceptability of the vertical-scale panel for use in the X-15 airplane. This evidence was obtained in the form of a wide variety of performance measures for 16 subjects, for two different representative mission profiles, and over three trials or runs 85 for each profile. The data were subjected to both parametric and nonparametric statistical analysis. 493. Love, James E.; and Young, William, R.: Operational Experience of the X-15 Airplane as a Reusable Vehicle System. Presented at SAE 2nd Annual Space Technology Conference, Palo Alto, California, May 9–11, 1967. 494. Wilson, R. J.; and Larson, R. R.: Statistical Analysis of Landing-Contact Conditions for the XB-70 Airplane. NASA TN D-4007, June 1967, 67N27617, #. Landing-contact conditions for 71 landings of the XB-70 airplanes are analyzed. Some of the conditions are similar to those that may be experienced by future supersonic vehicles. Results are presented as frequency histograms and cumulative frequency distributions in terms of probability. The landing-contact parameters examined include vertical velocity; indicated airspeed; angles of roll, pitch, attack, and sideslip; and rolling and pitching velocities. 495. Perry, J. J.; Dana, W. H.; and Bacon, D. C., Jr.: Flight Investigation of the Landing Task in a Jet Trainer With Restricted Fields of View. NASA TN D-4018, June 1967, 67N27294, #. A total of 155 landings were made in a T-33A jet aircraft in order to determine the relationship between the pilot’s field of view and his performance of the landing maneuver. The field of view was reduced from unrestricted to a minimum of 0.10 radian (5.7 degrees) horizontal and 0.52 radian (30 degrees) vertical. The pilot’s task was to fly a 180 poweron pattern and final approach and land the aircraft on a predetermined point on the runway. Also, power-off 360 degrees overhead and straight-in approaches were performed by one of the pilots. The quality of the performance of the power-on task was measured by recording touchdown error. Pilot comments were obtained for all flights. 496. Lewis, Charles E., Jr.; Jones, Walton L.; *Austin, Frank; and Roman, James: Flight Research Program: IX. Medical Monitoring of Carrier Pilots in Combat – II. Aerospace Medicine, Vol. 38, No. 6, June 1967. Cardiorespiratory functioning in flight was monitored on Naval aviators flying bombing missions against heavily defended targets in North Vietnam. Thirty-one missions suitable for data analysis were obtained. Continuous records of ECG, respiratory rate, acceleration and voice were recorded in flight. Both day and night missions were monitored. The pilots studied were of an unusually high experience level, averaging 1,952 total flying hours and 104 combat missions per man. The overall combat heart rate was 94.9 bpm. Overall bombing heart rate was 112.3 bpm, including day and night bombing; frequently in bad weather. Overall respiratory rate was 22.9 breaths per minute. In a comparison study on Marine reserve pilots, gravitational stress was determined to be of importance in elevating the bombing heart rate observed in this combat study. The stresses of combat flying, particularly the element of risk, is clearly shown to be ineffectual in evoking cardiovascular response in the group studied. *Capt., U. S. Marine Corps. 497. Taylor, L. W., Jr.; and Balakrishnan, A. V.: Identification of Human Response Models in Manual Control Systems. NASA TM X-60204. Presented at the IFAC Symposium on the Probl. of Identification of Automatic Control Systems, Prague, June 12–17, 1967, 1967, 68N25712, #. Frequency domain and time domain methods of analysis are reviewed with their regard to their application toward identifying pilot models. The models would subsequently be used to study the stability and performance of a man-machine system in which the human controller performs a compensatory tracking task. Sample linear model results are compared and discussed. The inherent requirement constraining the freedom of the form of the pilot model is also discussed. 498. Noll, R. B.; and McKay, J. M.: Theoretical Dynamic Analysis of the Landing Loads on a Vehicle With a Tricycle Landing Gear. NASA TN D-4075, August 1967, 67N32394, #. A theoretical analysis is presented for the landing dynamics of a vehicle equipped with a tricycle landing-gear system. The equations are simplified in order to provide a more convenient yet adequate analysis for most vehicles. The adequacy of the simplified analysis for simulating the landing dynamics and loads of a vehicle is illustrated by comparing results of calculations with flight-test data from the X-15 research airplane. The feasibility of using the modified analysis for investigating off-design landing contingencies is demonstrated by examples of studies performed for the X-15. 499. Pyle, J. S.; and Swanson, R. H.: Lift and Drag Characteristics of the M2-F2 Lifting Body During Subsonic Gliding Flight. NASA TM X-1431, August 1967, 71N70184. The subsonic flight lift and drag characteristics of the M2-F2 lifting-body configuration are presented at angles of attack from –4 degrees to 16 degrees. Flight results are compared with data obtained from full-scale wind-tunnel tests on the flight vehicle and with M2-F1 flight results. 86 M2-F2 Lifting Body 500. Beeler, De E.: Optimization of Aircraft Performance and Mission Completion Through Research on the Pilot and Aircraft as an Overall System. Aerospace Proceedings 1966, Royal Aeronautical Society, Centenary Congress and International Council of the Aeronautical Sciences, Congress, 5th, London, England, September 12– 16, 1966, Vol. 2, (see A66-42492), 1967, pp. 909–930, 68A19806. The philosophies and findings of the research aeroplane programme are in many respects directly applicable to the requirements for the development of successful future highperformance aircraft. One of the lessons could be the importance of an experimental or prototype aircraft as a basic requirement to the successful development of proposed aircraft of the future. 501. Gaidsick, H. G.; Layton, G. P., Jr.; and Dana, W. H.: Indirect Pilot Viewing for Reentry Vehicles and SSTs. Space/Aeronautics, Vol. 48, September 1967, pp. 118–121, 68A14761. Indirect pilot viewing systems for reentry vehicles and SST, discussing overlapping monoculars and panoramic display. 502. *McDonald, R. T.; and Roman, J.: Development of Respiration-Rate Transducers for Aircraft Environments. NASA TN D-4217, November 1967, 67N39753, #. Two types of sensors for monitoring respiration rate in aircraft environments have been developed: a low-pressure pneumotachometer designed to monitor the pilot’s respiration rate in aircraft that have a low-pressure breathingoxygen supply, and a high-pressure pneumotachometer designed to monitor the pilot’s respiration rate in aircraft with ECN-1088 a high-pressure breathing-oxygen supply. For both pneumotachometers, the sensor is placed in series with the ’ oxygen supply line and the pilot s oxygen line. The sensor detects gas flow that accompanies inspiration. *Northrop Corporation, Hawthorne, California. 503. Kordes, E. E.; and Love, B. J.: Preliminary Evaluation of XB-70 Airplane Encounters With HighAltitude Turbulence. NASA TN D-4209, November 1967, 67N39719, #. Measurements of airplane response to clear-air turbulence were obtained during supersonic flights of the XB-70 airplanes to an altitude of 74,000 feet (22,555 meters) over the Western United States. In general, the results for 75,757 miles (121,919 kilometers) of operation above 40,000 feet (12,192 meters) altitude show that turbulence was encountered an average of 7.2 percent of the miles flown between 40,000 feet (12,192 meters) and 65,000 feet (19,812 meters) and an average of 3.3 percent of the miles flown above 65,000 feet (19,812 meters) with less than l percent of the turbulent areas exceeding 100 miles (160.93 kilometers) in length. Power-spectral-density estimates of the acceleration response to turbulence show that the structural modes contribute an appreciable amount to the total response. 504. Montoya, E. J.; and Palitz, M.: Wind-Tunnel Investigation of the Flow Field Beneath the Fuselage of the X-15 Airplane at Mach Numbers From 4 to 8. NASA TM X-1469, November 1967, 68N11147, #. Wind-tunnel data were obtained on the local flow field beneath the fuselage of a model of the X-15 airplane approximately 5 to 8 fuselage diameters aft of the model nose from the model surface to the bow shock. Multipletube rakes, model surface pressure orifices, and a cone probe were used to survey the flow field. The cone probe was used to obtain Mach numbers in the flow field up to a free-stream Mach number of 6 and to obtain flow angularity up to a Mach number of 8. Test results were obtained at free-stream Mach numbers from 4 to 8 and angles of attack from –3 degrees to 20 degrees and were compared with theory and flight data. 505. Roman, J.; and Sato, R. N.: A Useful Modification of the Wright Spirometer. NASA TN D-4234, NASA, November 1967, 68N10057, #. The Wright spirometer is a useful gas flowmeter for physiological use, in that it is small and reliable. However, data collected with this device must be reduced manually, and the calibration curve is nonlinear at low flow values. The instrument was modified by fitting the output shaft with a spooked wheel that interrupts the light beam between a low-power light source and a photonsensor. This modification provides a digital electrical output that can be 87 computer-reduced, permitting correction of nonlinearity of the calibration curve. The 96 milliwatts, which is small enough to be battery supplies of most self-contained recorders. the data for the power drain is drawn from the miniature tape 506. Taylor, L. W., Jr.; and Smith, J. W.: An Analysis of the Limit-Cycle and Structural-Resonance Characteristics of the X-15 Stability Augmentation System. NASA TN D-4287, December 1967, 68N11545, #. This paper considers in some detail the limit-cycle and structural-resonance problems by using nonlinear mathematical models in the analysis of the system stability. The results of the analysis are compared with results obtained from ground flight tests. Limit cycle calculations involved multiple, nonseparable, nonlinear elements which demonstrate the use of describing functions. 507. Smith, H. J.: Human Describing Functions Measured in Flight and on Simulators. IEEE Transactions on Human Factors in Electronics, Vol. HFE-8, December 1967, pp. 264–268, 68A20662. (See also 450, 476.) Human describing functions measured in flight and on simulators, noting difference in variances. 508. Thompson, M. O.; Weil, J.; and Holleman, E. C.: An Assessment of Lifting Reentry Flight Control Requirements During Abort, Terminal Glide, and Approach and Landing Situations. NASA TM X-59119. Presented at Specialists meeting on Stability and Control, Cambridge, England, September 20–23, 1966, 1967, 68N27404. The results of the X-15 research airplane and M-2 lifting body flight programs and various simulation programs are summarized for pertinence to the control requirements for manned lifting reentry. Piloted reentries have been successfully accomplished with several degrees of controlsystem sophistication and at a variety of reentry conditions some more severe than expected during orbital reentry. 509. Walker, H. J.; and Thompson, M. O.: Handling Qualities of Hypersonic Cruise Aircraft. Conference on Hypersonic Aircraft Technology, 1967, pp. 155–169, 74N73060. 510. Reed, R. D.: Flight Testing of Advanced Spacecraft Recovery Concepts Using the Aeromodeler’s Approach. NASA Langley Research Center Inter-Agency Flexible Wing Technology Meeting, 1967, 86N72377. (See also 522.) Model flight investigations are being conducted at the NASA Flight Research Center to explore various spacecraft terminal-landing concepts. Modern model-airplane equipment and procedures have contributed to many successful flights on radio-controlled experimental gliding parachute models launched from a large radio-controlled launch model. In addition, three flights of a similar largescale parachute model were made with the aid of a helicopter. Launches took place approximately 1000 feet above the ground with the radio-controlled launch model and 4000 feet above the ground with the helicopter. Studies were made of parawing deployment transients, steering control, and model landing impact. Experiments with parawings in combination with the aerodynamically stable lifting-body payloads revealed potential directional-coupling characteristics at fullscale conditions. Also, approximately 20 tests on advanced variable-geometry free-flight and radio-controlled bodies provided some insight into low-speed stability and control characteristics with conventional wings “stowed” and “extended.” The information obtained was primarily qualitative, based on movies of the flight tests. Control inputs were recorded to provide some quantitative data on parawing turning rates as a function of control input. 511. Taylor, L. W., Jr.: A Comparison of Human Response Modeling in the Time and Frequency Domains. Three-D Annual NASA University Conference on Manual Control, 1967, pp. 137–153, (see N68-15901 06-05), 68N15910, #. (See also 487.) Frequency and time-domain methods of analyzing human control response while performing compensatory tracking tasks are reviewed. Sample linear model results using these methods are compared and discussed. The inherent requirement of constraining the freedom of the form of the pilot models is also discussed. The constraint in the frequency domain consists of smoothing with respect to frequency; whereas, the constraint for the time domain model is more natural and meaningful in that it consists simply of limiting the memory of the pilot model. The linear models determined by both methods were almost identical. 512. Taylor, L. W., Jr.: Relationships Between Fourier and Spectral Analyses. Three-D Annual NASA University Conference on Manual Control, 1967, pp. 183–186, (see N68-15901 06-05), 68N15913, #. About 2-1/2 years ago the Flight Research Center was preparing to analyze human response data for a joint NASAUSAF-Cornell program using a ground based simulator and the variable-stability T-33 airplane. A decision to use expressions of the cross- and power-spectral density functions involving Fourier transforms instead of the crossand auto-correlation functions led to certain simplifications which raised some questions. 88 513. Fischel, J.; and Gee, S. W.: Aeronautical FlightControl Systems Research. NASA SP-154, 1967, (see N68-33169 20-30), pp. 245–259, 68N33186. The primary flight control system configuration now being used, an electromechanical-hydraulic combination, has proved to be fairly reliable; however, this system is complex and has several inherent undesirable features and attendant problems. The increasing complexity of flight control systems resulting from the increased performance of future aircraft can be best resolved electrically by a fly-by-wire system. Consideration should be given to the integration and simplification of display parameters to alleviate pilot effort and for the display of new parameters providing essential information for improved flight control by the pilot. When this is accomplished, we can look toward completely automatic control and, possibly, remote control. 514. Gee, S. W.: A Review of Avionics Requirements for General Aviation. NASA SP-154, Aerospace Electronic Systems Technology, 1967, (see N68-33169 20-30), pp. 261– 272, 68N33187, #. 1968 Technical Publications 515. Taylor, L. W., Jr.: Nonlinear, Time-Domain Models of Human Controllers. NASA TM X-60996. Presented at Hawaii International Conference on System Sciences, Honolulu, Hawaii, January 29–30, 1968, January 1968, also Journal of Optimization Theory and Applications, January 1968, 68N28920, #. This paper presents results of analyses and discusses the method of selecting maximum memory time and order of the nonlinear model. In addition, there was discussion and results of orthogonal expansion of the weighting functions for reasons for data compression and reduced computation. 516. Dana, W. H.: From the Pilot’s Seat. Science News, February 24, 1968, pp. 188–189. 517. Thompson, M. O.; and Dana, W. H.: Flight Simulation of Night Landings of Lifting Entry Vehicles. AIAA Paper 68-259. Presented at the 2nd AIAA Flight Test, Simulation and Support Conference, Los Angeles, California, March 25–27, 1968, March 1968, 68A23680, #. As part of an evaluation of the operational capabilities of lifting entry vehicles, night approaches and landings were performed in fighter aircraft configured to provide a maximum L/D ratio of approximately 3.0. These approaches were performed to a lighted runway and to a dry lake bed illuminated by airborne parachute flares. Approach patterns were 270 degree overhead patterns begun at altitudes of 30,000 to 45,000 ft. Approach pattern control was accomplished by the pilot. Moonlight varied from full moon to no moon. Each night flight was paired with an identical flight flown the preceding afternoon. The approaches are compared with one another and with their daytime counterparts using radar tracking plots and touch-down miss distance as evaluation criteria. Landing-site comparison (runway or lake bed) is by pilot comment. Effects of moonlight on pattern control are discussed. Spot landing miss distances for the entire program are presented as evidence of the feasibility of night landings in lifting entry vehicles. 518. Bogue, R. K.; and Webb, L. D.: Advanced Air Data Sensing Techniques. NASA TM X-61115, 1968, 68N37326, #. (See also 519.) Determination of the feasibility of using a fluidic-type temperature sensor for measuring total temperature on an aircraft traveling at hypersonic speeds within the atmosphere. Problem areas that do not severely limit the application but do require further study to assess the total effect of each problem in the system operation were revealed. A partial totaltemperature time history of X-l5 flight 53 (Oct. 3, 1967) up to peak Mach number is shown, together with radar measurements of the X-l5 Mach number and altitude, in addition to the data obtained from the shielded thermocouple sensor and the fluidic sensor. 519. Bogue, R. K.; and Webb, L. D.: Advanced Air Data Sensing Techniques. Presented at the International Aerospace Instrumentation Symposium, 5th, Cranfield, Beds., England, March 25–28, 1968, Proceedings, 14 refs., (A69-16747 05-14), 1968, pp. 66–75, 69A16755, #. (See also 518.) 520. Taillon, N. V.: A Method for the Surface Installation and Fairing of Static-Pressure Orifices on a Large Supersonic-Cruise Airplane. NASA TM X-1530, March 1968, 68N19341, #. A method for installing and fairing static-pressure orifices on the wing surface of a supersonic airplane without penetrating the skin is described. Orifice discs were fixed to pressure tubes which were, in turn, attached to the ferrous skin by welded straps. The assembly was faired over with a temperature-resistant aerodynamic smoothing compound hand-milled flush with the orifices. Some deviation from the mold line is inherent in the method; however, analytical estimates indicate that the effect on local aerodynamic pressures is negligible for this installation. The smoothing compound has been found to be operationally suitable at a Mach number of 3. 521. Van Leynseele, F. J.: Evaluation of LateralDirectional Handling Qualities of Piloted Reentry Vehicles Utilizing a Fixed Base Simulation. NASA TN D-4410, March 1968, 68N19226, #. A simulator investigation was conducted to evaluate the lateral-directional handling qualities of piloted vehicles. The 89 lateral-directional parameters were chosen to represent a sample of dynamic characteristics typical of reentry-vehicle configurations. The evaluations were made by using a threedegree-of-freedom fixed-base simulator with a pseudooutside world visual display (contact analog). The investigation showed that the pilots preferred the ratio of the roll transfer function numerator frequency to the dutch roll frequency to be unity, independent of the magnitude of bank angle to sideslip angle ratio. They objected to an excessive amount of sideslip angle excitation with ailerons when the ratio of the roll transfer function numerator frequency to the dutch roll frequency differed from unity, the bank angle to sideslip angle ratio was low, and the yawing moment due to aileron was large. 522. Reed, R. D.: Flight Testing of Advanced Spacecraft Recovery Concepts Using the Aeromodeler’s Approach. AIAA Paper 68-242. Presented at the 2nd AIAA and Flight Test Simulation and Support Conference, Los Angeles, California, March 25–27, 1968, March 1968, 68A23665, #. (See also 510.) 523. Lasagna, P. L.; and McLeod, N. J.: Preliminary Measured and Predicted XB-70 Engine Noise. NASA TM X-1565, April 1968, 68N21834, #. This paper presents measured and predicted noise levels and computed perceived noise levels for the XB-70 airplane during takeoffs, a landing, and a flyby at Edwards Air Force Base, Calif. The SAE jet-noise prediction method was used to predict noise levels for comparison with measured values. 524. Carpenter, R.; and Roman, J.: Recording and Signal-Conditioning Techniques and Equipment Used in a 1,000-Flight Biomedical Study. NASA TN D-4487, April 1968, 68N21538, #. The NASA Flight Research Center recently concluded a biomedical monitoring program involving 1,000 flights in high-performance aircraft by students of the USAF Aerospace Research Pilot School and by NASA aerospace research pilots. To permit accurate and reliable data acquisition of electrocardiogram (ECG), respiration rate, and normal acceleration, it was necessary to design and develop a means of reliably recording and transcribing flight medical data in a format compatible with computer reduction. Signal conditioners and interconnecting harnesses were designed and fabricated, and guidelines were established for the construction of a five-channel analog tape recorder to record these data while the recorder is being carried on the pilot with minimum interference or discomfort. The equipment operated reliably and enabled satisfactory data acquisition of biomedical information both in extended biomedical instrumentation studies and in remote-site medical monitoring. 525. Carpenter, R.; and Roman, J.: FM Handling and Analog-to-Digital Conversion of Biomedical Data From a 1,000-Flight Study. NASA TN D-4488, April 1968, 68N21537. To collect, process, and analyze FM-recorded biomedical data from 1,000 flights in high-performance aircraft and test vehicles, it was necessary to devise a handling facility that would prepare these data in a standard format for high-speedcomputer processing. The handling system designed maintains the very high signal-to-noise radio inherent in the original data-acquisition equipment, provides push-button control for converting the medical information into a standard format for digital processing at either four or eight times faster than the original record speed, and provides an effective number of quality-control checkpoints. The system is described in detail, and system design considerations are discussed in relation to preventing data degradation in both FM handling and digital conversion. 526. Ehernberger, L. J.: Meteorological Aspects of High-Altitude Turbulence Encountered by the XB-70 Airplane. NASA TM X-61114, 1968, 68N37298, #. (See also 527.) This paper discusses the preliminary results of a study of meteorological features associated with turbulence encountered by the XB-70 airplane at flight levels above 40,000 feet (12,200 meters). Also, three of the larger temperature transients encountered during level flight at high altitudes are described. This study was conducted at the NASA Flight Research Center, Edwards, California, and covers XB-70 airplane flights made between April 1965 and March 1966 over the Western United States. 527. Ehernberger, L. J.: Meteorological Aspects of High-Altitude Turbulence Encountered by the XB-70 Airplane. Proceedings, 3rd National Conference On Aerospace Meteorology, New Orleans, Louisiana, May 6–9, 1968, (A68-35067 17-20), pp. 515–522. Boston, Massachusetts, American Meteorological Society, Conference Sponsored by the American Meteorological Society, the American Institute of Aeronautics and Astronautics, and the Institute of Environmental Sciences, 1968, 68A35132, #. (See also 526.) 528. Wolowicz, C. H.; Strutz, L. W.; Gilyard, G. B.; and Matheny, N. W.: Preliminary Flight Evaluation of the Stability and Control Derivatives and Dynamic Characteristics of the Unaugmented XB-70-1 Airplane Including Comparisons With Predictions. NASA TN D-4578, May 1968, 68N24498, #. Stability and control characteristics of the XB-70-1 airplane were evaluated from data obtained during the early phases of the flight-test program at Mach numbers extending to 2.56 and altitudes to 64,700 feet (19,700 meters). This report summarizes the results of the evaluation and compares the 90 flight-determined derivatives with those obtained from windtunnel tests and with estimated effects of aeroelasticity. 529. Powers, B. G.: A Review of Transport HandlingQualities Criteria in Terms of Preliminary XB-70 Flight Experience. NASA TM X-1584, May 1968, 68N23901, #. A preliminary flight evaluation of handling qualities of the unaugmented XB-70 airplane was made during the initial flight test and envelope-expansion program. The evaluations consisted of pilot ratings and comments on the longitudinal and lateral-directional characteristics. The pilot ratings were compared with several current handling-qualities criteria for transport aircraft to establish the applicability of these criteria to this class of airplane. 530. Sisk, T. R.; Matheny, N. W.; Kier, D. A.; and Manke, J. A.: A Preliminary Flying-Qualities Evaluation of a Variable-Sweep Fighter-Type Aircraft. NASA TM X-1583, H-509, May 1968, 75N70035. An evaluation of F-111A airplane number 6 (S/N 63-9771) consisting of 9 pilot-familiarization flights and 14 dataacquisition flights extending to Mach numbers approaching 1.9 at 47,500 feet (14,478 meters) altitude and 1.1 at 10,000 feet (3048 meters) altitude were completed. This preliminary evaluation allowed the assessment of flightcontrol-system and airplane response characteristics as various wing sweeps over the normal operating envelope of the aircraft with stability augmentation on and off. Augmentation-off flight is considered to be outside the normal flight environment of an operational aircraft and to be an emergency condition. 531. Watts, Joe D.; and Olinger, Frank V.: Heat Transfer Effects of Surface Protuberances on the X-15 Airplane. NASA TM X-1566, H-507, May 1968, 75N70033. Heat-transfer effects of separated flow were investigated in flight tests of two protuberance configurations on the X-15 airplane. The 0.20-inch forward-and-aft-facing step and the 0.20-inch-amplitude sine-wave oriented at a right angle to the stream direction resulted in local heat-transfer variation of 0.09 to 2.23 and 0.34 to 2.03 times the smooth surface value, respectively. 532. Webb, L. D.: Characteristics and Use of X-15 Air-Data Sensors. NASA TN D-4597, June 1968, 68N25317, #. The uses, techniques of correlation, and analysis of flightguidance and air-data sensors that have been flown on the X-15 airplane are examined. Methods by which meteorological balloons and high altitude rocketsondes were used to define the atmospheric envelope around the X-15 airplane are discussed. The application of onboard sensor data, meteorological data, and radar in obtaining altitude, velocity, Mach number, and dynamic pressure is explained. 533. Holleman, E. C.: Stability and Control Characteristics of the M2-F2 Lifting Body Measured During 16 Glide Flights. NASA TM X-1593, June 1968, 70N78443. Sixteen glide flights with the M2-F2 lifting-body research vehicle were analyzed to obtain a measure of some of the static and dynamic stability and control and handling characteristics for a Mach number range of 0.4 to 0.7. The vehicle was statically and dynamically stable in the regions in which it was predicted to be stable. The upper flap was about twice as effective as the lower flap as a pitch control. The flight stability and control results agreed reasonably well with the wind-tunnel predicted characteristics. The M2-F2 handling qualities with dampers on and rudder-to-aileron interconnect operative were rated satisfactory for the M2-F2 research mission by the four pilots in the program. The predicted unacceptable handling characteristics of the basic vehicle were observed in flight. Various handling-qualities criteria predicted handling that was in general agreement with the actual pilot evaluation for the M2-F2 vehicle. The vehicle has the lift capability and maneuverability for satisfactory approach and landing as a glider at a selected landing site. The approach and landing piloting task was demanding and required detailed preparation and practice for the flight and complete concentration during the maneuver. 534. Wolowicz, Chester H.; and Wykes, John H.: Stability Derivatives of an Elastic Airplane From Flight Test. 1968 Seminar on Elastic Airplane Stability, Control, and Response, University of Kansas, June 12, 1968. 535. Smith, R. H.; Bellman, D. R.; and Hughes, D. L.: Preliminary Flight Investigation of Dynamic Phenomena Within Air Breathing Propulsion Systems of Supersonic Aircraft. AIAA Paper 68-593. Presented at the 4th AIAA Propulsion Joint Specialist Conference, Cleveland, Ohio, June 10–14, 1968, June 1968, 68A33790, #. Many of the aerodynamic conditions that contribute to the propulsion system problems of aircraft are dynamic and require higher response instrumentation than is generally used for flight-test work. Such problems become increasingly prominent as aircraft advance into the supersonic speed region where the function and control of the inlet and engine become more critical. At the NASA Flight Research Center, the F-111A airplane and the XB-70A airplane, both capable of flying at Mach numbers greater than 2, have been instrumented to measure pressure fluctuations in the propulsion-system airstream at frequencies as high as 200 hertz. This paper presents design and development details of the two instrumentation systems as well as their characteristics as shown by laboratory tests. Temperature effects on the pressure transducers and means for their compensation and correction are discussed. XB-70A flight results for compressor-face pressures are included. 536. Beaulieu, W.; Campbell, R.; and Burcham, W.: Measurement of the XB-70A Propulsion Performance Incorporating the Gas Generator Method. AIAA Paper 68-594, Propulsion Joint Specialist Conference, Cleveland, Ohio, June 10–14, 1968, June 1968, 68A33791. (See also 578.) Propulsion performance of XB-70A aircraft calculated by gas generator method. 537. Burke, M. E.: X-15 Analog and Digital Inertial Systems Flight Experience. NASA TN D-4642, July 1968, 68N29404, #. Two different types of inertial flight data systems, an analog system and a digital system, have been used during the X-15 program to provide primary flight information for the X-15 pilot. This use has afforded an opportunity to compare the two mechanization concepts in the same operating environment. The two systems, although having basically different computers, use similar inertial measurement units. Equation mechanization is different primarily because of the difference in computers. The development problems on the analog system were considerably more complex than those with the digital system, inasmuch as the analog it could be refined. These development problems ultimately brought about the redesign of analog system and utilization of the digital system. 91 980078 M2-F2 Lifting Body Three-View Drawing 538. Burcham, F. W., Jr.: Wind-Tunnel Calibration of a 40 Deg Conical Pressure Probe at Mach Numbers From 3.5 to 7.4. NASA TN D-4678, July 1968, 68N28801, #. A wind-tunnel calibration of a 40° included-angle flow-field cone probe was made over a Mach number range of 3.5 to 7.4. The cone probe was designed and fabricated by the NASA Flight Research Center to obtain flow-field data on the X-15 airplane. Estimated accuracy of the calibration was ±2 percent in Mach number and ±0.2 degree in flow angularity at a Mach number of 7.4. Reynolds number effects were negligible over the test range of 0.65 million to 3.25 million per foot (0.20 million to 1. 0 million per meter). A rake designed for flight on the X-15 was used to mount two cone probes. Slightly different calibrations resulted for the two cones because of differences in the cone afterbody configurations. 539. Carpenter, L. R.; Lewis, C. E., Jr.; and McDonald, R. T.: Electrocardiograph Transmitted by RF and Telephone Links in Emergency Situations. FRC-10031, July 1968, 68B10233. 540. Roman, J.; Larmie, F. M.; and Figarola, T. R.: A Simple Laboratory Method for Reduction of Rhythm and Rate in Large-Scale Monitoring of Electrocardiogram. NASA TN D-4751, August 1968, 68N32100, #. A laboratory system for rapid reduction of large amounts of continuously recorded ECG information has been developed. The system consists of a 60-times-real-time playback device which generates one pulse for each cardiac cycle, appropriate signal conditioning and logic circuitry, and a counting and printing system. Practical means for culling out noisy information, at 60 times real time, are provided. 541. McLain, L. J.; and Palitz, M.: Flow-Field Investigations on the X-15 Airplane and Model Up to Hypersonic Speeds. NASA TN D-4813, September 1968, 68N35173, #. Flight-measured impact pressures and local Mach numbers near the surface of the rear-lower-fuselage centerline, wing lower surface, and upper vertical tail of the X-15 airplane are presented and compared with calculated results and windtunnel data. In addition, wind-tunnel-derived total pressures in the rear-lower-fuselage flow field are presented. The flight measurements are presented over a free-stream Mach number range of l to 5.7 and an angle-of-attack range of 0 degree to 20 degrees. The wind-tunnel measurements cover a Mach number range of 4.0 to 8.0. 542. Ehernberger, L. J.: Atmospheric Conditions Associated With Turbulence Encountered by the XB-70 Airplane Above 40,000 Feet Altitude. NASA TN D-4768, September 1968, 68N33416, #. 92 High altitude atmospheric turbulence has been encountered by the XB-70 airplane during flight tests over the Western United States. The encounters from 36 flights were used to obtain a preliminary assessment of the meteorological features associated with high altitude turbulence. This study used data from an NACA VGH recorder carried on the airplane and from rawinsonde observations made near turbulence encounters at altitudes above 40,000 feet (12,200 meters). These data showed that turbulence of significant intensity at high altitudes is related to wind velocity, vertical and wind shear, and the vertical temperature gradient. These findings are in general agreement with various turbulence-generating disturbances suggested previously in the literature. It is also indicated that the disturbances causing high-altitude turbulence can originate in both the lower atmosphere and the stratosphere. 543. Saltzman, Edwin J.; Goecke, Sheryll A.; and Pembo, Chris: Base Pressure Measurements on the XB-70 Airplane at Mach Numbers From 0.4 to 3.0. NASA TM X-1612, September 1968, 71N17132, #. Full-scale flight base pressure coefficients obtained from the XB-70 propulsion package are compared with predicted values based on a combination of cold jet flow wind tunnel models and data from a two-engine side-by-side jet, full-scale aircraft. At cruise mach numbers the base pressures of the full-scale aircraft were higher than predicted, resulting in a favorable increment of about 2 percent in terms of lift-drag ratio. At low supersonic speeds near a Mach number of 1.2, the negative base pressure coefficients were about three times larger than predicted, which would result in a significant liftdrag-ratio decrement. The investigation showed that the net calculated effect of underestimating the base drag a low supersonic climbout speeds, even though overestimating the base drag at cruise, can seriously reduce the range potential of the aircraft, depending on several operational factor that can influence transonic excess thrust. The trend of this range decrement (with respect to the transonic excess thrust) emphasizes the need for a base drag prediction based on models with a higher degree of similitude throughout the transonic and supersonic range. 544. Taylor, Lawrence W.; and Iliff, Kenneth W.: A Modified Newton-Raphson Method for Determining Stability Derivatives From Flight Data. Second International Conference on Computing Methods in Optimization Problems, San Remo, Italy, September 9–13, 1968. This paper presents the formulation of least squares and the Newton-Raphson method. The results are compared and discussed. The work reported was done jointly by the authors and Dr. A. V. Balakrishnan of the University of California at Los Angeles. Example solutions are included that show not only the quality of the fitted solution but also a comparison of the estimated coefficients with values obtained by other techniques. 545. Kordes, Eldon E.: Status of Structural-Response and Modal-Suppression Programs on the XB-70. Langley Meeting on Aircraft Response to Turbulence, NASA Langley Research Center, Hampton, Virginia, September 24–25, 1968. The XB-70 flight program has provided an excellent opportunity to study the dynamic response of a large flexible aircraft under full-scale conditions up to a Mach number of 3 at 70,000 feet altitude The purpose of this paper is to describe some of the research being carried out on the XB-70 in the area of turbulence, turbulence encountered, and airplane response and to present some preliminary results obtained 50 far in the program. 546. Mallick, Donald L.; Kluever, Emil E.; and Matranga, Gene J.: Flight Results Obtained With a NonAerodynamic, Variable Stability, Flying Platform. NASA TM X-59039. Presented at 10th Symposium and Banquet, Los Angeles, California, September 23, 1968. 547. Wilson, R. J.; and McKay, J. M.: Landing Loads and Accelerations of the XB-70-1 Airplane. NASA TN D-4836, October 1968, 68N36073, #. Data are presented on landing-contact conditions for the first 48 landings of the XB-70-1 airplane. Landing weights varied from 419, 800 pounds (190,400 kilograms) to 274,600 pounds (124,600 kilograms), Vertical velocities at touchdown ranged from 5.26 feet/second (1.603 meters/ second) to 1.49 feet/second (0.454 meter/second). Maximum indicated airspeed was 195.0 knots, with a minimum of 167.3 knots. 548. Watts, Joe D.: Flight Experience With Shock Impingement and Interference Heating on the X-15-2 Research Airplane. NASA TM X-1669, October 1968, 92N70863, #. Severe structural melting damage due to complex shock impingement and interference effects on local aerodynamic heating was experienced on a flight of the X-15-2 research airplane to a maximum Mach number of 6.7. Measured flight temperature data and observed structural damage resulting from shock impingement and interference heating on the airplane and its ablative coating were analyzed in the light of hypersonic wind-tunnel results. The best approximations of the flight results were made by increasing the undisturbed pylon leading-edge heat-transfer coefficient by a factor of 9 and the undisturbed heat-transfer coefficient in the two interference zones by a factor of 7. The calculated effect of increased heat transfer due to interference in radiationequilibrium temperature is presented for selected hypersonic cruise conditions. 93 549. Fulton, F. L., Jr.: Lessons From the XB-70 as Applied to the Supersonic Transport. NASA TM X-56014. Presented at the 21st Annual International Air Safety Seminar, Anaheim, California, October 7–11, 1968. 1968, 68N35734, #. The lessons from the XB-70A program that have been selected for discussion are only a few of the things that have been learned during the program. These things will certainly apply to the supersonic transport (SST). In some cases they will apply to any large airplane, and in a few cases they will apply to almost any airplane. The XB-70 is a very valuable research airplane; there is no other airplane in the world of similar size that can fly in the same speed environment. Many of its design features were pushing the state of the art; therefore, both positive and negative results were obtained, providing validation or correlation of design prediction techniques. It also provided information on operational factors applicable to a large supersonic aircraft The program has been expensive in money, time, and personal sacrifice, but if the knowledge gained from the XB-70A test program makes it possible to avoid even one catastrophic SST accident, the program will more than pay for itself. 550. Richardson, R. B.; and Harney, P. F.: Flight and Laboratory Testing of a Double Sideband FM Telemetry System. NASA TM X-56015. Proceedings of the International Telemetering Conference, Los Angeles, California, October 8–11, 1968, 1968, pp. 581–596, 69A19132. Double sideband suppressed carrier FM telemetry system as airborne data recorder, discussing noise, environmental conditions, laboratory and flight tests. 551. Kordes, Eldon E.: XB-70 Contributions to Environmental Technology of the Supersonic Transport. Presented at the 1968 ASME Transportation Engineering Conference, Washington, D.C., October 27–30, 1968. 552. Wilson, E. J.: Use of Strain Gages for Measurements of Flight Loads in a High-Temperature Environment. Proceedings, Instrument Society Of America, 5th Annual Test Measurement Symposium, New York, New York, October 28–31, 1968, 1968, pp. 555 1 to 555 5, 69A31277. Strain gages to measure flight loads in high temperature environment, discussing selection, calibration techniques and performance characteristics. 553. Reed, R. D.: Can the R/C’er Contribute to Aeronautical Research? R/C Modeler, Vol. 5, No. 10, October 1968, pp. 28–35. 554. Taylor, Lawrence W., Jr.; Iliff, Kenneth W.; and Powers, Bruce G.: A Comparison of Newton-Raphson and Other Methods for Determining Stability Derivatives From Flight Data. Third Technical Workshop on Dynamic Stability Problems, Ames Research Center, Moffett Field, California, November 4–7, 1968. (See also 562.) This paper presents the formulation of least squares, Shinbrot, and Newton-Raphson methods as applied to the problem of determining stability derivatives from flight data. The results are compared and discussed. The work reported was done jointly by the authors and Dr. A. V. Balakrishnan of the University of California at Los Angeles. Example solutions are included that show not only the quality of the fitted solution but also a comparison of the estimated coefficients with values obtained by other methods. 555. McTigue, J. G.; and Ryan, B. M.: Lifting-Body Research Vehicles in a Low-Speed Flight Test Program. New York Academy of Sciences, International Congress on Subsonic Aeronautics, New York, New York, April 3–6, 1967, New York Academy of Sciences, Annals, Vol. 154, November 1968, pp. 1014–1032, 69A15571. The potential advantages offered by the lifting-body concept for entry and landing prompted the NASA Flight Research Center at Edwards, Calif., to initiate a multiphased flight study to determine the handling qualities and maneuvering required to flare and land this class of vehicles. This paper discusses the background of the program, the objectives, and the results obtained to date. 556. Kock, B. M.; and Painter, W. D.: Investigation of the Controllability of the M2-F2 Lifting-Body Launch From the B-52 Carrier Airplane. NASA TM X-1713, December 1968, 71N15004, #. The launch characteristics of the M2-F2 lifting body after release from the B-52 carrier airplane were studied by using analytical methods and simulators to predict launch safety and to determine the piloting requirements during launch. The predicted launch characteristics and the flight results are compared. 557. Jenkins, J. M.; Tang, M. H.; and Pearson, G. P. E.: Vertical-Tail Loads and Control-Surface Hinge-Moment Measurements on the M2-F2 Lifting Body During Initial Subsonic Flight Tests. NASA TM X-1712, December 1968, 71N15003, #. Subsonic aerodynamic load characteristics are presented for the right vertical tail and the control surfaces on the M2-F2 lifting-body vehicle. The effects of vehicle attitude and control-surface deflection on the vertical-tail loads are determined. Coefficients defining the effects of angle of attack, angle of sideslip, upper-flap deflection, and rudder deflection on flight-measured vertical-tail loads are presented in terms of linear equations. Portions of two maneuver time histories are included to illustrate the magnitude of each of these effects. The effects of angle of attack and controlsurface deflection on the flight-measured rudder, upper-flap, and lower flap hinge moments are discussed. The measured loads data are presented in aerodynamic-coefficient form. Large vertical-tail loads were measured during flight tests. Flight-measured control-surface hinge-moment data are compared with wind-tunnel data obtained from full-scale vehicle tests. 558. McLeod, N. J.; Lasagna, P. L.; and Putnam, T. W.: Predicted and Measured XB-70 Ground-To-Ground Engine Noise. NASA SP-189. Presented in program of NASA Research Relating to Noise Alleviation of Large Subsonic Jet Aircraft, 1968, pp. 423–434, 69N11569, #. Measurements have been made of XB-70 engine noise during ground runs. The effect of engine power settings and engine spacing on the noise spectra during the ground runs is presented. Some of the data obtained during the ground runs were analyzed to determine the amplitude variation, and the effect of averaging time was determined. The SAE method was used to predict the noise levels for various test conditions, and comparisons of predicted and measured spectra are presented. Tests indicate some limitations for the SAE prediction method. During ground operation of the XB-70 airplane, the amplitude variations in the acoustic data during quasi-stable engine thrust indicate that similar variations might occur when flyover noise spectra are being measured. Atmospheric attenuation predictions were greater than the measured attenuation for the high-frequency octave bands at the specific conditions of these tests. 1969 Technical Publications 559. Taylor, L. W., Jr.; Smith, H. J.; and Iliff, K. W.: Experience Using Balakrishnan’s Epsilon Technique to Compute Optimum Flight Profiles. AIAA Paper 69-75. Presented at the 7th AIAA Aerospace Sciences Meeting, New York, New York, January 20–22, 1969, 70A28088, 69A18058. (See also 589, 605.) A technique for computing optimum profiles is developed which differs from the classical gradient method in that a term representing the constraint of satisfying the equations of motion is included in the cost function to be minimized. Although the number of unknown independent functions is increased to include the state variables, the dimensionally of the gradient of the modified cost is greatly reduced, resulting in considerable savings in complexity and time. The unknown control and state variables are expressed in s functional expansion to facilitate solution by means of Newton’s method. The effects of weighting terms and the 94 number of functions on the convergence properties are discussed. Comparisons are made of solutions using the classical gradient method, dynamic programming, and Balakrishnan’s epsilon technique. 560. *Newell, F. D.; and Smith, H. J.: Human Transfer Characteristics in Flight and Ground Simulation for a Roll Tracking Task. NASA TN D-5007, February 1969, 69N17814. Pilot transfer characteristics for three pilots have been measured in flight and in ground-based simulators for a compensatory roll tracking task with small bank-angle disturbances. The forcing function, in each case, consisted of the sum-of-ten-sine-waves with a bandwidth of l5 radians per second. A variable-stability T-33 airplane was used to obtain the flight measurements. Ground-based simulator measurements were obtained with both the T-33 airplane and a general-purpose simulator which used a contactanalog color display. Three different controlled elements were used, two of which were simple single-degree-offreedom controlled elements that had been studied previously. The third was a multiple-degree-of-freedom element representative of an airplane with good handling qualities and was considered and controlled as a singledegree-of-freedom configuration in roll. *Cornell Aeronautical Laboratory, Inc., Buffalo, New York. 561. Montoya, E. J.; and Nugent, J.: Wind-Tunnel Force and Pressure Tests of Rocket-Engine Nozzle Extensions on the 0.0667-Scale X-15-2 Model at Supersonic and Hypersonic Speeds. NASA TM X-1759, March 1969, 69N20873, #. Wind-tunnel force and pressure test results of nozzle extensions on the 0.0667-scale X-15-2 model over the freestream Mach number range from 2. 3 to 8.0 at angles of attack from –5 degree to 18 degree and Reynolds numbers of 2.0 × 10 (to the 6th) per foot and 3.4 × 10 (to the 6th) per foot (1.12 × 107 per meter) are presented. The effects of the presence of an aft-mounted ramjet shape and control-surface deflections are shown. 562. Taylor, L. W., Jr.; Iliff, K. W.; and Powers, B. G.: A Comparison of Newton-Raphson and Other Methods for Determining Stability Derivatives From Flight Data. AIAA Paper 69-315, 3rd AIAA and FTSS Conference, Houston, Texas, March 10–12, 1969, 69A22379, #. (See also 554.) A new technique of determining stability derivatives from flight data is formulated and compared with the simple equations, analog matching, least squares, and Shinbrot 95 methods of analysis. It is shown that the new technique, termed Newton-Raphson, is superior to the others whether flight data or a statistical model is used. Although the new method uses the Newton-Raphson technique, it is also similar to quasilinearization. The Newton-Raphson technique has been developed to enable the use of a priori (wind tunnel) information and to automatically adjust bias terms and initial conditions to compensate for errors. The technique has been successfully applied to the X-15, XB-70, F-111, and HL-10 vehicles and has application to many other system identification problems. 563. Quinn, Robert D.; and Olinger, Frank V.: HeatTransfer Measurements Obtained on the X-15 Airplane Including Correlations With Wind-Tunnel Results. NASA TM X-1705, NAS 1.15:X-1705, March 1969, 92N70607. Heat transfer measurements were obtained on the X-15 airplane from two flights under quasi-steady conditions at a freestream Mach number of 5.1 and an angle of attack of 2.0 degrees, and a freestream Mach number of 4.98 and an angle of attack of 16.3 degrees. These measurements were made at corresponding freestream Reynolds numbers of 2.45 × 10 (exp 6) and 1.31 × 10 (exp 6) per foot. Experimental heat transfer coefficients derived from temperatures obtained from 200 recording thermocouples on the skin of the airplane are tabulated. Correlations with wind tunnel results show that the wind tunnel data are in fair to good agreement with the flight data obtained on the wing, ventral tail, and vertical tail at low angles of attack but are generally in poor agreement with high angle of attack data and the low angle of attack fuselage data. 564. Wagner, C. A.: Visual Simulation Image Generation Using a Flying-Spot Scanner. NASA TN D-5151, April 1969, 69N23194, #. This paper analyzes the flying-spot scanner television camera used as a video signal generator for visual flight simulation. A description of the technique is included as well as a detailed theoretical analysis of the quality of the video that can be produced. A commercially available flying-spot scanner designed for flight simulation was tested, and its capabilities are presented. Discussion is limited to the television camera; display devices such as monitors are not discussed. The flying-spot scanner was found to be a low-cost device which can simulate all motions of an aircraft except roll. Large excursions at high rates are easily achieved. It is limited to a relatively small field of view, a fixed forward visibility, and simulation of flat terrain. The quality of the video is adequate over a moderate range of altitudes; a very high and very low altitudes result in poor performance. Possible approaches to increasing the useful altitude range are discussed. 565. Mallick, D. L.; and Fulton, F. L., Jr.: Flight Crew Preparation and Training for the Operation of Large Supersonic Aircraft. FAUSST VII Meeting, Paris, France, March 3–7, 1969. 566. Bellman, D. R.; and Hughes, D. L.: The Flight Investigation of Pressure Phenomena in the Air Intake of an F-111A Airplane. AIAA Paper 69-488. Presented at the 5th AIAA Propulsion Joint Specialist Conference, U. S. Airforce Academy, Colorado Springs, Colorado, June 9–13, 1969, June 1969, 69A32706, #. Balakrishnan’s epsilon technique is used to compute minimum time profiles for the F-104 airplane. This technique differs from the classical gradient method in that a quadratic penalty on the error in satisfying the equations of motion is included in the cost function to be minimized as a means of eliminating the requirement of satisfying the equations of motion. Although the number of unknown independent functions is increased to include the state variables, the evaluation of the gradient of the modified cost is simplified, resulting in considerable computational savings. The unknown control and state variables are approximated by a functional expansion with unspecified coefficients which are determined by means of Newton’s method. Typically 8 to 10 iterations are required for convergence when using the epsilon technique. Comparisons are made of solutions obtained by using this technique and the energy method. ECN-2092 F-111A Aardvark Airplane 567. Painter, W. D.; and Kock, B. M.: Operational Experiences and Characteristics of the M2-F2 Lifting Body Flight Control System. NASA TM X-1809, June 1969, 71N14526, #. F-104 Airplane, Three-View Drawing Flights of the M2-F2 lifting body demonstrated that the manual control system and the stability augmentation system met the operational flight control requirements for the test vehicle. The regions of pilot-induced oscillation predicted from ground simulation were encountered in flight. The pilots considered the control system to be adequate for the M2-F2 flight envelope flown. Limit-cycle data obtained during ground tests agreed with flight results. Structural frequencies of the vehicle control surfaces were never sustained in flight as a result of filtering in the stability augmentation system. 568. Taylor, Lawrence W., Jr.; Smith, Harriet J.: and Iliff, Kenneth W.: A Comparison of Minimum Time Profiles for the F-104 Using Balakrishnan’s Epsilon Technique and the Energy Method. Presented by Dr. A. V. Balakrishnan at Symposium on Optimization, Nice, France, June 29–July 5, 1969. (See also 649.) 96 569. Lock, W. P.; and Gee, S. W.: Flight Investigation of a Fluidic Autopilot System. NASA TN D-5298, July 1969, 69N30946, #. A flight investigation was made of an experimental fluidic flight control system capable of various modes of operation, including altitude hold, heading hold, wings leveler and turn control. The fluidic control system was tested in each mode at two flight conditions: cruise at 5000 feet, and cruise at 10,000 feet. Although stability problems were encountered early in the program, stable performance was achieved in each control mode for the flight conditions tested. High reliability was demonstrated, in that there were no failures with fluidic elements themselves. Failures were experienced, however, with the mechanical portion of the mechanical fluidic components. 980064 570. Gaidsick, H. G.; Dana, W. H.; and McCracken, R. C.: Evaluation of an Indirect Viewing System for LiftingBody Terminal-Area Navigation and Landing Tasks. NASA TN D-5299, July 1969, 69N29730, #. A short-eyed-relief optical system, consisting of two monocular periscopes with overlapping fields of view, was mounted in an F-104B airplane to evaluate the feasibility of using this type of indirect viewing system in place of normal vision for performing simulated lifting-body approaches and landings. Three approach techniques were used in the study. Performance was evaluated by measuring touchdown distance from a marked touchdown point and rate of sink and airspeed at touchdown. Results obtained with the optics system were compared with normal-vision results. The ability of the pilots to perform the simulated lifting-body tasks was not noticeably reduced with the optics system. The workload and other pilot acceptance factors, however, indicated that this particular system required improvement in design, even though the pilots could readily adapt to its use. 571. Swaroop, R.; West, K. A.; and Lewis, C. E., Jr.: A Simple Technique for Automatic Computer Editing of Biodata. NASA TN D-5275, July 1969, 69N29592, #. Before any data are statistically analyzed, it is always necessary to edit the data to some extent. Furthermore, when large quantities of data are collected, the editing must performed by automatic means. One common task in the editing process is the identification of observations which deviate markedly from the rest of the sample, commonly known outliers. A simple statistical technique for identifying the outliers and the necessary computer program is presented in this report. The program requires as input only the data set, sample size, and preselected levels of significance which outliers are to be identified. It is assumed that the data set is a random sample of size larger than two from a normal population. Two examples are presented to illustrate applications of the described technique. 572. Layton, Garrison P., Jr.: Interim Results of the Lifting-Body Flight-Test Program. NASA TM X-1827. Presented at AIAA Entry Vehicle Systems and Technology Meeting, Hampton, Virginia, December 3–5, 1968. July 1969, 71N14527, #. The significant results of the joint NASA/U. S. Air Force lifting-body flight-test program are presented in general terms, based on 16 M2-F2 glide flights, 14 HL-10 flights and wind-tunnel tests of the X-24A flight vehicle. The liftingbody flight-test program has demonstrated that lifting reentry vehicles can be maneuvered to an unpowered landing from initial conditions representing the entry of the terminal area for a reentry vehicle. 980548 HL-10 Lifting Body, Three-View Drawing 573. Borek, R. W.; and Richardson, R. B.: Flight and Laboratory Tests of an “L-Band” Telemetry RF System. Telemetry Journal, Vol. 4, July 1969, pp. 21–24, 69A35996. Laboratory and flight tests of airborne solid state UHF telemetry transmitter, discussing miniaturized coaxial hardware from RF power conservation viewpoint. 574. Taylor, L. W., Jr.; and Iliff, K. W.: Fixed-Base Simulator Pilot Rating Surveys for Predicting LateralDirectional Handling Qualities and Pilot Rating Variability. NASA TN D-5358, August 1969, 69N35762, #. Pilot ratings of lateral-directional handling qualities were collected for a wide range of simplified aircraft characteristics through the use of a simple fixed-base simulator with a color contact analog display. The results of the general survey were obtained with an engineer as the subject and are contained in 45 plots involving five parameters. The survey results show that the handling qualities for the specific simulations used were, in general, optimum. The results of the general survey are used in an empirical method for predicting lateral-directional pilot ratings for most airplane configurations and flight conditions. In another survey, utilizing the same simulation, ratings were 97 obtained from many pilots in order to study the variability in pilot ratings among pilots and the differences in pilot rating that result from changes in mission. The standard deviation of individual pilot ratings ranged from 1. 0 at the “good” end of the scale to 2. 0 in the middle of the scale. Ratings for specific missions were generally numerically higher (more adverse) than those for the general mission for the same vehicle characteristic. 575. Jenkins, J. M.; DeAngelis, V. M.; Friend, E. L.; and Monaghan, R. C.: Flight Measurements of Canard Loads, Canard Buffeting, and Elevon and Wing Tip Hinge Moments on the XB-70 Aircraft Including Comparisons With Predictions. NASA TN D-5359, August 1969, 69N32284, #. During a flight-test program with the XB-70 airplane, canard, elevon, and wingtip flight load measurements were made in the Mach number range from 0.40 to 3.00. The data are compared with the manufacturer’s rigid and aeroelastic predictions, wind-tunnel airfoil data, or results obtained from flight tests of other applicable aircraft. The magnitudes of the flight loads and the variation of surface loads with angle of attack or surface deflection corresponded generally with predictions. Canard buffeting was experienced at subsonic speeds. The characteristics of this effect are examined on the basis of the results of in-flight tuft studies and the analysis of flight-measured bending-moment data. 576. Kotfilia, R. P.; and Painter, W. D.: Design, Development, and Flight Test Experience With Lifting Body Stability Augmentation Systems. AIAA Paper 69-887. Presented at the AIAA Guidance, Control and Flight Mechanics Conference, Princeton, New Jersey, August 18–20, 1969, 69A39413, #. The aerodynamic characteristics of lifting body research vehicles tested thus far require stability augmentation systems (SAS) to improve the handling qualities of the vehicles in some areas of the flight envelope. The design goals for such a system are performance, reliability, and ease of testing. Validation of these goals requires realistic test criteria that establish vehicle-SAS performance based on ground tests. This paper describes the design of the SAS for the X-24A vehicle and the ground test techniques that were used for this validation. Test criteria and procedures are established for various ground tests, such as frequency response, limit cycle, and structural resonance. Ground test results are compared with subsequent flight test data obtained in the joint NASA-USAF lifting body flight research program. 980045 X-24A Lifting Body, Three-View Drawing 577. Rediess, H. A.; and *Whitaker, H. P.: A New Model Performance Index for the Engineering Design of Flight Control Systems. AIAA Paper 69-885. Presented at the AIAA Guidance, Control, and Flight Mechanics Conference, Princeton, Jew Jersey, August 18–20, 1969, 69A39410, #. The theory and application of a new performance index, the Model PI, that brings engineering design specifications into the analytical design process is presented. A parameter optimization design procedure is established that starts with practical engineering specifications and uses the Model PI as a synthesis tool to obtain a satisfactory design. The Model P1 represents a new criterion for approximating one dynamical system by another, based on a novel geometrical representation of linear autonomous systems. It is shown to be an effective performance index in designing practical systems and to be substantially more efficient to use than a comparable model-referenced integral squared error performance index. The design procedure is demonstrated by designing a lateral-directional stability augmentation system for the X-15 aircraft. *Massachusetts Massachusetts. Institute of Technology, Cambridge, 578. Beaulieu, W.; Campbell, R.; and Burcham, W.: Measurement of XB-70 Propulsion Performance Incorporating the Gas Generator Method. J. Aircraft, Vol. 6, No. 4, July–August 1969, 9A37152. (See also 536.) Propulsion performance of XB-70A aircraft calculated by gas generator method. 98 579. Sefic, W. J.; and Anderson, K. F.: NASA High Temperature Loads Calibration Laboratory. NASA TM X-1868, September 1969, 69N36224, #. The NASA High Temperature Loads Calibration Laboratory and the equipment it contains for simulating the loading and heating of aircraft or their components are described. Particular emphasis is placed on various fail-safe devices which are built into the equipment to minimize the possibility of damage to flight vehicles. The data-acquisition system is described. This system involves on-site pickup of data and signal conditioning coupled with conversion from analog to digital data and transmission to a central area for recording on magnetic tape and distribution to real-time displays. Instrumentation available in the facility for measuring load, position, and strain is also discussed. 580. Fields, R. A.; and Vano, A.: Evaluation of an Infrared Heating Simulation of a Mach 4.63 Flight on an X-15 Horizontal Stabilizer. NASA TN D-5403, September 1969, 69N35949, #. Temperatures recorded on the X-15 horizontal stabilizer during a Mach 4.63 flight were simulated in the laboratory. A liquid-nitrogen evaporative cooler was used to cool the structure to a prelaunch condition; the heating was provided by an infrared heating system with closed-loop control. The simulated flight produced temperatures from approximately – 50 degrees F (228 degrees K) to 750 degrees F (672 degrees K). The simulation was evaluated by comparing flight-measured temperatures with those measured during the simulation. 581. Lewis, C. E., Jr.; and Krier, G. E.: Flight Research Program: XIV—Landing Performance in Jet Aircraft After the Loss of Binocular Vision. Aerospace Medicine, Vol. 40, September 1969, pp. 957–963, 69A41675. Landing performance in T-33A aircraft with loss of binocular vision is compared to performance with both eyes. 582. *Lipana, J. G.; *Fletcher, J.; *Brown, W.; and *Cohen, G.: Effects of Various Respiratory Maneuvers on the Physiological Response to Angular Acceleration. Aerospace Medicine, Vol. 40, September 1969, pp. 976–980, 69A41679. The effects of breath holding, Ml, Valsalva and Mueller’s maneuvers were studied on healthy males during static condition at various postures and during pure axis rotations. The subject was seated inside a hollow spherical simulator (ARTS). Rotation was at the rate of 6 rpm with the axis of rotation through the body. Heart rates, EGG, blood pressures, respiratory rates, voice and TV were monitored via telemetry. 583. Adkins, E. J.; McLeod, N. J.; and Lasagna, P. L.: Variation in Engine Noise for Two Landing-Approach Configurations of a Jet Transport Aircraft. NASA TM X-1896, October 1969, 69N38104, #. A limited flight investigation was conducted to determine the effect of reduced flap deflections on power required and the resulting engine noise for a subsonic jet transport aircraft. Noise levels were measured during level flight at an altitude of 400 feet (122 meters) at approach speeds. Data were obtained during flybys with flap deflections of 50 degrees and 36 degree. The maximum overall sound pressure level (OASPL) from flybys with 36 degree of flap deflection averaged 3 decibels (ref. 0.00002 newtons/meter2) lower than for flybys with 50 degree of flap deflection. Buffet intensity was reduced by the use of lower-flap deflections. 584. Fisher, D. F.: Flight-Measured Aerodynamic Drag of Two Large External Tanks Attached to the X-15-2 Airplane at Mach Numbers of 1.6 to 2.3. NASA TM X-1895, October 1969, 69N38065, #. Full-scale power-on flight lift and drag measurements were made on the X-15-2 airplane just before and just after two large external fuel tanks were ejected. By subtracting the drag of the airplane after the tank ejection from the drag of the airplane before tank ejection, the incremental drag due to the tanks was determined. Analysis of the data showed that the percentage increase of incremental drag due to the tanks was almost equal to the percentage increase in cross-sectional area caused by the tanks. A buildup drag estimate based on free-stream conditions agreed well with the flight data; whereas, the wind tunnel data, although they had the same general trend of the tank drag coefficient with Mach number, were lower than both the estimate and the flight data. E-17242 *Systems Research Laboratories, Inc., City, State. 99 XB-70 and X-15-2 on Ramp 585. Irwin, K. S.; and Andrews, W. H.: Summary of XB-70 Airplane Cockpit Environmental Data. NASA TN D-5449, October 1969, 69N38010, #. Thermal, acoustical, and acceleration environmental data were obtained for the crew compartment of the XB-70 airplane during the 186-flight-hour airworthiness test program. More than 20 hours were flown at Mach numbers greater than 2.5. Temperature levels, gradients, and time histories are presented for the cockpit walls, floor, and windshields. Heat transfer through the walls and along the floor produced no crew discomfort. Thermal radiation from the hot inner windshield would have been objectionable to the crew if they had not been protected by insulated flight suits and helmets with faceplates. The acoustical environment of the crew compartment was similar to that of other military turbojet bomber aircraft. At Mach 3 the sound-pressure level in the XB-70 cockpit, primarily generated by onboard electrical and environmental equipment, was 90 decibels, which is about 10 decibels higher than that measured on a present subsonic jet transport. Subsonically, the cockpit noise levels exceeded military specification limits by as much as 10 decibels at frequencies above 400 hertz. At cruise conditions the cockpit noise levels exceeded the supersonic transport internal-noise-level design criteria in the frequency range above 300 hertz. Acceleration data for the XB-70 crew compartment and center of gravity are presented for taxi, takeoff, subsonic buffet, and atmospheric-turbulence conditions. The long, flexible fuselage produced unpleasant ride characteristics in the crew compartment under vibrational situations. However, flight control was always maintained, even in heavy turbulence. 586. McTigue, J. G.; and Layton, G. P., Jr.: Lifting Body Flight Tests and Analysis. SAE Paper 69-0662. Presented at the Society of Automotive Engineers, National Aeronautic and Space Engineering and Manufacturing Meeting, Los Angeles, California, October 6–10, 1969, 70A15840. Reusable lifting entry vehicle flight tests, investigating handling qualities and subsonic-transonic aerodynamics of M2-F2, M2-F3, HL-10 and X-24A. 587. Tang, M. H.; and DeAngelis, V. M.: Fin Loads and Control-Surface Hinge Moments Measured in Full-Scale Wind-Tunnel Tests on the X-24A Flight Vehicle. NASA TM X-1922, H-580, November 1969, 71N14501, #. Tests were conducted on the full-scale X-24A lifting-body in the 40- by 80-Foot Wind Tunnel at the NASA Ames Research Center. One purpose of the tests was to measure aerodynamic loads on the stabilizing fins and hinge moments on all the control surfaces. The tests were conducted at dynamic pressures of 60, 80, and 100 lb/ft2 (2870, 3830, and 4790 N/m2). The effects of variations in rudder deflections, flap deflection, and angles of attack and sideslip were studied. Also, limited tests were performed with a simulated ablative surface on the aerodynamic characteristics. Detailed results of the wind-tunnel tests are given in the form of load coefficients and hinge-moment coefficients. The results are compared with data from tests performed in other wind tunnels on small-scale models. 588. Quinn, Robert D.; and Olinger, Frank V.: FlightMeasured Heat Transfer and Skin Friction at a Mach Number of 5.25 and at Low Wall Temperatures. NASA TM X-1921, H-579, NAS 1.15:X-1921, November 1969, 92N70606, #. Turbulent skin friction and heat transfer coefficients were measured simultaneously on a test panel installed on the sharp leading edge upper vertical tail of the X-15 airplane at wall to recovery temperature ratios of 0.218 to 0.333 and at a nominal free stream Reynolds number of 1.54 x 10(exp 6) per foot. The data were obtained from one flight at a nominal free stream Mach number of 5.25. Reynolds analogy factors were derived from skin friction and heat transfer measurements. The measured data are compared with values predicted by various theories. 589. Taylor, Lawrence W., Jr.; Smith, Harriet J.; and Iliff, Kenneth W.: Experiences Using Balakrishnan’s Epsilon Technique to Compute Optimum Flight Profiles. Fourth NASA Inter-Center Control Systems Conference, Boston, Massachusetts, November 4–5, 1969. (See also 559, 605.) A technique for computing optimum profiles is developed which differs from the classical gradient method in that a term representing the constraint of satisfying the equations of motion is included in the cost function to be minimized. Although the number of unknown independent functions is increased to include the state variables, the dimensionally of the gradient of the modified cost is greatly reduced, resulting in considerable savings in complexity and time. The unknown control and state variables are expressed in s functional expansion to facilitate solution by means of Newton’s method. The effects of weighting terms and the number of functions on the convergence properties are discussed. Comparisons are made of solutions using the 100 ECN-2353 X-24A, M2-F3, and HL-10 Lifting Bodies classical gradient method, dynamic programming, and Balakrishnan’s epsilon technique. 590. Deets, Dwain A.: Optimal Regulator of Conventional Setup Techniques for a Model Following Simulator Control System. Fourth NASA Inter-Center Control Systems Conference, Boston, Massachusetts, November 4–5, 1969. (See also 969.) This paper compares the optimal regulator technique for determining simulator control system gains with the conventional servo analysis approach. Practical considerations associated with airborne motion simulation using a model-following system provided the basis for comparison. The simulation fidelity specifications selected were important in evaluating the relative advantages of the two methods. Frequency responses for a JetStar aircraft following a roll mode model were calculated digitally to illustrate the various cases. A technique for generating forward loop lead in the optimal regulator model-following problem was developed which increases the flexibility of that approach. In this study it appeared to be the only way in which the optimal regulator method could meet the fidelity specifications. 591. Rediess, Herman A.: Linear Optimal Control Via a Model Performance Index. Fourth NASA Inter-Center Control Systems Conference, Boston, Massachusetts, November 4–5, 1969. 592. Iliff, K. W.; and Taylor, L. W., Jr.: A Modified Newton-Raphson Method for Determining Stability Derivatives From Flight Data. Computing Methods in Optimization Problems—2, Proceedings of the Second International Conference, San Remo, Italy, September 9–13, 1968, 1969, pp. 353-364, 70A19272. Approach. NASA TN D-5676, H-587, February 1970, 70N19804, #. Approaches and landings during the XB-70 program were performed at various approach speeds, glide-slope angles, gross weights, runway offsets, and operational conditions. Representative time histories, pilot comments, and pilot ratings were obtained from these maneuvers. Stability and control data and limited correlations with predictions and handling-qualities criteria were also obtained. The XB-70 flight experience indicated that the height of the cockpit above the runway in combination with nose-high landing attitudes and high approach speeds made the landing task more difficult than that for current subsonic jet transports. Three-degree glide slopes were considered unsatisfactory at the 200-knot indicated airspeed approaches required by the XB-70 The high rate of descent reduced the time available to accomplish the flare and, therefore, increased the possibility of a hard landing. Large changes in lift due to elevon deflection were satisfactory because of the high control effectiveness. Laterally, the aircraft was sensitive to turbulence. Lateral-offset maneuvers simulating breakout from an overcast were not difficult; however, because of the higher approach speeds, excessive runway distances would be covered prior to touchdown and the adverse yaw accompanying aileron deflection was considered excessive. Sidestep maneuvering performance was adequately predicted by a simple technique. 596. Burcham, F. W., Jr.; and Nugent, J.: Local Flow Field Around a Pylon-Mounted Dummy Ramjet Engine on the X-15-2 Airplane for Mach Numbers From 2.0 to 6.7. NASA TN D-5638, H-566, February 1970, 70N18035, #. The flow field around a pylon-mounted dummy ramjet engine on the X-15-2 airplane was surveyed at Mach numbers from 2.0 to 6.7 in preparation for flight tests of a hydrogenburning hypersonic ramjet engine. Impact pressures, local Mach number, and flow angularity were determined and compared with wind-tunnel data and theoretical calculations. The wing, camera fairing, and side fairing of the X-15-2 generated shock waves which impinged on the dummy ramjet and pylon. However, a region free of significant shock-wave impingement on the ramjet inlet existed for flight at a freestream angle of attack of 5 degree or less for free-stream Mach numbers from 3. 0 to 8.0. In flight regions free of shock-wave impingement, impact pressure, local Mach number, and angle of attack generally showed good agreement with wind-tunnel data. Shock-wave locations determined from impact-pressure data and wind-tunnel schlieren photograph data showed good agreement. Strong flow-interference effects occurred at the pylon-fuselage intersection. The separated-flow region and the resulting separation shock wave remained within 10 inches (25.4 centimeters) of the fuselage surface in front of the pylon for all flight conditions. The extent of separated flow was 101 1970 Technical Publications 593. Thompson, M. O.; and Welsh, J. R.: Flight Test Experience With Adaptive Control Systems. AGARDCP-58, Paper 11. Advanced Control System Concepts. January 1970, pp. 139–147, 70N23037, #. 594. Taylor, L. W., Jr.: Nonlinear Time-Domain Models of Human Controllers. Journal of Optimization Theory and Applications, Vol. 5, January 1970, pp. 23–38, 70A26396. Human controllers nonlinear time domain mathematical model for analyzing compensatory tracking task data. 595. Berry, D. T.; and Powers, B. G.: Handling Qualities of the XB-70 Airplane in the Landing sensitive to angle of attack and extremely sensitive to small deviations from 0 degree in angle of sideslip. 599. Wykes, J. H.; and Kordes, E. E.: Analytical Design and Flight Tests of a Modal Suppression System on the XB-70 Airplane: Part 1 - Design Analysis, Part 2 - Flight Tests. AGARD-CP-46, Paper 22. Aeroelastic Effects From a Flight Mech. Standpoint, March 1970, (see N70-29401 15-02), 70N29423, #. A control system designed to damp the structural motion of flexible airframes was flight tested on the XB-70 airplane, a flexible, low-aspect-ratio supersonic configuration. Even though the system—known as ILAF (Identical Location of Accelerometer and Force)—was an exploratory device and was not developed as an optimum system, the flight tests provided valuable information applicable to aircraft of the supersonic-transport type. This paper reviews the design processes and presents some preliminary results obtained from flight data. 600. Szalai, K. J.; and Deets, D. A.: An Airborne Simulator Program to Determine If Roll-Mode Simulation Should Be a Moving Experience. AIAA Paper 70-351. Presented at the AIAA Visual and Motion Simulation Technology Conference, Cape Canaveral, Florida, March 16–18, 1970, 70A24202, #. Motion and visual cue influences on pilot performance and opinion in roll motion tasks were investigated in an airborne simulator flight program. Human describing functions were obtained from a compensatory tracking task for three roll damping conditions (low, medium, and high) for each of three simulator configurations (fixed base, IFR; moving base, IFR; and moving base, VFR). In addition, the pilots made handling qualities evaluations of the nine different cases. The gross effects of motion and visual stimuli, as determined by pilot comments and ratings, were compared with the conclusions drawn from compensatory tracking task results. The quantitative results from the tracking task experiment showed definite motion cue influences. However, the handling qualities experiment indicated that motion was not necessary to obtain valid results. It was concluded that the quantitative results from tracking task are not adequate to establish the need for motion in a handling qualities research simulator. 601. Sisk, Thomas R.; Enevoldson, Einar K.; and Krier, Gary E.: Factors Affecting Tracking Precision. Presented at AIAA Fighter Airplane Conference, St. Louis, Missouri, March 5–7, 1970. A tracking study conducted at the Flight Research Center on three aircraft has demonstrated that the degradation in tracking precision due solely to buffet intensity is on the order of 5 to 6 mils for buffet intensity levels to 0.2, and it is believed that these trends are generally applicable to other fighter aircraft. Also, the results of this study indicate that only a portion of the buffet-free normal-force coefficient attained with maneuver flaps on the F-104 aircraft may be 102 EC88-180-2 X-15 Airplane With Dummy Ramjet 597. Gee, S. W.; Kock, B. M.; and Schofield, B. L.: Operational Experiences With Unpowered Terminal Area Instrument Approaches. Proceedings, Inst. of Navigation, National Space Meeting on Space Navigation Theory and Practice in the Post Apollo Era, NASA Ames Research Center, Moffett Field, California, February 17–19, 1970, 1970, pp. 211–223, 70A30465, #. 598. Putnam, T. W.; and Smith, R. H.: XB-70 Compressor-Noise Reduction and Propulsion-System Performance for Choked Inlet Flow. NASA TN D-5692, H-578, March 1970, 70N20484, #. An investigation was conducted with the XB-70 airplane attached to a thrust stand to observe compressor-radiated noise and propulsion-system performance as the inlet throat area was reduced to form an aerodynamically choked flow. The anticipated compressor-noise reduction was not experienced. Tests at constant engine speed settings disclosed only minor noise reductions as a result of choking the inlet. Additional tests were performed with the inlet fixed full open and the engines set at successively higher speeds from 60-percent to 57-percent rpm. These tests disclosed considerable noise reduction (10 decibels) in the compressorblade passing frequencies at engine settings above 80-percent rpm. Thus, it was concluded that most of the noise reduction expected for the throat-closing tests, which were done at engine settings of 87-percent rpm and higher, had already occurred and that only slight additional suppression could he expected when the throat area was decreased to choke the flow. Choking the flow resulted in thrust losses due to decreased total-pressure recovery and airflow in the inlet. These test results suggest that major noise reduction may be obtained without a full-choked flow and the concurrent propulsion-system performance loss. utilized if tracking precision is to be maintained because of the introduction of wing rock. Finally, it has been demonstrated that certain stability and control or handlingqualities factors can degrade tracking precision to an equal or greater degree than buffet intensity. 602. Evans, Robert D., Jr.: Development and Testing of a Triaxial Angular Accelerometer for High-Performance Aerospace Vehicles. Presented at the 6th International Aerospace Instrumentation Symposium, Cranfield, England, March 23–26, 1970. 603. Fields, R. A.: A Study of the Accuracy of a FlightHeating Simulation and Its Effect on Load Measurement. NASA TN D-5741, H-597, April 1970, 70N24320, #. A series of laboratory heating tests simulating the flight heating on an X-15 horizontal stabilizer was conducted. The initial test simulated, as nearly as reasonably possible, the temperatures that were recorded during an X-15 flight to a Mach number of 4.63. Ten additional heating tests were conducted during which inaccuracies were introduced into the flight-heating simulation. The objective of these tests was to establish the effect of the inaccuracies on the strain-gage responses and ensuing load measurements. Strain-gagebridge responses from all tests were reviewed and compared with those calculated for a stabilizer load of 6000 pounds force (26,700 newtons). The tests were shown to be useful for selecting bridges for use in load-equation derivations and for selecting the equations that yield load measurements with the lowest overall error. 604. Andrews, W. H.; Robinson, G. H.; Krier, G. E.; and Drinkwater, F. J., III: Flight-Test Evaluation of the Wing Vortex Wake Generated by Large Jet-Transport Aircraft. FAA Comp. of Work Papers Concerning Wake Turbulence Tests, (see N70-40911 23-02), April 1970, 70N40912, #. 605. Taylor, L. W., Jr.; Smith, H. J.; and Iliff, K. W.: Experience Using Balakrishnan’s Epsilon Technique to Compute Optimum Flight Profiles. AIAA Paper 69-75. Presented at the 7th AIAA Aerospace Sciences Meeting, New York, New York, January 20–22, 1969. Journal of Aircraft, Vol. 7, No. 2, March–April 1970, (see A69-18058), April 1970, pp. 182–187, 70A28088, #. (See also 559, 589.) A technique for computing optimum profiles is developed which differs from the classical gradient method in that a term representing the constraint of satisfying the equations of motion is included in the cost function to be minimized. Although the number of unknown independent functions is increased to include the state variables, the dimensionally of the gradient of the modified cost is greatly reduced, resulting in considerable savings in complexity and time. The unknown control and state variables are expressed in a functional expansion to facilitate solution by means of 103 Newton’s method. The effects of weighting terms and the number of functions on the convergence properties are discussed. Comparisons are made of solutions using the classical gradient method, dynamic programming, and Balakrishnan’s epsilon technique. 606. Berry, D. T.; and Powers, B. G.: Flying Qualities of a Large, Supersonic Aircraft in Cruise and Landing Approach. AIAA Paper 70-566. Presented at the AIAA Atmospheric Flight Mechanics Conference, Tullahoma, Tennessee, May 13–15, 1970, 70A29031, #. XB-70 handling-qualities flight-test experience is reviewed and its implications for handling qualities criteria are analyzed. Pilot ratings, pilot comments, and flightdetermined handling qualities parameters are presented. Results throughout the entire flight envelope are considered, but emphasis is placed on high-speed supersonic cruise and landing approach. 607. Fischel, Jack; and Friend, Edward L.: Preliminary Assessment of Effects of Wing Flaps on High Subsonic Flight Buffet Characteristics on Three Airplanes. Presented at the American Institute of Aeronautics and Astronautics, Atmospheric Flight Mechanics Conference, Tullahoma, Tennessee, May 13–15, 1970. 608. Lane, James W.; and Evans, Robert D.: A High Altitude Altimeter Utilizing a Vibrating Diaphragm Transducer. Presented at the ISA 16th National Aeronautics Institute Symposium, Seattle, Washington, May 11–13, 1970. 609. Saltzman, Edwin J.; and Fisher, David F.: Some Turbulent Boundary-Layer Measurements Obtained From the Forebody of an Airplane at Mach Numbers Up to 1.72. NASA TN D-5838, H-567, June 1970, 70N29908, #. Boundary-layer-profile data were obtained from the smooth undersurface of the A-5A airplane fuselage during the demonstration of sensors for measuring boundary-layer characteristics. The data represent Mach numbers from 0.51 to 1.72. angles of attack up to 7°, and Reynolds numbers up to 74 million. The data are interpreted in terms of local skin friction and momentum thickness, and the velocity profiles from which these data are derived are tabulated. Local transformed friction coefficients obtained from a Clauser type of determination from velocity profiles were close to the incompressible values of Karman-Schoenherr when presented as a function of momentum thickness Reynolds number. Turbulent momentum thickness values were significantly influenced by angle of attack. The flight values of momentum thickness for angles of attack near 6° and 7° were lower than the flat plate values, approaching the level for slender cones. At angles of attack near 0° and 1°, momentum thickness from flight was higher than flat plate values. The aircraft nose boom and the protuberances on the boom are believed to be major reasons for the additional thickness at low angles of attack. simulation program was used to determine the reaction of the started inlet over a range of performance levels to free-stream temperature gradients in smooth air and in clear-air turbulence. Atmospheric-temperature variations were found to be nearly as significant as severe turbulence for operation of that type of inlet. Unstart margins for the manually controlled XB-70 inlet in the presence of these disturbances were obtained thorough use of the simulation program. 612. Martin, R. A.: Dynamic Analysis of XB-70-1 Inlet Pressure Fluctuations During Takeoff and Prior to a Compressor Stall at Mach 2.5. NASA TN D-5826, H-595, June 1970, 70N28273, #. Instrumentation in the left inlet of the XB-70-1 airplane was used to record high-response total- and static-pressure data from 0 to 200 hertz during takeoff and immediately prior to a compressor stall at Mach 2.5 and an altitude of 63,100 feet (19,200 meters). Since the statistical assumptions of stationarity, randomness, and normality were found to be approximately valid for the inlet pressure data, random dataanalysis techniques were applied. Values of mean turbulence parameter as high as 12 percent were obtained during takeoff and from 14 percent to as high as 31 percent prior to stall. The flight inlet turbulence-producing mechanism, namely, normal-shock boundary-layer interaction, can he simulated successfully in ground test facilities up to at least 40 hertz as evident in the pressure-wave power spectra; however, higher turbulence values were experienced in flight than in model tests by an engine compressor prior to stall. 613. Rediess, Herman A.: Is Modern Control Theory Relevant to Flight Control Systems? — A Design Challenge. Presented at the Joint Automatic Control Conference, Atlanta, Georgia, June 22–23, 1970. 614. Burcham, F. W., Jr.; and Hughes, D. L.: Analysis of In-Flight Pressure Fluctuations Leading to Engine Compressor Surge in an F-111A Airplane for Mach Numbers to 2.17. AIAA Paper 70-624. Presented at the 6th AIAA Propulsion Joint Specialist Conference, San Diego, California, June 15–19, 1970, 70A33543, #. 615. Jarvis, C. R.; Loschke, P. C.; and Enevoldson, E. K.: Evaluation of the Effect of a Yaw-Rate Damper on the Flying Qualities of a Light Twin-Engine Airplane. NASA TN D-5890, H-584, July 1970, 70N32770, #. A flight-test program was conducted with a light twin-engine airplane to determine the effect of a parallel yaw damper and aileron-to-rudder interconnect on the flying qualities of this class of aircraft. Both quantitative and qualitative results are presented for several flight tasks and conditions, including flight in turbulence. Airplane handling qualities and ride qualities are summarized. The effect of the yaw damper on the stall and post-stall motions of the test airplane and the motions resulting from sudden engine failure are also discussed. 104 ECN-231 A-5A Vigilante Airplane 610. Kier, D. A.: Flight Comparison of Several Techniques for Determining the Minimum Flying Speed for a Large, Subsonic Jet Transport. NASA TN D-5806, H-590, June 1970, 70N28674, #. A flight investigation was conducted to define the minimum flying speed for a large, subsonic jet transport by using three techniques: (1) the Federal Aviation Regulations (FAR) Part 25 demonstration technique; (2) a flight-path 1-g-break technique; and (3) a constant-rate-of-climb technique. The effect of thrust on minimum speed is analyzed. Results indicate that the flight-path 1-g-break technique was the best overall technique. The constant-rate-of-climb technique, or minimum level-flight speed, though highly affected by the deceleration dynamics of the maneuver, was found to be an acceptable alternate for the 1-g-break technique. The FAR demonstration technique, when analyzed by two current analysis methods, was found to yield the least conservative results. However, if the analysis were based on actual airplane maximum lift capability, the technique would yield acceptable results. 611. Gallagher, R. J.: Investigation of a Digital Simulation of the XB-70 Inlet and Its Application to Flight-Experienced Free-Stream Disturbances at Mach Numbers of 2.4 to 2.6. NASA TN D-5827, H-585, June 1970, 70N28343, #. The capability of a digital inlet simulation program to predict both the performance of the started XB-70 inlet system at free-stream Mach numbers of 2.4 to 2.6 and the dynamic aspects of an inlet unstart at these Mach numbers was analyzed. The simulation-predicted performance was compared with flight measurements, and reasonable agreement was obtained for started and unstarting inlet modes. The agreement between empty-fill (buzz) simulations and flight data was less precise. This was attributed to the need for additional information on the boundary-layerseparation process during this mode of inlet operation. The 616. Kordes, E. E.: Secondary Structures and Mechanisms — Design Trouble Area for the Space Shuttle. TM X-52876, Space Transportation System Technology Symposium, Vol. 3, July 1970, pp. 93–99, 70N42983, #. The design of secondary structures and mechanisms for a reusable space shuttle can produce problems as serious as those involved in designing the primary structure and thermal protection system. Several events involving failure of secondary structures during the X-15 flight program have been selected to illustrate potential problem areas. Similar problems may be expected on future reusable vehicles until additional research provided adequate design information in these areas. 617. Capasso, V. N., Jr.: Space Shuttle Related Maintenance Experience With the X-15 Aircraft. NASA TM X-52876. Proceedings, Space Transportation System Technology Symposium, Vol. 5, July 1970, pp. 33–44, 70N39605, #. This paper discusses the maintenance activity between X-15 flights and the number and types of repair items which would be related to space shuttle requirements. The increased size and complexity of the shuttle systems will magnify the number of repair items, making the required turnaround time difficult or impossible to achieve unless careful consideration is given to problem prevention and access for system repair and maintainability. 618. Schofield, B. L.; Gaidsick, H. G.; and Gee, S. W.: Experience With Unpowered Terminal-Area Instrument Approaches. NASA TM X-52876. Proceedings, Space Transportation System Technology Symposium, Vol. 6, July 1970, (see N70-40951 23-31), pp. 133–147, 70N40960. The first part of this paper will discuss a terminal-area guidance technique recently developed by the Air Force Flight Test Center around the F-111A inertial navigation system. The results of flying under instrument flight rules using this guidance scheme will be reported as will the results of ground controlled approaches (GCA) using an NB-52B airplane in an unpowered, low L/D configuration. The latter portion of this paper will discuss a circular approach guidance scheme under development at the NASA Flight Research Center. 619. Thompson, M. O.: Lifting-Body Progress Report. NASA TM X-66712. Presented at the ELDO/NASA Space Transportation Systems Briefing, Bonn, July 7–8, 1970, 71N18428, #. 620. Montoya, L. C.: Drag Characteristics Obtained From Several Configurations of the Modified X-15-2 Airplane Up to Mach 6.7. NASA TM X-2056, H-598, August 1970, 70N35693, #. Flight tests were made with and without the lower ventral fin, dummy ramjet (including the modified fixed ventral fin), and 105 ablative coating over the entire wetted area. Data were obtained at Mach numbers from 0. 5 to 6.7 and free-stream Reynolds numbers from approximately 1.7 × 10 (to the 8th) to 3.6 × 10 (to the 7th) based on fuselage length. Angle of attack ranged from 0 degree to 11 degrees, dynamic pressure from about 300 lb/ft2 (14,364 N/m2) to 770 lb/ft2 (36,868 N/ m square), and altitude from approximately 3000 ft (914 m) to 102.000 ft (31,090 m). Supersonic flight results showed an increase in drag coefficient caused by the ablative coating of 0.008 at a lift coefficient of 0 and 0.022 at a lift coefficient of 0.3. At subsonic speeds the average increase in drag coefficient was about 0.013 for lift coefficients of 0.3 and 0.4. The flight incremental increase in drag coefficient due to the ablative coating showed good agreement with compressible predicted values at low lift coefficients. The average incremental increase in drag coefficient caused by the lower ventral fin and dummy ramjet was about 0.010 at subsonic speeds at a lift coefficient of about 0.3. The flight incremental increase in drag coefficient of the lower ventral fin between Mach 3.8 and 4.9 was about 0.006 at a lift coefficient of 0.1 and 0.014 at a lift coefficient of 0.3. 980044 X-15-2 Airplane, Three-View Drawing 621. Holleman, E. C.: Flight Investigation of the Roll Requirements for Transport Airplanes in Cruising Flight. NASA TN D-5957, H-616, September 1970, 70N38625, #. An airborne simulator provided a wide range of maximum roll control power (0.05 to 3.5 rad/sec square) and time constants (0.1 to 10 sec) for pilot evaluation and rating. Roll criteria were developed and compared favorably with previously reported criteria. Maximum roll angular acceleration, maximum roll rate, roll time constant, time to bank. and bank-angle change in a given time all appear to be effective roll-criteria parameters. Steady-state roll rates of about 20 deg/sec and roll time constants of 1.8 seconds or less were required for satisfactory pilot ratings. With experienced test pilots, valid evaluation of single-degree-of-freedom roll response ran be obtained with a fixed-base simulator. 622. Andrews, W. H.: Flight-Test Evaluation of the Wing Vortex Wake Generated by Large Jet-Transport Aircraft. Presented at the Wake Turbulence Meeting, Seattle, Washington, September 1–3, 1970. (See Aircraft Wake Turbulence and Its Detection, Plenum Press, 1971.) 623. Lewis, C. E., Jr.; and Rezek, T. W.: A Miniature Respiratory Minute Volume Sensor for the Flight Environment. Space Life Sciences, Vol. 2, September 1970, pp. 206–218, 70A44840. 624. Carpenter, Richard: Description of an Energy Absorbing Seat Designed for Medium-Velocity Impacts. Presented at Annual Safety and Risk Management Conference, Lewis Research Center, September 30– October 1, 1970. 625. Baker, P. A.; Schweikhard, W. G.; and Young, W. R.: Flight Evaluation of Ground Effect on Several Low-Aspect-Ratio Airplanes. NASA TN D-6053, H-550, October 1970, 70N42738, #. A constant-angle-of-attack-approach technique was used to measure ground effect on several low-aspect-ratio aircraft. The flight results were compared with results from constantaltitude flybys, wind-tunnel studies, and theoretical prediction data. It was found that the constant-angle of-attack technique provided data that were consistent with data obtained from constant-altitude flybys and required fewer runs to obtain the same amount of data. The test results from an F5D-1 airplane modified with an ogee wing, a prototype F5D-1 airplane, two XB-70 airplanes, and an F-104A airplane indicate that theory and wind-tunnel results adequately predict the trends caused by ground effect as a function of height and aspect ratio. However, the magnitude of these predictions did not always agree with the flightmeasured results. In addition, there was consistent evidence that the aircraft encountered ground effect at a height above one wing span. 626. Wagner, C. A.: Frequency Responses and Other Characteristics of Six Fast-Decay Phosphors Applicable to Flying-Spot Scanners. NASA TN D-6036, H-609, October 1970, 70N42118, #. Several tests were conducted on six fast-decay phosphors to measure their characteristics applicable to flying-spot scanners. P-16, P-24, P-36, P-37, and P-SP phosphors are currently available in production cathode-ray tubes; P-X42 is an experimental phosphor. A pulse test measured rise and decay times, and two scanning tests measured frequency responses. Screen noise, burn and aging resistance, and variation of light output as a function of electron beam dwell time were also measured. The frequency responses of P-16, P-37, and P-SP phosphors did not vary during the tests. The frequency response of P-36 phosphor varied with electron beam dwell time, and the frequency response of P-24 106 phosphor varied with both dwell time and beam current. The bandwidths (frequencies at which the gain is 0.1) were as follows: P-16, 22 megahertz; P-24, 12 megahertz; P-36, 12 megahertz; P-37, 34 megahertz; P-SP, 21 megahertz. The bandwidth of P-X42 phosphor, too high to measure with the equipment used, was estimated to be in excess of 100 megahertz. However, this phosphor produced relatively little light and had a screen that was too noisy to be useful in flying-spot scanners. 627. Anon.: Flight Test Results Pertaining to the Space Shuttlecraft Symposium Papers. NASA TM X-2101, October 1970, 71N10101, #. This compilation consists of papers presented at the NASA Symposium on Flight Test Results Pertaining to the Space Shuttlecraft, held at the NASA Flight Research Center, Edwards, Calif, on June 30, 1970. The symposium was divided into the following sessions: Lifting Body Flight Test Results and Additional Space Shuttle Oriented Studies. Papers were presented by representatives from the NASA Flight Research Center and the U. S. Air Force Flight Test Center. A list of attendees is included. 628. McTigue, J. G.: Background and Current Status of the Lifting Body Program. NASA TM X-2101, October 1970, pp. 1–10, 71N10102, #. The purpose of this paper is to present the results from the flight-test program and the correlations of these data with predictions. Evaluations and impressions of the lifting body vehicles’ overall handling qualities will also be presented by several of the pilots in the program. Finally, future plans for the lifting body flight program will be discussed. 629. Kempel, R. W.; Strutz, L. W.; and *Kirsten, P.: Stability and Control Derivatives of the Lifting Body Vehicles. NASA TM X-2101, October 1970, pp. 11–27, 71N10103, #. In this paper the more important longitudinal and lateraldirectional aerodynamic stability and control derivatives obtained from flight are compared with small- and full-scale wind-tunnel results where applicable. Significant trends and important differences are pointed out, and the implications discussed. *Air Force Flight Test Center, Edwards, California. 630. Manke, J. A.; *Retelle, J. P.; and Kempel, R. W.: Assessment of Lifting Body Vehicle Handling Qualities. NASA TM X-2101, October 1970, pp. 29–41, 71N10104, #. (See also 659.) Handling qualities have always been vitally important to the pilot. Before the current series of lifting body flight tests, there was speculation and concern about how this class of wingless vehicle would handle. The general behavior of the three lifting bodies in flight is described in broad terms in this paper, and some specific examples of behavior that may be of special interest from the pilot’s viewpoint are presented. In addition, comments are offered concerning simulation requirements. *Air Force Flight Test Center, Edwards, California. 631. Pyle, J. S.; and *Ash, L. G.: Performance Characteristics of the Lifting Body Vehicle. NASA TM X-2101, October 1970, pp. 43–58, 71N10105, #. Flight and wind-tunnel lift and drag data obtained on three lifting body vehicles has been compared. With the exception of the drag-due-to-lift factor, the flight and small-scale windtunnel results generally agree; however, it is extremely important that the model contours match the flight vehicle and the flow on the model be carefully observed, because minor separation problems on the model may become severe on the full-scale vehicle in flight. The full-scale wind-tunnel results obtained with the flight vehicle generally predicted higher zero-lift drag coefficients and lower drag-due-to-lift factors than were observed during the flight tests. *Air Force Flight Test Center, Edwards, California. 632. Tang, M. H.: Correlation of Flight-Test Loads With Wind-Tunnel Predicted Loads on Three Lifting Body Vehicles. NASA TM X-2101, October 1970, pp. 59-72, 71N10106, #. An essential area of research with the unique M2-F2, HL-10, and X-24A lifting body configurations is the assessment of the ability to predict flight loads from wind-tunnel tests. Flight measurements and correlation with predictions are necessary in verifying the structural integrity of existing vehicles and establishing the groundwork for weight savings on future vehicles of similar shapes. As part of the overall lifting body flight investigation at the Flight Research Center, detailed aerodynamic-load studies are being made on each of the three vehicles. This paper presents the preliminary results from these studies. 633. Dana, W. H.; and *Gentry, J. R.: Pilot Impressions of Lifting Body Vehicles. NASA TM X-2101, October 1970, pp. 73–88, 71N10107, #. Piloting aspects of the lifting body vehicles are discussed in this paper by two of the pilots assigned to the flight program. Subjects discussed include: approach, landing, and energy management considerations; field of view requirements; stability considerations; and vehicle riding qualities, including the effects of turbulence. Remarks pertinent to the various subject areas are made by each pilot. *Air Force Flight Test Center, Edwards, California. 634. Layton, G. P., Jr.: Summary of Primary Results of the Lifting Body Program. NASA TM X-2101, October 1970, pp. 89–97, 71N10108, #. 107 This summary paper will point out results of the lifting body program that have a bearing on the design of a large space shuttle vehicle. The initial program objectives, the primary program results, and the pertinence of these results to the shuttle will be outlined, as will the future direction of the program. 635. Kock, B. M.; and Fulton, F. L., Jr.: Approach and Landing Studies. NASA TM X-2101, October 1970, pp. 99–108, 71N10109, #. This paper will discuss the application of recent approach and landing studies to the proposed space shuttle. These studies were conducted basically in two areas: powered approaches with the HL-10 lifting body, and unpowered types of approaches with shuttle-size vehicles, the B-52 and CV-990 airplanes. 636. Holleman, E. C.: Rationale for Proposed FlyingQualities Specifications. NASA TM X-2101, October 1970, pp. 127–145, 71N10111, #. The NASA Flight Research Center is reviewing the applicability of flying-qualities experience to the shuttle mission for the purpose of preparing a shuttle flying-qualities specification. This paper is a progress report on this review; the result presented are, of course, preliminary. 637. Thompson, M. O.: Flight Test Results Pertaining to the Space Shuttlecraft — Final Remarks and Future Plans. NASA TM X-2101, October 1970, pp. 147–151, 71N10112, #. This paper sums up the flight test results pertaining to the space shuttlecraft symposium held at the Flight Research Center. Additional remarks by the author discuss future plans to continue work on unpowered approach and landing techniques and the feasibility of a lifting entry. 638. Carpenter, Lewis R.: Biotechnology Problems Relative to the Space Shuttle Vehicle. Presented at AIAA 7th Annual Meeting and Technology Display, Houston, Texas, October 19–22, 1970. 639. Egger, R. L.; and Wilson, E. J.: Design and Operation of a 1500 F Thermal-Null Strain Gage. Instrument Society of America, Annual Conference, 25th, Philadelphia, Pennsylvania, Oct. 26–29, 1970, Proceedings: Part 2, 1970, pp. 631.1–631.6, 71A22721. High temperature thermal null strain gage with sensing unit and electronic control unit to measure mechanical strain in terms of induced thermal strain. 640. Roman, J.; Lewis, C. E., Jr.; and Allen, W. H.: Hazards of the G-Suit in Lower Extremity Thrombophlebitis. Aerospace Medicine, Vol. 41, October 1970, pp. 1198–1199, 70A45347. Trauma is widely accepted as an etiologic factor in venous thrombosis and thrombophlebitis of the lower extremities. Because of the frequent participation of military pilots and test pilots in athletic activities, the incidence of venous thrombosis of the extremities may be expected to be significant in this population. This group is likely to fly high performance vehicles and, therefore, likely to use the g-suit. On theoretical grounds, use of the g-suit in the face of recent venous thrombosis in the lower extremities should be hazardous. This problem is considered in this paper. 641. Thompson, M. O.: A Progress Report on the Lifting Body Flight Test Program. Proceedings, Space Technology and Earth Problems, American Astronautical Society, Symposium, Las Cruces, New Mexico, October 23– 25, 1969, 1970, pp. 23–27, 71A14819. 642. Wilson, E. J.: Installation and Testing of Strain Gages for High-Temperature Aircraft Applications. Presented at the Society for Experimental Stress Analysis, Fall Meeting, Boston, Massachusetts, October 18–22, 1970, 71A13781, #. 643. Burke, M. E.: Flight Research Experience With Guidance and Control Computers Related to General Applications. NASA TM X-66491, November 1970, 71N12610, #. Several guidance and control research programs involving the X-15 and F-104 airplanes are discussed, with the discussion oriented toward airborne digital computer utilization. An analog and a digital systems mechanization are compared, and the performance advantages of the digital system are pointed out. The flexibility of the digital computer as a research tool is indicated, as are advantages of decentralized computers. Application of a general purpose computer to the solution of strapdown system equations was successful in the laboratory in preparation for a flight program. The effects of input-output mechanizations on software complexity are discussed. The utility of a general purpose digital computer is shown by its flexibility in being used for various research tasks. This utility is degraded, however, by the effort required to write programs in machine language for real-time applications. 644. Arnaiz, H. H.; and Schweikhard, W. G.: Validation of the Gas Generator Method of Calculating Jet-Engine Thrust and Evaluation of XB-70-1 Airplane Engine Performance at Ground Static Conditions. NASA TN D-7028, H-596, December 1970, 71N13419, #. Deficiencies in established techniques of measuring aircraft thrust in flight led to the application of the gas generator method of calculating engine thrust to the XB-70-1 airplane. A series of tests on a ground static-thrust stand was performed on the airplane to establish at ground static conditions the accuracy of this method, to measure the installed thrust of the YJ93-GE-3 engine, and to determine the effect of instrumentation errors and nonuniform flows at the engine compressor face on the thrust calculation. Tests 108 with an aerodynamically choked inlet, an opened inlet-bypass system, and varying combinations of operating engines were also conducted. Results showed that the accuracy of the gas generator method was ±2 percent for the normal operation of the XB-70-1 airplane at ground static conditions and for the upper 70 percent of the engine’s throttle range. They also showed that the effect of individual instrument errors on the thrust calculation was reduced because of the large number of measurements and that abnormally high inlet flow distortion affects the thrust calculation. When corrected for inlet losses, the installed thrust of the YJ93-GE-3 engine agreed favorably with the engine manufacturer’s uninstalled estimated thrust for all power settings except those at the low end. 645. Rediess, H. A.; and Whitaker, H. P.: A New Model Performance Index for Engineering Design of Flight Control Systems. AIAA Paper 69-885, presented at AIAA Guidance, Control, and Flight Mechanics Conference, Princeton, New Jersey, August 18–20, 1969, 71A12682, #. (Also J. Aircraft, Vol. 7, No. 21, December 1970, pp. 542–549, A69-39410.) Model performance index /Pi/ providing criterion for approximating one dynamic flight control system by another based on geometrical representation of linear autonomous systems. 646. *Allison, R. D.; Lewis, C. E.; and Rezek, T. W.: Vascular Dynamics — Impedance Plethysmograph Study During a Standardized Tilt Table Procedure. Space Life Sciences, Vol. 2, December 1970, pp. 361–393, 71A17958. A study, using the four-electrode impedance plethysmograph system, was completed to evaluate simultaneous variations in conduction of upper and lower body segments relative to displacement of blood volume during change in body position. Measurements of cardiac output were compared with simultaneous results by dye dilution methods as a means of ascending the use of impedance technique to determine cardiac output during tilt table studies. Two groups, 48 healthy private pilots and 22 patients with diabetes mellitus, were tested and the results were compared. *Scott and White Clinic, Temple, Texas. 647. Dana, W. H.; and Gentry, G.: Flying the Lifting Bodies. Flight International, Vol. 98, December 31, 1970, pp. 1016–1020, 71A16680. 648. Taylor, L. W., Jr.: Nonlinear Time-Domain Models of Human Controllers. NASA SP-215, 1970, pp. 49–65, 70N30880. This paper presents some of the results of subsequent analyses and discusses the method of selecting the maximum memory time and the order of the non-linear model. In addition, some results of orthogonal expansion of the weighting functions for reasons of data compression and reduction computation are presented and discussed. 649. Taylor, L. W., Jr.; Smith, H. J.; and Iliff, K. W.: A Comparison of Minimum Time Profiles for the F-104 Using Balakhrishnan’s Epsilon Technique and the Energy Method. Proceedings, International Federation For Information Processing, Symposium On Optimization, Nice, France, June 29–July 5, 1969, 1970, pp. 327–335, 71A28831. (See also 568.) Balakrishnan’s epsilon technique is used to compute minimum time profiles for the F-104 airplane. This technique differs from the classical gradient method in that a quadratic penalty on the error in satisfying the equations of motion is included in the coat function to be minimized as a means of eliminating the requirement of satisfying the equation of motion. Although the number of unknown independent functions is increased to include the state variables, the evaluation of the gradient of the modified cost is simplified, resulting in considerable computational savings. The unknown control and state variables are approximated by 3 functional expansion with unspecified coefficients which are determined by means of Newton’s Method. Typically 8 to 10 iterations are required for convergence when using the epsilon technique. Comparisons are made of solutions obtained by using this technique and the energy method. increased or the number of adjacent engines operating was increased. As the distance between two operating engines became greater, the overall sound pressure level increased as the angle between the microphone position and the exhaust axis decreased. The overall sound pressure level agreed best with the SAE prediction levels at an angle of 120 degree for exhaust velocities between 1500 feet/second (457 meters/ second) and 3000 feet/second (914 meters/second). The SAE method adequately estimated the noise spectrum of the XB-70 airplane for subsonic exhaust flow and underestimated the high-frequency spectral levels for supersonic flow. Some shielding of the high frequencies was observed when two or more adjacent engines were operating in supersonic exhaust flow conditions. The noise spectrum shape was independent of jet exhaust velocity for the XB-70 engines with supersonic flow. 651. Carpenter, R.; Thompson, M. O.; and *Bowers, J. H.: Instrumentation and Drop-Testing Techniques for Investigating Flight Vehicles and Personnel Protective Systems. NASA TM X-2149, H-621, January 1971, 71N14780, #. A vehicle and flight systems dynamic impact laboratory, with appropriate instrumentation, was constructed recently at the NASA Flight Research Center. The purpose of the laboratory is to investigate energy-absorbing characteristics of various flight vehicle configurations and personnel protection and restraint systems during low-energy impacting. The laboratory was designed to permit investigation of a large range of horizontal and vertical impact velocities that would occur in low-level crash situations. It has provided acquisition of dynamic structural data on both vehicle configurations and personnel restraint systems. *Northrop Corporation, Field Teams at Flight Research Center, Edwards, California. 652. Rediess, Herman A.; Mallick, Donald L.; and Berry, Donald T.: Recent Flight Test Results on Minimum Longitudinal Handling Qualities for Transport Aircraft. Presented at FAUSST VIII Meeting, Washington, D.C., January 1971. 653. Pecoraro, Joseph N.; and Carpenter, Lewis R.: Shuttle: Life Support, Protective Systems, and Crew Systems Interface Technology. Astronautics and Aeronautics, Vol. 9, February 1971, pp. 58–63. 654. Pyle, J. S.: Lift and Drag Characteristics of the HL-10 Lifting Body During Subsonic Gliding Flight. NASA TN D-6263, March 1971, 71N18867, #. Subsonic lift and drag data obtained during the HL-10 lifting body glide flight program are presented for four configurations for angles of attack from 5.0 to 26.0 and Mach numbers from 0.35 to 0.62. These flight data, where applicable, are compared with results from small-scale windtunnel tests of an HL-10 model, full-scale wind-tunnel 109 E-6213 F-104 Airplane 1971 Technical Publications 650. Lasagna, P. L.; and Putnam, T. W.: Engine Exhaust Noise During Ground Operation of the XB-70 Airplane. NASA TN D-7043, H-599, January 1971, 71N15820, #. XB-70 engine noise was measured from 90 to 1600 from the airplane heading at a radius of 500 feet (152 meters). Overall sound pressure levels, perceived noise levels, and normalized spectra are presented for jet exhaust velocities up to 3300 feet/second (1006 meters/second), various engine spacings, and various numbers of adjacent engines operating. The direction of propagation of maximum noise levels moved from 135 degree to 120 degree as either the jet velocity was results obtained with the flight vehicle, and flight results for the M2–F2 lifting body. The lift and drag characteristics obtained from the HL-10 flight results showed that a severe flow problem existed on the upper surface of the vehicle during the first flight test. This problem was corrected by modifying the leading edges of the tip fins. The vehicle attained lift-drag ratios as high as 4.0 during the landing flare (performed with the landing gear up), which is approximately 14 percent higher than demonstrated by the M2-F2 vehicle in similar maneuvers. attitude. Additional work is required to include rotation variables in standardizing test data and in more fully defining the rotation effects on performance from a limited number of takeoff tests. 656. Staff of the Flight Research Center: Experience With the X-15 Adaptive Flight Control System. NASA TN D-6208, H-618, March 1971, 71N18422, #. The X-15 adaptive flight control system is briefly described, and system development and flight-test experiences in the X-15 research airplane are discussed. Airplane handling qualities with the system and system reliability are also discussed. 657. McLeod, N. J.: Acoustic Attenuation Determined Experimentally During Engine Ground Tests of the XB-70 Airplane and Comparison With Predictions. NASA TM X-2223, H-633, March 1971, 71N17951, #. Acoustic data obtained during ground runs of the XB-70 airplane enabled the attenuation of sound to be determined for ground-to-ground propagation over distances of 500 feet (152 meters) and 1000 feet (305 meters) for one set of atmospheric conditions. The considerable scatter in the experimentally determined attenuations seemed to be associated with variations in the wind parameters. For downwind propagation, reducing the predicted acousticattenuation values obtained from the Society of Automotive Engineers, Inc., APP 866 by 50 percent for octave bands above 1000 hertz resulted in good agreement between the present experimentally determined and predicted values. Previous investigators recommended the modification to the predicted attenuation values and obtained similar results. Some unexplained differences remain between the experimentally determined and modified predicted attenuation values, indicating that additional research is required. 658. Gee, S. W.; Gaidsick, H. G.; and Enevoldson, E. K.: Flight Evaluation of Angle of Attack as a Control Parameter in General-Aviation Aircraft. NASA TN D-6210, H-603, March 1971, 71N18442, #. The use of angle-of-attack information for a pilot’s display in a general-aviation airplane was investigated to determine whether this form of information would improve performance and flight safety. An angle-of-attack system consisting of a wing-mounted vane, an electronic computer unit, and a display instrument was installed and flight tested in a typical twin-engine, general-aviation airplane. The flight-test maneuvers were limited to the low-speed flight region where the benefits of angle-of-attack presentation were likely to be greatest. Some of the expected advantages of this parameter, such as visual indication of stall margin and its independence of gross weight and flap position, were realized; however, certain aerodynamic characteristics of the airplane, such as the phugoid and directional-control 110 ECN-2203 HL-10 Lifting Body With B-52 Flyover 655. Larson, T. J.; and Schweikhard, W. G.: Verification of Takeoff Performance Predictions for the XB-70 Airplane. NASA TM X-2215, H-574, March 1971, 71N18509, #. XB-70 airplane standardized takeoff data are compared with simple predictions based on aerodynamic and engine estimates. Effects of atmospheric and aircraft variables on takeoff distance are evaluated. Although experimentation with various techniques for aircraft rotation to lift-off attitudes was limited, the effect of the pilot techniques used are discussed and compared. Predictions of distance from brake release to initiation of rotation as 5 functions of velocity were found to he accurate to approximately 100 feet (30 meters). Because of the significant drag at the high aircraft attitudes required for takeoff, the standardized ground roll distance for a given velocity was increased nominally by 400 feet (120 meters) over the distance which would occur with no increase in drag. Standardized performance during climb from lift-off to a height of 35 feet (10.7 meters) with all engines operating was marginal because of low longitudinal accelerations, resulting from high induced drag at lift-off capability, were found to limit and tended to negate some of the expected advantages. As a result. this use of angle of attack did not show a significant improvement in performance and flight safety. 659. Manke, J. A.; Kempel, R. W.; and *Retelle, J. P.: Assessment of Lifting Body Vehicle Handling Qualities. AIAA Paper 71-310, AIAA Space Shuttle Development Testing and Operations Conference, Phoenix, Arizona, March 15–17, 1971, 71A22622, #. (See also 630.) Before the current series of lifting body flight tests, there was speculation and concern about how this class of wingless vehicle would handle. The general behavior of the three lifting bodies in flight is described in broad terms in this paper, and some specific examples of behavior that may be of special interest from the pilot’s viewpoint are presented. *USAF, Flight Test Center, Edwards AFB, California. 660. Burcham, F. W., Jr.: An Investigation of Two Variations of the Gas Generator Method to Calculate the Thrust of the Afterburning Turbofan Engines Installed in an F-111A Airplane. NASA TN D-6297, H-643, April 1971, 71N22614, #. The NASA Flight Research Center investigated two variations of the gas generator method for calculating the net thrust of the afterburning turbofan engines installed in an F-111A airplane. An influence coefficient study and two ground thrust tests were performed. It was found that the gas generator method can be successfully applied to an afterburning turbofan engine. At static conditions with two engines operating, ±2 percent accuracy can be achieved for most power settings using either the method based primarily on nozzle total pressure and area (PTA) or the method based primarily on nozzle total temperature and weight flow (TTW). For in-flight conditions the influence coefficient results indicated that the accuracy of the TTW method was about ±3 percent, whereas the accuracy of the PTA method was about ±5 percent for a military power setting. With either calculation method, additional errors in calculated thrust of ±2 percent could result from high inlet flow distortion. If accurate thrust values are required, both thrust calculation methods should be used. 661. Holleman, E. C.; and Gilyard, G. B.: In-Flight Evaluation of the Lateral Handling of a Four-Engine Jet Transport During Approach and Landing. NASA TN D-6339, H-642, May 1971, 71N25537, #. As part of a program to document the stability, control, and flying qualities of jet transport airplanes, the lateral handling of a typical jet transport was evaluated during up-and-away and approach flight in the landing configuration. Sidestep maneuvers to a landing were performed with several levels of lateral control power in smooth-air conditions. A roll control power capability of about 15 deg/sec2 was required for 111 satisfactory lateral control, but 61-meter (200-foot) lateral offsets to the runway could be safely corrected with very low levels of lateral control power, approximately 2 to 5 deg/sec2, using altered piloting techniques. The pilot evaluation results were in general agreement with results from other studies. 662. Tang, M. H.: Vertical-Fin Loads and Rudder Hinge-Moment Measurements on a 1/8 Scale Model of the M2-F3 Lifting Body Vehicle at Mach Numbers From 0.50 to 1.30. NASA TM X-2286, H-650, May 1971, 71N24581, #. Outboard-fin loads and rudder hinge-moment measurements were obtained from a 1/8-scale model of the M2-F3 lifting body vehicle tested in the Ames Research Center’s 11-Foot Transonic Wind Tunnel. The tests were conducted at Mach 0.50 to 1.30. The effects of variations in rudder deflection, upper flap deflection, lower-flap deflection, and angles of attack and sideslip were studied. The left-outboard-fin loads increased with increase in angle of attack, Mach number, rudder deflection, lower-flap deflection, and negative sideslip and decreased with increasing upper-flap deflection. The rudder hinge moment increased with increase in rudder deflection and, generally, with Mach number. E-21533 M2-F3 Lifting Body 663. Goecke, S. A.: Comparison of Wind Tunnel and Flight-Measured Base Pressures From the SharpLeading-Edge Upper Vertical Fin of the X-15 Airplane for Turbulent Flow at Mach Numbers From 1.5 to 5.0. NASA TN D-6348, H-602, May 1971, 71N23922, #. Pressures measured at six locations on the base of the sharpleading edge upper vertical fin of the X-15 airplane during the power-off portion of eight flights are compared with previous flight data obtained from a blunt-leading-edge fin, theory, and wind-tunnel data. The flight and wind-tunnel base pressure ratios for the Mach number range from 1.5 to 5.0 are presented as a linearized function of turbulent boundary-layer height and base width by using a Mach-number-dependent factor derived in the study. The resulting curve seems to provide another criterion for determining whether flow is laminar or turbulent. The difference between base pressure and free-stream pressure for any specific Mach number of the study is found to be a linear function of both free-stream pressure and dynamic pressure. Data from the sharp-leadingedge upper vertical fin agree with data from the bluntleading-edge upper vertical fin. The flight data show the variation in pressure across the base to be negligible. 664. Ehernberger, L. J.; and Wilson, R. J.: Analysis of Subjective Ratings for the XB-70 Airplane Response to Atmospheric Turbulence and Controlled Inputs. Proceedings, Royal Aeronautical Society, International Conference on Atmospheric Turbulence, London, England, May 18–21, 1971, 1971, 71A29774, #. This paper presents the XB-70 crew’s subjective evaluation of the turbulence response in comparison with measured accelerations; evaluates the turbulence response of the crew station and other fuselage locations; and examines the crew’s reaction to controlled sinusoidal excitations of the airframe during flight. During the XB-70 flight tests, turbulence data were obtained as turbulence was encountered during scheduled flights; no flights were scheduled solely to obtain turbulence information. 665. Carpenter, R.: Evaluation of an Energy Absorbing Crew Seat Integrated With a Rocket Extraction System. Space Shuttle Technology Conference, Vol. 2, May 3, 1971, pp. 19–34, 71N35268, #. Consideration has been given to equipping the scaled prototype shuttle vehicle with a lightweight energy absorbing seat integrated with a crew extraction rocket. Such a system would provide protection for low velocity vehicle impacts and also offer a means of escape during higher velocity conditions. This system has been developed and fabricated at the Flight Research Center (FRC). The energy absorbing seat has been tested in a dynamic impact laboratory will satisfactory results. The escape system has been evaluated by extracting dummies by tractor rockets from a typical cockpit configuration. These tests indicate unsatisfactory performance during high roll rates. 666. Wilson, E. J.: Installation and Testing of Strain Gages for High-Temperature Aircraft Applications. Proceedings, Strain Gages and Extreme Environments, Society for Experimental Stress Analysis, Technical Session, Huntsville, Alabama, May 19, 1970, May 1971, pp. 1–10, 71A30681, #. Survey of a research program conducted to define the optimal selection, installation, and calibration criteria for strain gages used in aircraft flight-load measurements in a hightemperature environment caused by aerodynamic heating. Tests have been made to determine apparent strain, hysteresis, gage factor, insulation resistance, and attachment 112 methods. The necessity of an evaluation program to select the optimum strain gage for specific applications, strain-gageinstallation procedures, fatigue problems associated with weldable gages, and new strain-measurement concepts are discussed. Data from laboratory tests to determine the performance characteristics of bonded and weldable strain gages for high-temperature applications are presented. 667. Marshall, R. T.: Flight Determined Acceleration and Climb Performance of an F-104G Airplane for Use in an Optimum Flight Path Computer Program. NASA TN D-6398, H-636, June 1971, 71N27002, #. A flight-test investigation was conducted to determine the standard-day performance characteristics (excess thrust, fuel flow, and climb potential) at maximum afterburner power for an F-104G airplane. The tests were conducted at Mach numbers from 0.5 to 2.0 and at altitudes from 5000 feet (1524 meters) to 50,000 feet l5,240 meters). The standardday excess thrust and fuel-flow data obtained from the investigation were used to define a computer model of the performance of the test airplane. In addition, the climbpotential (specific excess power) data obtained from the flight tests were compared with the available predicted climbpotential data. From the comparisons, it was found that the predicted data for the “average” F-104G airplane did not represent the performance of the test airplane as accurately as required for the computation of meaningful flight trajectories. Therefore, to compute meaningful flight trajectories for the test airplane, the flight-derived model should be used. 668. Burcham, F. W., Jr.: Use of the Gas Generator Method to Calculate the Thrust of an Afterburning Turbofan Engine. AIAA Paper 71-680. Presented at the 7th AIAA and Society of Automotive Engineers, Propulsion Joint Specialist Conference, Salt Lake City, Utah, June 14–18, 1971, 71A30744, #. The feasibility of using the gas generator method to calculate the thrust of an afterburning turbofan engine was investigated. The NASA Flight Research Center’s F-111A airplane, powered by TF30 afterburning turbofan engines, was used. Two variations of the gas generator method were utilized, one based primarily on the exhaust nozzle total pressure and area and the other on the nozzle total temperature and weight flow. An influence coefficient study was performed for static and flight conditions. Results showed that the accuracy of the two calculation methods was about equal at static conditions, but for flight conditions the total temperature and weight flow calculation was superior. Two ground thrust stand tests were also performed, and the thrust calculated by using both methods was compared to the measured thrust; results using either method were generally within ±3 percent except at low power settings. High inlet flow distortion caused an additional scatter of as much as ±2 percent. 669. Wilson, R. J.; Love, B. J.; and Larson, R. R.: Evaluation of Effects of High-Altitude Turbulence Encounters on the XB-70 Airplane. NASA TN D-6457, H-631, July 1971, 71N30718, #. A turbulence response investigation was conducted with the XB-70 airplane. No special turbulence penetration techniques, speeds, or other restrictions were specified for the investigation, nor were any flights made solely to obtain turbulence data. During 79 flights, turbulence was encountered, and recorded on a VGH recorder, 6.2 percent of the total flight distance at supersonic speeds above an altitude of 12,192 meters (40,000 feet). Geographical locations are given for selected turbulence encounters. For 22 flights the airplane was instrumented to measure true gust velocities and the structural acceleration response to turbulence. The turbulence intensities measured were very low in comparison with those measured at high altitudes in other investigations. Acceleration response spectra, frequency response transfer functions, and coherence functions were computed from three turbulence encounters at Mach numbers of 0.88, 1.59, and 2.35. Results are compared with calculated studies. Frequencies from the vertical and lateral structural modes, dominant in the airplane acceleration responses, were compared with the natural frequencies of the human body in the vertical and lateral directions. 670. Saltzman, Edwin J.: In-Flight Use of Traversing Boundary-Layer Probes. NASA TN D-6428, H-640, July 1971, 71N28872, #. Two prototype traversing boundary-layer pitot probes were demonstrated in flight. A motor-operated screw-driven type of probe was used on two jet aircraft for defining boundary layers at profile edge Mach numbers from 0.2 to 2.2. The other type of traversing probe was motor driven through a Scotch yoke mechanism and was operated on several flights of the X-15 airplane. The highest free-stream Mach number reached during this series of flights was 5.6. The mechanical and electrical features of these probes are described, and photographs and conceptual drawings are included. Problems encountered during the development of the devices are described, and the solutions that were found are explained. Boundary-layer profile data are presented in several forms, and local friction coefficients derived from the profile through a Clauser type of determination are shown. 671. Swaroop, R.; and Winter, W. R.: A Statistical Technique for Computer Identification of Outliers in Multivariate Data. NASA TN D-6472, H-657, August 1971, 71N31371, #. A statistical technique and the necessary computer program for editing multivariate data are presented. The technique is particularly useful when large quantities of data are collected and the editing must be performed by automatic means. One task in the editing process is the identification of outliers, or observations which deviate markedly from the rest of the 113 sample. A statistical technique, and the related computer program, for identifying the outliers in univariate data was presented in NASA TN D-5275. The current report is a multivariate analog which considers the statistical linear relationship between the variables in identifying the outliers. The program requires as inputs the number of variables, the data set, and the level of significance at which outliers are to be identified. It is assumed that the data are from a multivariate normal population and the sample size is at least two greater than the number of variables. Although the technique has been used primarily in editing biodata, the method is applicable to any multivariate data encountered in engineering and the physical sciences. 672. Saltzman, E. J.; and Bellman, D. R.: A Comparison of Some Aerodynamic Drag Factors as Determined in Full-Scale Flight With Wind-Tunnel and Theoretical Results. NASA TM X-67413, AGARD-CP-83-71, Paper 15. Facilities and Tech. for Aerodyn. Testing at Transonic Speeds and High Reynolds Number, August 1971, pp. 16-1 to 16-9, 72N11869, #. Reliable techniques for defining flight values of overall aircraft drag and turbulent skin friction, and the drag associated with local regions of separated flow are reported. Selected results from these studies are presented for several types of aircraft, including the X-15, the XB-70, lifting bodies, and military interceptors. These flight results are compared with predictions derived from windtunnel models or, for friction, with the Karman-Schoenherr relationship. The flight experiments have defined the turbulent skin friction to Reynolds numbers somewhat above 10 to the 8th power, the overall drag of two airplanes, base pressure coefficients for aircraft and for an aft-facing step immersed in a thick boundary layer. A flight application of a splitter plate for reducing base drag is discussed along with examples of the drag associated with afterbody flow separation for shapes having relatively large afterbody closure angles. 673. Kempel, R. W.: Analysis of a Coupled Roll Spiral Mode, Pilot Induced Oscillation Experienced With the M2-F2 Lifting Body. NASA TN D-6496, H-633, September 1971, 71N33307, #. During the 16 glide flights of the M2-F2 lifting body vehicle, severe lateral pilot-induced oscillations occurred on three occasions in the low-angle-attack, final-approach, preflare situation. These oscillations were analyzed qualitatively to determine the type and similarity and by a systems analysis to determine the root cause. The analysis was complemented by a piloted simulator study, which verified the results. The systems analysis revealed the presence of a coupled rollspiral mode which caused the pilots to generate a closed-loop lateral instability in the low-angle-of-attack, preflare flight region. A systems analysis, a piloted simulator study, and flight data showed that the addition of a fixed center fin lessened the pilot-induced-oscillation tendencies in the critical flight region. 674. Martin, R. A.; and Hughes, D. L.: Comparisons of In-Flight F-111A Inlet Performance for On and Off Scheduled Inlet Geometry at Mach Numbers of 0.68 to 2.18. NASA TN D-6490, H-654, September 1971, 71N33211, #. 675. Wilson, E. J.: Strain Gage and Thermocouple Installation on a Research Airplane. Presented at the Western Regional Strain Gage Committee Meeting, Tempe, Arizona, September 20–21,1971. Strain gages with modified Karman filaments and backings of glass-fiber-reinforced epoxy resin matrices were selected use on a YF-12 aircraft. The gages were installed with an epoxy adhesive and are being used in a flight-loads measurements program. Individual laboratory performance tests indicated that the gages would be capable of operating at expected flight temperature. 676. Tang, M. H.; and Pearson, G. P. E.: FlightMeasured HL-10 Lifting Body Center Fin Loads and Control Surface Hinge Moments and Correlation With Wind-Tunnel Predictions. NASA TM X-2419, H-669, October 1971, 71N38700, #. Subsonic, transonic, and supersonic aerodynamic loads data are presented for the center fin and the control surfaces of the HL-10 lifting body vehicle. The effects of variations in angle of attack, angle of sideslip, aileron deflection, rudder deflection, and Mach number on the center fin loads are presented in terms of coefficient slopes. The effects of vehicle attitude, control surface deflection, Mach number, and rocket engine operation on the outboard and inboard tip fin flaps, rudder, elevon flap, and elevon hinge-moment coefficients are discussed. The flight test aerodynamic loads are compared with full-scale and small-scale wind-tunnel data. 677. Szalai, K. J.: Validation of a General Purpose Airborne Simulator for Simulation of Large Transport Aircraft Handling Qualities. NASA TN D-6431, H-591, October 1971, 71N37823, #. A flight simulation program was conducted to validate the general purpose airborne simulator (GPAS) for handlingqualities studies of large transport aircraft in cruise. Pilots compared flying qualities of the XB-70-1 with those simulated on the GPAS during consecutive flights of the two vehicles. In addition, various handling-qualities parameters and time histories for the XB-70 and the airborne simulator were compared to assess simulator fidelity. The GPAS was shown to be capable of accurate and realistic simulation of the XB-70 at two flight conditions (Mach 1.2 at 12,200 meters (40,000 feet) altitude and Mach 2.35 at 16,800 meters (55,000 feet) attitude. In-flight changes to the programmed model were required to obtain a satisfactory simulation from the pilot’s point of view. In most instances, these changes were necessary to improve model 114 representation of the XB-70 rather than to correct for possible simulator-introduced distortions. 678. Szalai, K. J.: Motion Cue and Simulation Fidelity Aspects of the Validation of a General Purpose Airborne Simulator. NASA TN D-6432, H-648, October 1971, 71N36672, #. In the validation of the general purpose airborne simulator, certain motion and visual cues could not be duplicated because the airborne simulator could not be independently controlled in six degrees of freedom. According to pilot opinion (NASA TN D-6431), however, the XB-70 airplane at two flight conditions had been simulated satisfactorily. Because of the dependence of simulation results on simulator configuration, two areas were investigated after the validation program was completed. The first was the effect of mismatched cues on observed handling qualities. Experiments which varied lateral acceleration at the pilot’s location and yaw rate, while keeping constant the lateraldirectional dynamics displayed on the pilot’s instruments, showed pilot sensitivity to directional motion cues to be different for the simulation of two XB-70 flight conditions. A technique for allowing consecutive evaluations of movingand fixed-base configurations in flight was used successfully to determine motion cue effects. The second area investigated was the measurement and description of simulation fidelity. In-flight frequency-response measurement of the modelfollowing system were taken to examine model-following fidelity for directly matched variables such as sideslip and roll rate as well as uncontrolled parameters such as lateral acceleration. 679. Carpenter, R.; and Roman, J.: Digital Automatic Data Reduction Techniques Used in a 1000-Flight Biomedical Study. NASA TN D-6601, H-651, December 1971, 72N12059, #. Techniques developed to automatically process a large quantity of physiological data obtained during a 1000-flight study are described. To reduce this data reliably, a study program was conducted using physiological data from X-15 flights as a data source for experimenting with signal enhancement and noise elimination techniques. The techniques include an automatic means for counting heart rates, averaging electrocardiogram waveforms, plotting histograms of heart rate versus frequency, and counting respiration rates. These techniques were used to reduce more than 2000 hours of physiological data recorded in flight. 680. Kempel, R. W.; and Thompson, R. C.: FlightDetermined Aerodynamic Stability and Control Derivatives of the M2-F2 Lifting Body Vehicle at Subsonic Speeds. NASA TM X-2413, H-520, December 1971, 72N11900, #. Aerodynamic derivatives were obtained for the M2-F2 lifting body flight vehicle in the subsonic flight region between Mach numbers of 0.41 and 0.64 and altitudes of 7000 feet to 45,000 feet. The derivatives were determined by a flight time history curve-fitting process utilizing a hybrid computer. The flight-determined derivatives are compared with wind-tunnel and predicted values. Modal-response characteristics, calculated from the flight derivatives, are presented. 681. Lewis, T. L.; and Banner, R. D.: Boundary Layer Transition Detection on the X-15 Vertical Fin Using Surface-Pressure-Fluctuation Measurements. NASA TM X-2466, H-660, December 1971, 72N12994, #. A flush-mounted microphone on the vertical fin of an X-15 airplane was used to investigate boundary layer transition phenomenon during flights to peak altitudes of approximately 70,000 meters. The flight results were compared with those from wind tunnel studies, skin temperature measurements, and empirical prediction data. The Reynolds numbers determined for the end of transition were consistent with those obtained from wind tunnel studies. Maximum surface-pressure-fluctuation coefficients in the transition region were about an order of magnitude greater than those for fully developed turbulent flow. This was also consistent with wind tunnel data. It was also noted that the power-spectral-density estimates of the surface-pressure fluctuations were characterized by a shift in power from high frequencies to low frequencies as the boundary layer changed from turbulent to laminar flow. Large changes in power at the lowest frequencies appeared to mark the beginning of transition. 682. Burcham, F. W., Jr.; and Bellman, D. R.: A Flight Investigation of Steady State and Dynamic Pressure Phenomena in the Air Inlets of Supersonic Aircraft. NASA TM X-67495, AGARD-CP-91-71, Paper 24. UDC-533.697. Inlets and Nozzles for Aerospace Eng., (see N72-16685 07-28), December 1971, 72N16709, #. The difficulty of achieving adequate inlet performance and stability and avoiding engine compressor stalls at supersonic speeds has led to the investigation of pressure phenomena in the inlets of several supersonic aircraft. Results of tests with the F-111A airplane are presented showing the inlet steady state and dynamic performance. The inlet total pressure distortion that causes compressor stall is discussed, and the requirement for high response instrumentation is demonstrated. A duct resonance encountered at Mach numbers near 2.0 is analyzed and shown to be due to a normal shock oscillation at the duct fundamental frequency. Another type of resonance, in the engine fan duct, is shown to be a possible cause of reduced engine stall margin in afterburning operation. Plans for a comprehensive inlet study of the YF-12 airplane are discussed including flight tests and full scale, 1/3 scale, and 1/12 scale wind tunnel tests. 683. Lipana, J. G.; *Masters, R. L.; and Winter, W. R.: An Operating Environmental Health Program. Proceedings of the Annual Conference of NASA Clinic 115 Directors, Environmental Health Officers and Medical Program Advisors, 1971, pp. 204–223, 73N17093. Some concepts of an operational program for medical and environmental health are outlined. Medical services of this program are primarily concerned with emergency care, laboratory examinations, advice to private physician with patient permission, medical monitoring activities, and suggestions for treatment or control of the malfunction. *Lovelace Foundation for Medical Education and Research, Albuquerque, New Mexico. 684. Barber, M. R.; and Fischel, J.: General Aviation: the Seventies and Beyond. NASA SP-292. Vehicle Technology for Civil Aviation, 1971, pp. 317–332, 72N13013, #. The possible advancements in general aviation through the applications of technology during the next decade are discussed in terms of aircraft performance, utility, safety, and public acceptance. 685. Burcham, F. W., Jr.; Calogeras, J. E.; Meyer, C. L.; Povolny, J. H.; and Rudey, R. A.: Effects of Engine Inlet Disturbances on Engine Stall Performance. NASA SP-259. Aircraft Propulsion, 1971, (see N72-19451 09-28), pp. 313–341, 71N19461, #. 686. Gee, S. W.; and Burke, M. E.: NASA Flight Research Center Fly by Wire Flight Test Program. Proceedings of the Space Shuttle Integrated Electron. Conference, Vol. 1, 1971, (see N71 33051 20-31), pp. 365–391, 71N33066, #. 687. Andrews, W. H.; Robinson, G. H.; and Larson, R. R.: Aircraft Response to the Wing Trailing Vortices Generated by Large Jet Transports. NASA SP-270. NASA Aircraft Safety and Operating Probl., Vol. 1, 1971, (see N71 30756 18-02), pp. 115–126, 71N30765, #. 688. Loschke, P. C.; Barber, M. R.; Jarvis, C. R.; and Enevoldson, E. K.: Handling Qualities of Light Aircraft With Advanced Control Systems and Displays. NASA SP–270, in NASA Aircraft Safety and Operating Probl., Vol. 1, 1971, (see N71 30756 18-02), pp. 189–206, 71N30771, #. Flight tests to determine the benefits of advanced control systems and displays on the handling qualities of general aviation aircraft, primarily during ILS (Instrument Landing System) approaches in turbulence, have shown that very significant benefits can be achieved. The use of a flightdirector display and an attitude-command control system in combination was shown to transform a typical light aircraft into a flying machine that borders on being perfect from a handling-qualities standpoint during ILS approaches in turbulent air. The singular use of either the flight director display or the attitude-command control system provided significant benefits. A rate-command control system was found to provide significantly less benefit than attitudecommand control system. NASA SP-301, 1972, (see N77 85474 24-02), pp. 59–70, 77N85479. This paper reviews the lift and drag results obtained from the first series of flights with the F-8 supercritical wing configuration. To concentrate on the performance of the wing and eliminate extraneous effects of the fuselage and propulsion system, the internal drag and base drag components have been removed from the flight and windtunnel data. Although removing these variables provides for the best comparison of the wind-tunnel and flight wing drag, which is the immediate purpose of this paper, it is somewhat unreal for purposes of assessing the ability of a designer to use wind-tunnel results to predict the absolute drag level of a complete airplane. 692. Montoya, L. C.; and Banner, R. D.: F-8 Supercritical Wing Pressure Distribution Evaluation. NASA SP-301, 1972, (see N77 85474 24-02), pp. 71–84, 77N85480. Pressure measurements were made on the F-8 Supercritical Wing in flight. This paper presents some of these data, compares them with Langley 8-foot wind-tunnel results, and relates the measurements to the drag and buffet characteristics of the complete configuration. 693. DeAngelis, V. M.; and Banner, R. D.: Buffet Characteristics of the F-8 Supercritical Wing Airplane. NASA SP-301, 1972, (see N77 85474 24-02), pp. 85–96, 77N85481. Airplane and wing structural response measurements were made to show the buffet characteristics of the F-8 Supercritical Wing Flight Program at transonic speeds. This paper presents some of the preliminary results of that investigation. Wing structural response was used to sense the buffet of the wing, and these data are compared with windtunnel-model data and the wing flow characteristics at transonic speeds. 694. McMurtry, T. C.; Matheny, N. W.; and Gatlin, D. H.: Piloting and Operational Aspects of the F-8 Supercritical Wing Airplane. NASA SP-301, 1972, (see N77-85474 24-02), pp. 97–110, 77N85482. This paper considers both the overall handling characteristics of the test vehicle and the correlation of flight data with windtunnel results. It should be pointed out that the basic intent of the program is to validate the wing concept and design approach. An effort was made to achieve acceptable handling qualities; however, time and cost constraints made it impossible to optimize them. 695. Weil, J.; and *Dingeldein, R. C.: Summary and Future Plans. NASA SP-301, 1972, (see N77-85474 24-02), pp. 121–133, 77N85484. 116 1972 Technical Publications 689. Anon.: Supercritical Wing Technology: A Report on Flight Evaluations. NASA SP-301, 1972, 77N85474. The papers in this compilation were presented at the NASA Symposium on “Supercritical Wing Technology: A Progress Report on Flight Evaluations,” held at the NASA Flight Research Center, Edwards, Calif, on February 29, 1972. The purpose of the symposium was to present timely information on flight results obtained with the F-8 and T-2C supercritical wing configurations, discuss comparisons with wind-tunnel predictions, and project follow-on flight programs planned for the F-8 and F-111 (TACT) airplanes. Papers were presented by representatives of the NASA Flight Research Center, the NASA Langley Research Canter, and North American Rockwell-Columbus Division. 690. Andrews, W. H.: Status of the F-8 Supercritical Wing Program. NASA SP-301, 1972, (see N77-85474 24-02), pp. 49–58 , 77N85478. This paper discusses the modifications incorporated in the test airplane and the status of the program. The flight program of the F-8 supercritical wing test-bed airplane has proceeded in an orderly manner, particularly in view of the difficulties of testing in the transonic speed range in either wind tunnels or flight. An attempt has been made to acquire accurate data from precise state-of-the-art instrumentation and test techniques. EC73-3468 F-8 Supercritical Wing Airplane 691. Pyle, J. S.: Preliminary Lift and Drag Characteristics of the F-8 Supercritical Wing Airplane. A summary of where we have been and where we think we may be going in supercritical wing proof-of-concept flight testing. *NASA Langley Research Center, Hampton, Virginia. 696. Bellman, D. R.; Burcham, F. W., Jr.; and Taillon, N. V.: Techniques for the Evaluation of Air-Breathing Propulsion Systems in Full-Scale Flight. NASA TM X-68305, AGARD-CP-85, Paper 7. Flight Test Tech. (see N72 20976 12-02), February 1972, 72N20983, #. Techniques for evaluating air breathing propulsion systems in full scale flight are discussed. Examples of flight test techniques being used to measure the performance of turbojet propulsion systems are presented. Included are the determination of jet engine thrust, the study of inlet pressure phenomena, the measurement of exhaust nozzle characteristics, and the use of tufts at supersonic speeds. A flow diagram of a gas generator method of thrust calculation is illustrated. 697. Layton, G. P., Jr.; and Thompson, M. O.: Lifting Body Flight-Test Techniques. NASA TM X-68306, AGARD-CP-85, Paper 10. Flight Test Tech., (see N72-20976 12-02), February 1972, 72N20986, #. Specific techniques and procedures for conducting flight tests of lifting body type aircraft are presented. The characteristics of the aircraft in transonic and supersonic flight regions were investigated. The data collection and analysis techniques with which the flight results were analyzed are outlined. Included are analog and digital matching techniques for derivative extractions and a method for extracting lift and drag data. Problems encountered in the flight test program and methods for solving these problems are discussed. 698. Sisk, Thomas R.: A Proposed Flight-Test Technique to Assess Fighter Aircraft Maneuverability. Presented at Air-to-Air Combat Analysis and Simulator Symposium, Kirkland AFB, New Mexico, February 29–March 2, 1972. Recent emphasis on air-to-air combat has led the National Aeronautics and Space Administration to intensify research into methods for improving the transonic maneuverability of fighter aircraft. As a part of this effort, the Flight Research Center has been conducting flight programs utilizing various aircraft to determine the factors affecting tracking precision and to investigate ways of improving transonic maneuverability. Fixed-reticle gunsights and cameras have been installed in various aircraft, and closed-loop tracking maneuvers have been performed throughout the transonic Mach range to the maximum load factor capability of each aircraft. Analysis of the gun camera film in conjunction with airplane response parameters from onboard instrumentation has been found to provide excellent means of assessing the overall maneuvering capability of different aircraft and 117 alternate configurations of the same aircraft. A flight technique for evaluating aircraft agility (i.e., combined performance and handling qualities) has grown out of these studies and is proposed for consideration in assisting in the air superiority evaluation task. An agility index is suggested for use with early design data. 699. Fields, R. A.; Olinger, F. V.; and Monaghan, R. C.: Experimental Investigation of Mach 3 Cruise Heating Simulations on a Representative Wing Structure for Flight Loads Measurement. NASA TN D-6749, H-676, March 1972, 72N19922, #. Radiant heating experiments were performed in the laboratory on an instrumented multispar wing structure to investigate: (1) how accurately the structural temperatures of a Mach 3 cruise-flight profile could be simulated, (2) what the effects of the heating and heating inaccuracies would be on the responses of strain-gage bridges installed on the structure, and (3) how these responses would affect flight loads measurements. Test temperatures throughout the structure agreed well with temperatures calculated for a Mach 3 profile. In addition, temperatures produced by two identical tests were repeatable to less than ±6 K deg. Thermally induced strain-gage-bridge responses were large enough to be detrimental to a high-speed flight loads program with a goal of establishing aerodynamic loads (exclusive of thermal loads). It was shown that heating simulation can be used effectively for thermal calibration (that is, to provide corrections for a high-temperature environment), and that thermal calibration may not be needed if the simulation data are used to carefully select bridges and load equations. 700. Iliff, K. W.; and Taylor, L. W., Jr.: Determination of Stability Derivatives From Flight Data Using a Newton-Raphson Minimization Technique. NASA TN D-6579, H-626, March 1972, 72N19659, #. A modified Newton-Raphson or quasilinearization minimization technique for determining stability derivatives from flight data was developed and compared with simpleequations, analog-matching, least-squares, and Shinbrot methods of analysis. For the data analyzed, the solutions computed by using the estimates obtained from the NewtonRaphson technique fit the data and determined coefficients adequately. A further modification to include a priori information was found to be useful. A model statistically similar to the flight data was analyzed using the same methods (excluding analog matching), and the NewtonRaphson technique was found to yield superior estimates. An approximate Cramer-Rao bound was compared with the error covariance matrix of the model and was found to provide information about the reliability of the individual estimates obtained. The technique was successfully applied to data obtained from a light airplane, a large supersonic airplane, and a lifting body vehicle. It was shown that the reliability of the estimates of a given coefficient obtained from these vehicles depends upon the data analyzed. 701. Kock, B. M.; Fulton, F. L.; and Drinkwater, F. J., III: Low-Lift-to-Drag-Ratio Approach and Landing Studies Using a CV-990 Airplane. NASA TN D-6732, H-672, March 1972, 72N19022, #. The results are presented of a flight-test program utilizing a CV-990 airplane, flow in low-lift-to-drag-ratio (L/D) configurations, to simulate terminal area operation, approach, and landing of large unpowered vehicles. The results indicate that unpowered approaches and landings are practical with vehicles of the size and performance characteristics of the proposed shuttle vehicle. Low L/D landings provided touchdown dispersion patterns acceptable for operation on runways of reasonable length. The dispersion pattern was reduced when guidance was used during the final approach. High levels of pilot proficiency were not required for acceptable performance. 702. Hughes, D. L.; Holzman, J. K.; and Johnson, H. J.: Flight-Determined Characteristics of an Air Intake System on an F-111A Airplane. NASA TN D-6679, H-661, March 1972, 72N18996, #. Flow phenomena of the F-111A air intake system were investigated over a large range of Mach number, altitude, and angle of attack. Boundary-layer variations are shown for the fuselage splitter plate and inlet entrance stations. Inlet performance is shown in terms of pressure recovery, airflow, mass-flow ratio, turbulence factor, distortion factor, and power spectral density. The fuselage boundary layer was found to be not completely removed from the upper portion of the splitter plate at all Mach numbers investigated. Inlet boundary-layer ingestion started at approximately Mach 1.6 near the translating spike and cone. Pressure-recovery distribution at the compressor face showed increasing distortion with increasing angle of attack and increasing Mach number. The time-averaged distortion-factor value approached 1300, which is near the distortion tolerance of the engine at Mach numbers above 2.1. 703. Larson, R. R.: Statistical Analysis of Landing Contact Conditions for Three Lifting Body Research Vehicles. NASA TN D-6708, H-684, March 1972, 72N18895, #. The landing contact conditions for the HL-10, M2-F2/F3, and the X-24A lifting body vehicles are analyzed statistically for 81 landings. The landing contact parameters analyzed are true airspeed, peak normal acceleration at the center of gravity, roll angle, and roll velocity. Ground measurement parameters analyzed are lateral and longitudinal distance from intended touchdown, lateral distance from touchdown to full stop, and rollout distance. The results are presented in the form of histograms for frequency distributions and cumulative frequency distribution probability curves with a Pearson Type 3 curve fit for extrapolation purposes. 118 704. Andrews, W. H.; Robinson, G. H.; and Larson, R. R.: Exploratory Flight Investigation of Aircraft Response to the Wing Vortex Wake Generated by Jet Transport Aircraft. NASA TN D-6655, H-671, March 1972, 72N18003, #. The effect of intercepting wing tip vortices generated by large jet transports, including jumbo jets, over separation distances from 1 nautical mile to 15 nautical miles is evaluated on the basis of the response of a vortex probe airplane in the roil mode. The vortex probe test aircraft included a representative general aviation airplane, an executive jet, a fighter, and light and medium weight jet transports. The test conditions and airplane configurations were comparable to those normally used during takeoff, landing, or holding pattern operations. For flight safety the tests were performed at altitudes from 9500 feet to 12,500 feet. In addition to an evaluation of the probe airplane response, a flight test technique is suggested for determining minimum separation distance, using as variable the ratio of vortex-induced roll acceleration to maximum lateral control acceleration and the gross weight of the generating aircraft. 705. Nugent, J.; Sakamoto, G. M.; and Webb, L. D.: Flight-Test Results From Two Total Temperature Probes for Air-Data Measurements Up to 2014 K (3625 R). NASA TN D-6748, H-668, March 1972, 72N20399, #. An experimental temperature probe package containing a fluidic oscillator temperature probe and a shielded thermocouple temperature probe was tested during several X-15 flights. The X-15 flights provided greatly varying test conditions, including a wide range of rapidly changing total temperatures and Mach numbers which extended from subsonic to hypersonic speeds. Within restricted ranges of free-stream Mach number, free-stream unit weight flow, and local stagnation pressure, both probes yielded ramp outputs of temperature parallel to ramp inputs of free-stream total temperature. Within these ranges both probes were used to determine total temperature in the Mach 6 temperature environment. Because ambient temperature was known, both probes were used to estimate velocity and Mach number. 706. Loschke, P. C.; Barber, M. R.; Jarvis, C. R.; and *Enevoldson, E. K.: Flight Evaluations of the Effect of Advanced Control Systems and Displays on the Handling Qualities of a General Aviation Airplane. SAE Paper 720316, March 1972, 72A25580. Flight tests have shown that, by means of improved displays and advanced control systems, it is possible to transform a typical light airplane into a flying machine that borders on being perfect from a handling-qualities standpoint. A flightdirector display and an attitude-command control system used in combination transformed a vehicle with poor handling qualities during ILS approaches in turbulent air into a vehicle with extremely good handling qualities. The attitude-command control system also improved the ride qualities of the airplane. A rate-command control system was less beneficial than an attitude-command control system. Although this paper deals primarily with general aviation aircraft, the results presented pertain to other types of aircraft. Short-takeoff-and-landing (STOL) aircraft would be a natural application of the control systems because, as a result of their low speeds, they encounter many of the handling-qualities problems noted on light aircraft. The improved ride qualities should be of interest to all airline operations, and for STOL aircraft in particular, because of their prolonged exposure to low-altitude turbulence. 707. Carpenter, R.; and Manke, J.: Flight Experiments to Determine Visibility Requirements for Approaches and Landings. Presented at the AIAA and NASA, Space Shuttle Operations, Maintenance, and Safety Technology Conference, Cocoa Beach, Florida, March 29, 1972, 72A31697, #. Some of the effects of horizontal visual restriction on the front cockpit of a T-33 aircraft were studied. These studies are pertinent to the establishment of guidelines that will be used in canopy design for limited visibility situations. Results of the study revealed that runway extension lines are helpful for restricted visibility situations. The superiority of a 300-foot runway over a 200-foot runway was greater than expected from geometric considerations. It was also shown that practice learning has a noticeable effect on performance. Finally, visibility restrictions that force a pilot into shallow glides should be avoided, and the available visibility should be sufficient to provide adequate information so that the pilot can solve the lateral-directional and pitch tasks simultaneously. 708. Webb, L. D.; and Washington, H. P.: Flight Calibration of Compensated and Uncompensated PitotStatic Airspeed Probes and Application of the Probes to Supersonic Cruise Vehicles. NASA TN D-6827, H-665, May 1972, 72N24016, #. Static pressure position error calibrations for a compensated and an uncompensated XB-70 nose boom pitot static probe were obtained in flight. The methods (Pacer, accelerationdeceleration, and total temperature) used to obtain the position errors over a Mach number range from 0.5 to 3.0 and an altitude range from 25,000 feet to 70,000 feet are discussed. The error calibrations are compared with the position error determined from wind tunnel tests, theoretical analysis, and a standard NACA pitot static probe. Factors which influence position errors, such as angle of attack, Reynolds number, probe tip geometry, static orifice location, and probe shape, are discussed. Also included are examples showing how the uncertainties caused by position errors can affect the inlet controls and vertical altitude separation of a supersonic transport. 709. Taylor, L. W., Jr.; and Iliff, K. W.: Systems Identification Using a Modified Newton-Raphson 119 Method: A FORTRAN Program. NASA TN D-6734, L-8028, May 1972, 72N22581, #. A FORTRAN program is offered which computes a maximum likelihood estimate of the parameters of any linear, constant coefficient, state space model. For the case considered, the maximum likelihood estimate can be identical to that which minimizes simultaneously the weighted mean square difference between the computed and measured response of a system and the weighted square of the difference between the estimated and a priori parameter values. A modified Newton-Raphson or quasilinearization method is used to perform the minimization which typically requires several iterations. A starting technique is used which insures convergence for any initial values of the unknown parameters. The program and its operation are described in sufficient detail to enable the user to apply the program to his particular problem with a minimum of difficulty. 710. Gilyard, G. B.: Flight-Determined Derivatives and Dynamic Characteristics of the CV-990 Airplane. NASA TN D-6777, H-693, May 1972, 72N23027, #. Flight-determined longitudinal and lateral-directional stability and control derivatives are presented for the CV-990 airplane for various combinations of Mach number, altitude, and flap setting throughout the flight envelope up to a Mach number of 0.87. Also presented are the dynamic characteristics of the aircraft calculated from the flightobtained derivatives and the measured phugoid characteristics. The derivative characteristics were obtained from flight records of longitudinal and lateral-directional transient oscillation maneuvers by using a modified NewtonRaphson digital derivative determination technique. Generally the derivatives exhibited consistent variation with lift coefficient in the low-speed data and with Mach number and altitude in the high-speed data. Many also varied with flap deflection, notably spoiler effectiveness and directional stability. 711. Holleman, E. C.; and Gilyard, G. B.: In-Flight Pilot Evaluations of the Flying Qualities of a Four-Engine Jet Transport. NASA TN D-6811, H-680, May 1972, 72N22026, #. The flying qualities of the CV-990 jet transport were evaluated over the normal operating flight envelope and in smooth air to provide baseline data for transport airplanes. Pilot ratings of airplane handling characteristics for specific test conditions and configurations from approach to normal cruise were compared with various flying qualities criteria. In general, the CV-990 flying qualities were evaluated as satisfactory, and the evaluations supported transport flying qualities criteria. The dutch roll damping was rated more satisfactory than was predicted by the flying qualities criteria. The pilots found rudder coordination for the yaw generated during high roll rates very difficult. They preferred to control with roll and pitch controls and to use the yaw damper to provide the required rudder coordination. 712. *Leverett, S. D., Jr.; *Davis, H. M., Jr.; and Winter, W. R.: Physiological Response in Pilot/Back-Seat Man During Aerial Combat Maneuvers in F-4E Aircraft. Aerospace Medical Association, 43rd Annual Scientific Meeting, Bal Harbour, Florida, May 8–11, 1972, Preprints, 1972, (see A72-28251 12-04), pp. 192–193, 72A28317, #. Comparison of objective/subjective physiological data between the pilot and the back-seat man during training within the G maneuvering envelope. It appears that the psychological requirements for the pilot to be mentally alert and physiologically adapted to a continually changing environment places additional responsibility on him to the extent the physiological signs monitored are indicative of a high stress condition and are increased by a significant amount over the back-seat man who is, in most instances, riding passively. *USAF School of Aerospace Medicine, Brooks AFB, Texas. 713. Painter, W. D.; and Sitterle, G. J.: Ground and Flight Test Methods for Determining Limit Cycle and Structural Resonance Characteristics of Aircraft Stability Augmentation Systems. NASA TN D-6867, H-682, June 1972, 72N26017, #. Performance criteria and test techniques are applied to stability augmentation systems (SAS) during ground testing to predict objectionable limit cycles and preclude structural resonance during flight. Factors that give rise to these problems, means of suppressing their effects, trade-offs to be considered, and ground test methods that have been developed are discussed. SAS performance predicted on the basis of these tests is compared with flight data obtained from three lifting body vehicles and the X-15 research airplane. Limit cycle and structural resonance test criteria, based upon ground and flight experience and data, were successfully applied to these vehicles. The criteria used were: The limit cycle amplitude (SAS gain multiplied by peak-to-peak angular rate) shall not exceed 0.5 deg for the highest product of control power and SAS gain that will be used in flight; the maximum in-flight SAS gain should never exceed 50 percent of the value at which a structural resonance can be sustained during ground test. 714. Wolowicz, C. H.; and Yancey, R. B.: Longitudinal Aerodynamic Characteristics of Light, Twin-Engine, Propeller-Driven Airplanes. NASA TN D-6800, H-646, June 1972, 72N26006, #. Representative state-of-the-art analytical procedures and design data for predicting the longitudinal static and dynamic stability and control characteristics of light, propeller-driven airplanes are presented. Procedures for predicting drag characteristics are also included. The procedures are applied to a twin-engine, propeller-driven airplane in the clean configuration from zero lift to stall conditions. The calculated characteristics are compared with wind-tunnel and flight 120 data. Included in the comparisons are level-flight trim characteristics, period and damping of the short-period oscillatory mode, and windup-turn characteristics. All calculations are documented. 715. Bikle, P.: Sailplane Performance Measured in Flight. Presented at the 12th OSTIV Congress, Alpine, Texas, 1970, Aero-Revue, June 1972, pp. 333–338, 72A34215. (Also published in Technical Soaring, January 1972, Vol. 1, No. 3.) Description of the T-6, a modified HP-14 sailplane, the performance data obtained, and the test techniques, as well as the comparison tests and results obtained for seven other sailplanes. These are the Kestrel, Cirrus, Phoebus C, 16.5-m Diamant, Phoebus A, BG-12, and 1-26. The T-6 is of allmetal construction, has a shoulder-high wing, a retractable gear, simple hinged flaps with no speed brakes or tail parachute, and is of medium aspect ratio and wing loading. 716. Putnam, T. W.; and Lasagna, P. L.: Externally Blown Flap Impingement Noise. AIAA Paper 72-664, June 1972, 72A35961, #. (See also 737.) An investigation of externally blown flap impingement noise was conducted using a full-scale turbofan engine and aircraft wing. The noise produced with a daisy nozzle installed on the engine exhaust system was greater than that produced by a conical nozzle at the same thrust. The daisy nozzle caused the jet velocity to decay about 35 percent at the flap. The presence of the wing next to the conical nozzle increased the noise, as did increasing the flap deflection. Compared with the conical nozzle, the daisy nozzle produced slightly less noise at a flap deflection of 60 deg but produced more noise at the lower flap deflections tested. 717. Powers, B. G.: Statistical Survey of XB-70 Airplane Responses and Control Usage With an Illustration of the Application to Handling Qualities Criteria. NASA TN D-6872, H-663, July 1972, 72N27013, #. The magnitude and frequency of occurrence of aircraft responses and control inputs during 27 flights of the XB-70 airplane were measured. Exceedance curves are presented for the airplane responses and control usage. A technique is presented which makes use of these exceedance curves to establish or verify handling qualities criteria. This technique can provide a means of incorporating current operational experience in handling qualities requirements for future aircraft. 718. Smith, J. P.: Research Aircraft Simulators. Western Simulator Council, Los Angeles, California, July 27, 1972. 719. Friend, E. L.; and Sefic, W. J.: Flight Measurements of Buffet Characteristics of the F-104 Airplane for Selected Wing-Flap Deflections. NASA TN D-6943, H-666, August 1972, 72N30004, #. A flight program was conducted on the F-104 airplane to investigate the effects of moderate deflections of wing leading- and trailing-edge flaps on the buffet characteristics at subsonic and transonic Mach numbers. Selected deflections of the wing leading and trailing-edge flaps, individually and in combination, were used to assess buffet onset, intensity, and frequency; lift curves; and wing-rock characteristics for each configuration. Increased deflection of the trailing-edge flap delayed the buffet onset and buffet intensity rise to a significantly higher airplane normal-force coefficient. Deflection of the leading-edge flap produced some delay in buffet onset and the resulting intensity rise at low subsonic speeds. Increased deflection of the trailing-edge flap provided appreciable lift increments in the angle-ofattack range covered, whereas the leading-edge flap provided lift increments only at high angles-of-attack. The pilots appreciated the increased maneuvering envelope provided by the flaps because of the improved turn capability. 720. Edwards, J. W.: Analysis of an Electrohydraulic Aircraft Control Surface Servo and Comparison With Test Results. NASA TN D-6928, H-629, August 1972, 72N30002, #. An analysis of an electrohydraulic aircraft control-surface system is made in which the system is modeled as a lumped, two-mass, spring-coupled system controlled by a servo valve. Both linear and nonlinear models are developed, and the effects of hinge-moment loading are included. Transfer functions of the system and approximate literal factors of the transfer functions for several cases are presented. The damping action of dynamic pressure feedback is analyzed. Comparisons of the model responses with results from tests made on a highly resonant rudder control-surface servo indicate the adequacy of the model. The effects of variations in hinge-moment loading are illustrated. 721. Anon.: Basic Research Review for the NASA OAST Research Council. NASA TM-74910, August 2, 1972, 77N84624. In preparing this report on the basic research activities included in the flight research programs at this Center, it became apparent that many elements of the multidisciplinary programs would be reported because of the difficulty of isolating basic research. Furthermore, in reporting the progress during the past year, it was felt that some cognizance and response should be given to the comments and suggestions of the Office of Aeronautics and Space Technology (OAST) Research Council after the last annual review. Details of the present day flight research programs at the Flight Research Center, including basic research elements, are presented in this review. 121 722. Deets, D. A.; and Szalai, K. J.: Design and Flight Experience With a Digital Fly-By-Wire Control System Using Apollo Guidance System Hardware on an F-8 Aircraft. AIAA Paper 72-881, presented at the AIAA Guidance and Control Conference, Stanford, California, August 14–16, 1972, 72A40060, #. This paper discusses the design and initial flight tests of the first digital fly-by-wire system to be flown in an aircraft. The system, which used components from the Apollo guidance system, was installed in an F-8 aircraft. A lunar module guidance computer is the central element in the three-axis, single-channel, multimode, digital, primary control system. An electrohydraulic triplex system providing unaugmented control of the F-8 aircraft is the only backup to the digital system. Emphasis is placed on the digital system in its role as a control augmentor, a logic processor, and a failure detector. A sampled-data design synthesis example is included to demonstrate the role of various analytical and simulation methods. The use of a digital system to implement conventional control laws was shown to be practical for flight. Logic functions coded as an integral part of the control laws were found to be advantageous. Verification of software required an extensive effort, but confidence in the software was achieved. Initial flight results showed highly successful system operation, although quantization of pilot’s stick and trim were areas of minor concern from the piloting standpoint. ECN-3276 F-8 Digital Fly-by-Wire Airplane 723. Edwards, J. W.: Flight Test Experience in Digital Control of a Remotely Piloted Vehicle. AIAA Paper 72-883, presented at the AIAA Guidance and Control Conference, Stanford, California, August 14–16, 1972. 724. Strutz, L. W.: Flight-Determined Derivatives and Dynamic Characteristics for the HL-10 Lifting Body Vehicle at Subsonic and Transonic Mach Numbers. NASA TN D-6934, H-708, September 1972, 72N30903, #. The HL-10 lifting body stability and control derivatives were determined by using an analog-matching technique and compared with derivatives obtained from wind-tunnel results. The flight derivatives were determined as a function of angle of attack for a subsonic configuration at Mach 0.7 and for a transonic configuration at Mach 0.7, 0.9, and 1.2. At an angle of attack of 14 deg, data were obtained for a Mach number range from 0.6 to 1.4. The flight and wind-tunnel derivatives were in general agreement, with the possible exception of the longitudinal and lateral damping derivatives. Some differences were noted between the vehicle dynamic response characteristics calculated from flight-determined derivatives and those predicted by the wind-tunnel results. However, the only difference the pilots noted between the response of the vehicle in flight and the response of a simulator programmed with wind-tunnel-predicted data was that the damping generally was higher in the flight vehicle. 725. Schweikhard, William G.; and Montoya, E. J.: Research Instrumentation Requirements for Flight/ Wind-Tunnel Tests of the YF-12 Propulsion System and Related Experience. Presented at the Symposium on Instrumentation for Airbreathing Propulsion, Naval Postgraduate School, Monterey, California, September 19–21, 1972. 726. Smith, Ronald H.; and Burcham, Frank W., Jr.: Instrumentation for In-Flight Determination of Steady State and Dynamic Inlet Performance in Supersonic Aircraft. Presented at the Symposium on Instrumentation for Airbreathing Propulsion, Naval Postgraduate School, Monterey, California, September 19–21, 1972. (See also 819.) Advanced instrumentation and techniques for in-flight measurements of air inlet performance of the XB-70, F-111A, and YF-12 supersonic airplanes were developed and evaluated in flight tests at the NASA Flight Research Center. A compressor face rake with in-flight zeroing capability was flown on the F-111A and found to give excellent steady state as well as high frequency response pressure data. The severe temperature environment of the YF-12 necessitated development of special high temperature transducers. Mounting these transducers to give the required 500-hertz frequency response required some special rake designs. Vibration test requirements necessitated some modifications to the rakes. The transducers and rakes were evaluated in flight tests and were found to function properly. Preliminary data have been obtained from the YF-12 propulsion program in flights that began in May 1972. One example shows the terminal shock wave effects on cowl surface pressures during bypass and spike motions. 727. Gilyard, G. B.; Berry, D. T.; and Belte, D.: Analysis of a Lateral-Directional Airframe/Propulsion System Interaction of a Mach 3 Cruise Aircraft. AIAA Paper 72-961, presented at the 2nd AIAA Flight Mechanics 122 Conference, Palo Alto, California, September 11–13, 1972, 72A42348, #. Mach 3 flight data from a YF-12 airplane are analyzed to determine the causes of the significant reduction in dutch roll damping induced by automatic inlet operation. Two stability derivative extraction techniques, the time vector and the modified Newton-Raphson, were applied to the flight data to determine the forces and moments resulting from the variable geometry of the engine inlet. The sideslip angle measurement, which is fed to the inlet for terminal shock stabilization, was found to have a sensor time lag of approximately 0.5 second. These results are then included in a root locus analysis of the overall airframe/propulsion system. 728. Lewis, T. L.; and Dods, J. B., Jr.: Wind Tunnel Measurements of Surface Pressure Fluctuations at Mach Numbers of 1.6, 2.0, and 2.5 Using 12 Different Transducers. NASA TN D-7087, H-700, October 1972, 72N33387, #. The turbulent boundary layer on the wall of a 9 by 7 foot wind tunnel was measured with 12 different transducers at Mach numbers of 1.6, 2.0, and 2.5. The results indicated that the wall surface-pressure-fluctuation field was more homogeneous at a Mach number of 2.5 than at Mach numbers of 1.6 or 2.0. A comparison of power-spectral-density data at Mach 2.5 with a summary of similar data (Mach 0.1 to 3.45) showed good agreement. The measurement uncertainty was greatest when frequencies were low and the surface-pressurefluctuation field was homogeneous. The uncertainty at higher frequencies increased as the surface-pressure-fluctuation field became more inhomogeneous. Since transducer mounting effects and system noise levels were determined not to have contributed appreciably to measurement uncertainties, the result was attributed to an interaction between the surface-pressure-fluctuation field and the transducers. Corcos’ correction for size effects improved the comparison between transducers at the high frequencies, but did not eliminate an apparent size effect at the lower frequencies. 729. Holleman, E. C.; and Powers, B. G.: Flight Investigation of the Roll Requirements for Transport Airplanes in the Landing Approach. NASA TN D-7062, H-711, October 1972, 72N33019, #. An in-flight evaluation of transport roll characteristics in the landing approach was made with a general purpose airborne simulator. The evaluation task consisted of an instrument approach with a visual correction for a (200-foot) lateral offset. Pilot evaluations and ratings were obtained for approaches made at 140 knots and 180 knots indicated airspeed with variations of wheel characteristics, maximum roll rate, and roll time constant. 730. Wolowicz, C. H.; and Yancey, R. B.: LateralDirectional Aerodynamic Characteristics of Light, TwinEngine, Propeller Driven Airplanes. NASA TN D-6946, H-694, October 1972, 73N11016, #. Analytical procedures and design data for predicting the lateral-directional static and dynamic stability and control characteristics of light, twin engine, propeller driven airplanes for propeller-off and power-on conditions are reported. Although the consideration of power effects is limited to twin engine airplanes, the propeller-off considerations are applicable to single engine airplanes as well. The procedures are applied to a twin engine, propeller driven, semi-low-wing airplane in the clean configuration through the linear lift range. The calculated derivative characteristics are compared with wind tunnel and flight data. Included in the calculated characteristics are the spiral mode, roll mode, and dutch roll mode over the speed range of the airplane. 731. Gee, S. W.; and Wolf, T. D.: NASA Ride Quality Program at the Flight Research Center. Symposium on Vehicle Ride Quality, October 1972, (see N73-10012 01-02), pp. 247–251, 73N10025. A flight test program to determine the effects of low frequency vibrations on passengers in short haul aircraft is discussed. The objective of the program is to accumulate flight test data on aircraft ride quality in terms of vehicle motion and acceleration and human responses. The subjects discussed are: (1) test procedures, (2) data processing, and (3) the program schedule. 732. *Wells, T. L.; *Canon, R. F.; *Rolls, G. C.; and Wilson, E. J.: Development of a High Temperature Fatigue Sensor. Proceedings, Instrument Society of America, 27th Annual Conference, Part 2, New York, New York, October 9–12, 1972, 1972, 73A22504. An experimental program was conducted to extend the Tracor Safety Gauge (patent pending) to elevated temperature service. The Safety Gauge is based on a conductive composite device which can be fabricated to function as a fatigue sensor that undergoes an irreversible resistance increase which results from cumulative strain damage. Prototype sensors were developed which appear capable of 1000 deg F operation for short periods of time (approaching one hour); however, bonding difficulties currently limit their use to about 775 deg F. The resistance change of the sensor was generally on the order of 400% or greater as the fatigue life of a titanium alloy (Ti-5Al-2.5Sn) test specimen was approached. *Tracor, Inc., Austin, Texas. 733. Robinson, G. H.; and Larson, R. R.: A Flight Evaluation of Methods for Predicting Vortex Wake 123 Effects on Trailing Aircraft. NASA TN D-6904, H-712, November 1972, 73N12033, #. The results of four current analytical methods for predicting wing vortex strength and decay rate are compared with the results of a flight investigation of the wake characteristics of several large jet transport aircraft. An empirical expression defining the strength and decay rate of wake vortices is developed that best represents most of the flight-test data. However, the expression is not applicable to small aircraft that would be immersed in the vortex wake of large aircraft. 734. Wolowicz, C. H.; Iliff, K. W.; and Gilyard, G. B.: Flight Test Experience in Aircraft Parameter Identification. AGARD-CP-119, Paper 23. Stability and Control, November 1972, 73N17012. An automatic method for determining stability and control derivatives from flight data is presented. The technique, a modification of the Newton-Raphson method for derivative extraction, has a priori provision that makes use of initial estimates of the derivatives and provides a means of checking the validity of the results. Recommendations for applications of the method are included. 735. Hughes, D. L.: Survey of Wing and Flap LowerSurface Temperatures and Pressures During Full-Scale Ground Tests of an Externally Blown Flap System. NASA SP-320, 1972 (see N73-32934 24-02), 73N32947. Full-scale ground tests of an externally blown flap system were made using the wing of an F-111B airplane and a CF700 engine. Pressure and temperature distributions were determined on the undersurface of the wing, vane, and flap for two engine exhaust nozzles (conical and daisy) at several engine power and engine/wing positions. The tests were made with no airflow over the wing. The leading-edge wing sweep angle was fixed at 26 deg, the angle of incidence between the engine and the wing was fixed at 3 deg, and the tests were conducted with the flap retracted, extended and deflected 35 deg, and extended and deflected 60 deg. The integrated local pressures on the undersurface of the flap produced loads approximately three times as great at the 60 deg flap position as at the 35 deg flap position. With both nozzle configurations, more than 90 percent of the integrated pressure loads were contained within plus or minus 20 percent of the flap span centered around the engine exhaust centerline. The maximum temperature recorded on the flaps was 218 C (424 F) for the conical nozzle and 180 C (356 F) for the daisy nozzle. 736. Gee, S. W.; Barber, M. R.; and McMurtry, T. C.: A Flight Evaluation of Curved Landing Approaches. NASA SP-320, 1972, pp. 245–258, 73N32953. The development of STOL technology for application to operational short-haul aircraft is accompanied by the requirement for solving problems in many areas. One of the most obvious problems is STOL aircraft operations in the terminal area. The increased number of terminal operations needed for an economically viable STOL system as compared with the current CTOL system and the incompatibility of STOL and CTOL aircraft speeds are positive indicators of an imminent problem. The high cost of aircraft operations, noise pollution, and poor short-haul service are areas that need improvement. A potential solution to some of the operational problems lies in the capability of making curved landing approaches under both visual and instrument flight conditions. 737. Lasagna, P. L.; and Putnam, T. W.: Externally Blown Flap Impingement Noise. NASA SP-320, 1972, pp. 427–441, 73N32964. (See also 716.) Tests of the noise produced by the impingement of the jet exhaust on the wing and flap for an externally blown flap system were conducted with a CF700 turbofan engine and an F-111B wing panel. The noise produced with a daisy nozzle installed on the engine was greater than that produced by a conical nozzle at the same thrust. The presence of the wing next to the test nozzles increased the noise, as did increasing the flap deflection angle. Compared with the conical nozzle, the daisy nozzle produced slightly less noise at a flap deflection of 60 deg but produced more noise at the lower flap deflections tested. Tests showed that the single-slotted flap deflected 60 deg, produced less noise than the double-slotted flaps. Also, maintaining the maximum distance between the exit nozzle and flap system resulted in a minor reduction in noise. 738. Kier, D. A.; Powers, B. G.; Grantham, W. D.; and Nguyen, L. T.: Simulator Evaluation of the Flying Qualities of Externally Blown Flap and Augmentor Wing Transport Configurations. NASA SP-320, 1972, (see N73-32934 24-02), pp. 157–800 , 73N32948. Concurrent simulations of powered-lift STOL transport aircraft having either an externally blown flap configuration or an augmentor wing configuration were conducted. The following types of simulators of varying sophistication were used: (1) a simple fixed-base simulation with a simple visual display, (2) a more complex fixed-base simulation using a realistic transport cockpit and a high-quality visual display, and (3) a six-degree-of-freedom motion simulator that had a realistic transport cockpit and a sophisticated visual display. The unaugmented flying qualities determined from these simulations were rated as unacceptable for both the externally blown flap and augmentor wing configurations. The longitudinal, lateral-directional, and single-engine-failure characteristics were rated satisfactory with extensive augmentation, including pitch and roll command systems, flight-path (or speed) augmentation, turn coordination, and effective yaw damping. However, the flare and landing characteristics from any approach glide-path angle in excess of 4 deg were rated as unsatisfactory but acceptable. 124 739. Carpenter, R.; and Winter, W. R.: A Flight-Rated Liquid-Cooled Garment for Use Within a Full-Pressure Suit. NASA SP-302 (see N72-27106 18-05), 1972, 72N27121. A flight rated liquid cooled garment system for use inside a full pressure suit has been designed, fabricated, and tested. High temperature tests with this system have indicated that heat is absorbed at a rate decreasing from 224 kg-cal/hr to 143 kg-cal/hr over a 40-min period. The first 30 min are very comfortable; thereafter a gradual heat load builds that results in mild sweating at the end of the 40-min period. In flight tests during hot weather when this cooling system was worn under a regulation flight suit, the pilot reported that temperatures were comfortable and that the garment prevented sweating. 740. Dorsch, R. G.; Lasagna, P. L.; Maglieri, D. J.; and Olsen, W. A.: Flap Noise. NASA SP-311, 1972, (see N7312012 03-02), pp. 259–290, 73N12024. Externally-blown-flap noise research can be summarized by the following remarks: With lower-surface blowing, the sources of the flap noise are beginning to be understood and the noise scaling laws have been established. Further, progress has been made on suppressing the flap interaction noise at the large flap deflections used during landing. Recent small-scale noise tests of configurations using external upper-surface blowing indicate that engine-over-the-wing configurations may be promising. 1973 Technical Publications 741. Hughes, D. L.: Pressures and Temperatures on the Lower Surfaces of an Externally Blown Flap System During Full-Scale Ground Tests. NASA TN D-7138, H-729, January 1973, 73N14984, #. Full-scale ground tests of an externally blown flap system were made using the wing of an F-111B airplane and a CF700 engine. Pressure and temperature distributions were determined on the undersurface of the wing, vane, and flap for two engine exhaust nozzles (conical and daisy) at several engine power levels and engine/wing positions. The test were made with no airflow over the wing. The wing sweep angle was fixed at 26 deg; and the angle of incidence between the engine and the wing was fixed at 3 deg; and the flap was in the retracted, deflected 35 deg, and deflected 60 deg positions. The pressure load obtained by integrating the local pressures on the undersurface of the flap, F sub p was approximately three times greater at the 60 deg flap position than at the 35 deg flap position. At the 60 deg flap position, F sub p was between 40 percent and 55 percent of the engine thrust over the measured range of thrust. More than 90 percent of F sub p was contained within plus or minus 20 percent of the flap span centered around the engine exhaust centerline with both nozzle configurations. Maximum temperatures recorded on the flaps were 218 C (424 F) and 180 C (356 F) for the conical and daisy nozzles, respectively, 742. Holzman, J. K.; and Payne, G. A.: Design and Flight Testing of a Nullable Compressor Face Rake. NASA TN D-7162, H-733, January 1973, 73N16247, #. A compressor face rake with an internal valve arrangement to permit nulling was designed, constructed, and tested in the laboratory and in flight at the NASA Flight Research Center. When actuated by the pilot in flight, the nullable rake allowed the transducer zero shifts to be determined and then subsequently removed during data reduction. Design details, the fabrication technique, the principle of operation, brief descriptions of associated digital zero-correction programs and the qualification tests, and test results are included. Sample flight data show that the zero shifts were large and unpredictable but could be measured in flight with the rake. The rake functioned reliably and as expected during 25 hours of operation under flight environmental conditions and temperatures from 230 K (–46 F) to greater than 430 K (314 F). The rake was nulled approximately 1000 times. The in-flight zero-shift measurement technique, as well as the rake design, was successful and should be useful in future applications, particularly where accurate measurements of both steady-state and dynamic pressures are required under adverse environmental conditions. 743. Gilyard, G. B.: Explicit Determination of LateralDirectional Stability and Control Derivatives by Simultaneous Time Vector Analysis of Two Maneuvers. NASA TM X-2722, H-751, February 1973, 73N16010, #. An extension of the time vector technique for determining stability and control derivatives from flight data is formulated. The technique provides for explicit determination of derivatives by means of simultaneous analysis of two maneuvers which differ by a dependent control input. The control derivatives for the dependent input are also explicitly determined. This extended technique is preferable to the application of the time vector method to single maneuvers in that no estimates of derivatives are required. An example illustrating the application of the technique is given. 744. Marshall, R. T.; and Schweikhard, W. G.: Modeling of Airplane Performance From Flight-Test Results and Validation With an F-104G Airplane. NASA TN D-7137, H-723, February 1973, 73N16008, #. A technique of defining an accurate performance model of an airplane from limited flight-test data and predicted aerodynamic and propulsion system characteristics is developed. With the modeling technique, flight-test data from level accelerations are used to define a 1g performance model for the entire flight envelope of an F-104G airplane. 125 The performance model is defined in terms of the thrust and drag of the airplane and can be varied with changes in ambient temperature or airplane weight. The model predicts the performance of the airplane within 5 percent of the measured flight-test data. The modeling technique could substantially reduce the time required for performance flight testing and produce a clear definition of the thrust and drag characteristics of an airplane. 745. Montoya, E. J.: Wind-Tunnel Calibration and Requirements for In-Flight Use of Fixed Hemispherical Head Angle-of-Attack and Angle-of-Sideslip Sensors. NASA TN D-6986, H-702, March 1973, 73N18014, #. Wind-tunnel tests were conducted with three different fixed pressure-measuring hemispherical head sensor configurations which were strut-mounted on a nose boom. The tests were performed at free-stream Mach numbers from 0.2 to 3.6. The boom-angle-of-attack range was –6 to 15 deg, and the angle-of-sideslip range was –6 to 6 deg. The test Reynolds numbers were from 3.28 million to 65.6 million per meter. The results were used to obtain angle-of-attack and angle-of-sideslip calibration curves for the configurations. Signal outputs from the hemispherical head sensor had to be specially processed to obtain accurate real-time angle-of-attack and angle-of-sideslip measurements for pilot displays or aircraft systems. Use of the fixed sensors in flight showed them to be rugged and reliable and suitable for use in a high temperature environment. 746. Lewis, T. L.; Dods, J. B., Jr.; and Hanly, R. D.: Measurements of Surface-Pressure Fluctuations on the XB-70 Airplane at Local Mach Numbers Up to 2.45. NASA TN D-7226, H-714, March 1973, 73N18013, #. Measurements of surface-pressure fluctuations were made at two locations on the XB-70 airplane for nine flight-test conditions encompassing a local Mach number range from 0.35 to 2.45. These measurements are presented in the form of estimated power spectral densities, coherence functions, and narrow-band-convection velocities. The estimated power spectral densities compared favorably with wind-tunnel data obtained by other experimenters. The coherence function and convection velocity data supported conclusions by other experimenters that low-frequency surface-pressure fluctuations consist of small-scale turbulence components with low convection velocity. 747. McKay, J. M.; Kordes, E. E.; and *Wykes, J. H.: Flight Investigation of XB-70 Structural Response to Oscillatory Aerodynamic Shaker Excitation and Correlation With Analytical Results. NASA TN D-7227, H-713, April 1973, 73N24892, #. The low frequency symmetric structural response and damping characteristics of the XB-70 airplane were measured at four flight conditions: heavyweight at a Mach number of 0.87 at an altitude of 7620 meters (25,000 feet); lightweight at a Mach number of 0.86 at an altitude of 7620 meters (25,000 feet); a Mach number of 1.59 at an altitude of 11,918 meters (39.100 feet); and a Mach number of 2.38 and an altitude of 18,898 meters (62,000 feet). The flight data are compared with the response calculated by using early XB-70 design data and with the response calculated with mass, structural, and aerodynamic data updated to reflect as closely as possible the airplane characteristics at three of the flight conditions actually flown. *North American Rockwell Corp., Los Angeles, California. 748. Barber, M. R.: Application of Advanced Control System and Display Technology to General Aviation. SAE Paper 730321. Presented at the Society of Automotive Engineers, Business Aircraft Meeting, Wichita, Kansas, April 3–6, 1973, 73A34679. This paper reviews NASA’s progress in research directed toward providing the technology necessary for the application of advanced control systems and displays to general aviation aircraft, and its plans for this effort in the future. Flight evaluations of such systems as wing levelers, fluidic autopilots, yaw dampers, and angle of attack displays have been made, and test conditions and major results of some of this work are reported. Potentially valuable systems evaluated thus far are an attitude command control system and a flight-director display. As presently configured, both are prohibitively expensive for us in general aviation, however, and efforts are underway to apply technology to the goal of reducing their cost. Perhaps the most promising development in this area is called separate surface stability augmentation and plans for its implementation and flight testing are described. 749. Gee, S. W.; and *Servais, N. A.: Development of a Low-Cost Flight Director System for General Aviation. SAE Paper 730331. Presented at the Society of Automotive Engineers, Business Aircraft Meeting, Wichita, Kansas, April 3–6, 1973, 73A34684. The NASA Flight Research Center awarded a contract to Astronautics Corporation of America to develop a low-cost flight director system for general aviation. The system that was designed is expected to cost the consumer less than $3,000, a reduction of nearly 70 percent in the total cost of available systems. The features that permit lower cost without excessive degradation in performance are use of belt drives, high-volume-production standard parts, single-box construction including gyros, and post-plate construction techniques. *Astronautics Wisconsin. Corporation, of America, Milwaukee, Paper 730304. Presented at the Society of Automotive Engineers, Business Aircraft Meeting, Wichita, Kansas, April 3–6, 1973, 73A34665. The purpose of this paper is to describe an investigation of separate surface stability augmentation systems for general aviation aircraft. The program objective were twofold: first, a wind tunnel program to determine control effectiveness of separate surfaces in the presence of main surfaces, and hinge moment feedback from separate surfaces via the main surfaces to the pilot; second, a theoretical study to determine the minimum performance of actuators and sensors that can be tolerated, the best slaving gains to be used with separate surfaces, and control authority needed for proper operation under direct pilot control, under autopilot control, and in failure situations. On the basis of the results obtained, it has been concluded that separate surface systems are feasible and advantageous for use in general aviation aircraft. *Kansas, University, Lawrence, Kansas. 751. Goecke, S. A.: Flight-Measured Base Pressure Coefficients for Thick Boundary-Layer Flow Over an Aft-Facing Step for Mach Numbers From 0.4 to 2.5. NASA TN D-7202, H-740, May 1973, 73N24317, #. A 0.56-inch thick aft-facing step was located 52.1 feet from the leading edge of the left wing of an XB-70 airplane. A boundary-layer rake at a mirror location on the right wing was used to obtain local flow properties. Reynolds numbers were near 10 to the 8th power, resulting in a relatively thick boundary-layer. The momentum thickness ranged from slightly thinner to slightly thicker than the step height. Surface static pressures forward of the step were obtained for Mach numbers near 0.9, 1.5, 2.0, and 2.4. The data were compared with thin boundary-layer results from flight and wind-tunnel experiments and semiempirical relationships. Significant differences were found between the thick and the thin boundary-layer data. 752. Johnson, H. J.; and Montoya, E. J.: Local Flow Measurements at the Inlet Spike Tip of a Mach 3 Supersonic Cruise Airplane. NASA TN D-6987, H-722, May 1973, 73N24037, #. The flow field at the left inlet spike tip of a YF-12A airplane was examined using at 26 deg included angle conical flow sensor to obtain measurements at free-stream Mach numbers from 1.6 to 3.0. Local flow angularity, Mach number, impact pressure, and mass flow were determined and compared with free-stream values. Local flow changes occurred at the same time as free-stream changes. The local flow usually approached the spike centerline from the upper outboard side because of spike cant and toe-in. Free-stream Mach number influenced the local flow angularity; as Mach number increased above 2.2, local angle of attack increased and local sideslip angle decreased. Local Mach number was generally 126 750. *Roskam, J.; Barber, M. R.; and Loschke, P. C.: Separate Surfaces for Automatic Flight Controls. SAE 3 percent less than free-stream Mach number. Impactpressure ratio and mass flow ratio increased as free-stream Mach number increased above 2.2, indicating a beneficial forebody compression effect. No degradation of the spike tip instrumentation was observed after more than 40 flights in the high-speed thermal environment encountered by the airplane. The sensor is rugged, simple, and sensitive to small flow changes. It can provide accurate inputs necessary to control an inlet. 753. Edwards, John W.: Flight Test of a Remotely Piloted Vehicle Using a Remote Digital Computer for Control Augmentation. Presented at the Symposium on Applications of Control Theory to Modern Weapons Systems, California City, California, May 9–10, 1973. 754. Wilson, E. J.: Strain Gage Installation on the YF12 Aircraft. Presented at the Society for Experimental Stress Analysis, Spring Meeting, Los Angeles, California, May 13–18, 1973, 73A35444. A flight-loads measurement program on the YF-12 aircraft required the mounting of 101 strain-gauge bridges in the fuselage, fuel tanks, control surfaces, and three stations on the left wing. The sensors were to be installed primarily on titanium and were required to operate between –70 and +600 F. Strain gauges with modified Karman filaments and backings of glass-fiber reinforced epoxy resin matrices were selected and were installed with an epoxy adhesive. Attention is given to the calibration, mounting, and performance of the sensors in flight-load measurements. 755. Fisher, D. F.; and Saltzman, E. J.: Local Skin Friction Coefficients and Boundary Layer Profiles Obtained in Flight From the XB-70-1 Airplane at Mach Numbers Up to 2.5. NASA TN D-7220, H-710, June 1973, 73N25276, #. Boundary-layer and local friction data for Mach numbers up to 2.5 and Reynolds numbers up to 3.6 × 10 to the 8th power were obtained in flight at three locations on the XB-70-1 airplane: the lower forward fuselage centerline (nose), the upper rear fuselage centerline, and the upper surface of the right wing. Local skin friction coefficients were derived at each location by using (1) a skin friction force balance, (2) a Preston probe, and (3) an adaptation of Clauser’s method which derives skin friction from the rake velocity profile. These three techniques provided consistent results that agreed well with the von Karman-Schoenherr relationship for flow conditions that are quasi-two-dimensional. At the lower angles of attack, the nose-boom and flow-direction vanes are believed to have caused the momentum thickness at the nose to be larger than at the higher angles of attack. The boundarylayer data and local skin friction coefficients are tabulated. The wind-tunnel-model surface-pressure distribution ahead of the three locations and the flight surface-pressure distribution ahead of the wing location are included. 127 756. Tang, M. H.; and Pearson, G. P. E.: FlightMeasured X-24A Lifting Body Control Surface Hinge Moments and Correlation With Wind Tunnel Predictions. NASA TM X-2816, H-748, June 1973, 73N25049, #. Control-surface hinge-moment measurements obtained in the X-24A lifting body flight-test program are compared with results from wind-tunnel tests. The effects of variations in angle of attack, angle of sideslip, rudder bias, rudder deflection, upper-flap deflection, lower-flap deflection, Mach number, and rocket-engine operation on the control-surface hinge moments are presented. In-flight motion pictures of tufts attached to the inboard side of the right fin and the rudder and upper-flap surfaces are discussed. ECN-2006 X-24A Lifting Body 757. Rediess, Herman A.: A Survey of Parameter Identification Applications on Aircraft Flight Testing. Presented at the Joint Automatic Control Conference, Columbus, Ohio, June 20–22,1973. 758. Burcham, F. W.; Hughes, D. L.; and Holzman, J. K.: Steady-State and Dynamic Pressure Phenomena in the Propulsion System of an F-111A Airplane. NASA TN D-7328, H-741, July 1973, 73N29806, #. Flight tests were conducted with two F-111A airplanes to study the effects of steady-state and dynamic pressure phenomena on the propulsion system. Analysis of over 100 engine compressor stalls revealed that the stalls were caused by high levels of instantaneous distortion. In 73 percent of these stalls, the instantaneous circumferential distortion parameter, k sub theta, exhibited a peak just prior to stall higher than any previous peak. The K sub theta parameter was a better indicator of stall than the distortion factor, k sub d, and the maximum-minus-minimum distortion parameter, d, was poor indicator of stall. Inlet duct resonance occurred in both F-111A airplanes and is believed to have been caused by oscillations of the normal shock wave from an internal to an external position. The inlet performance of the two airplanes was similar in terms of pressure recovery, distortion, and turbulence, and there was good agreement between flight and wind-tunnel data up to a Mach number of approximately 1.8. 759. Wolowicz, C. H.; and Yancey, R. B.: Comparisons of Predictions of the XB-70-1 Longitudinal Stability and Control Derivatives With Flight Results for Six Flight Conditions. NASA TM X-2881, H-773, August 1973, 73N30940, #. Preliminary correlations of flight-determined and predicted stability and control characteristics of the XB-70-1 reported in NASA TN D-4578 were subject to uncertainties in several areas which necessitated a review of prediction techniques particularly for the longitudinal characteristics. Reevaluation and updating of the original predictions, including aeroelastic corrections, for six specific flight-test conditions resulted in improved correlations of static pitch stability with flight data. The original predictions for the pitch-damping derivative, on the other hand, showed better correlation with flight data than the updated predictions. It appears that additional study is required in the application of aeroelastic corrections to rigid model wind-tunnel data and the theoretical determination of dynamic derivatives for this class of aircraft. 760. Wilson, R. J.; Cazier, F. W., Jr.; and Larson, R. R.: Results of Ground Vibration Tests on a YF-12 Airplane. NASA TM X-2880, H-736, August 1973, 73N29944, #. Ground vibration tests were conducted on a YF-12 airplane. To approximate a structural free-free boundary condition during the tests, each of the landing gears was supported on a support system designed to have a low natural frequency. The test equipment and the procedures used for the ground vibration tests are described. The results are presented in the form of frequency response data, measured mode lines, and elastic mode shapes for the wing/body, rudder, and fuselage ventral fin. In the frequency range between 3.4 cps and 28.8 cps, nine symmetrical wing/body modes, six antisymmetrical wing/body modes, two rudder modes, and one ventral fin mode were measured. 761. Wolf, T. D.; and McCracken, R. C.: Ground and Flight Experience With a Strapdown Inertial Measuring Unit and a General Purpose Airborne Digital Computer. NASA TM X-2848, H-735, August 1973, 73N29713, #. Ground and flight tests were conducted to investigate the problems associated with using a strapdown inertial flight data system. The objectives of this investigation were to develop a three axis inertial attitude reference system, to evaluate a self-alignment technique, and to examine the problem of time-sharing a general purpose computer for the several tasks required of it. The performance of the 128 strapdown platform/computer system that was developed was sufficiently accurate for the tasks attempted. For flights on the order of 45 minutes duration, attitude angle errors of ± .035 radian (± 2 deg) in all axes were observed. Laboratory tests of the self-alignment technique gave accuracies of ±.00075 radian in pitch and roll axes and ± 0.0045 radian in the yaw axis. Self-alignment flight results were inconsistent, since a stable solution was not obtained on windy days because of aircraft rocking motions. 762. Putnam, T. W.: Investigation of Coaxial Jet Noise and Inlet Choking Using an F-111A Airplane. NASA TN D-7376, H-685, August 1973, 73N28989, #. Measurements of engine noise generated by an F-111A airplane positioned on a thrustmeasuring platform were made at angles of 0 deg to 160 deg from the aircraft heading. Sound power levels, power spectra, and directivity patterns are presented for jet exit velocities between 260 feet per second and 2400 feet per second. The test results indicate that the total acoustic power was proportional to the eighth power of the core jet velocity for core exhaust velocities greater than 300 meters per second (985 feet per second) and that little or no mixing of the core and fan streams occurred. The maximum sideline noise was most accurately predicted by using the average jet velocity for velocities above 300 meters per second (985 feet per second). The acoustic power spectrum was essentially the same for the single jet flow of afterburner operation and the coaxial flow of the nonafterburning condition. By varying the inlet geometry and cowl position, reductions in the sound pressure level of the blade passing frequency on the order of 15 decibels to 25 decibels were observed for inlet Mach numbers of 0.8 to 0.9. 763. Monaghan, R. C.; and Fields, R. A.: Experiments to Study Strain Gage Load Calibrations on a Wing Structure at Elevated Temperatures. NASA TN D-7390, H-763, August 1973, 73N28883, #. Laboratory experiments were performed to study changes in strain-gage bridge load calibrations on a wing structure heated to temperatures of 200 F, 400 F, and 600 F. Data were also obtained to define the experimental repeatability of strain-gage bridge outputs. Experiments were conducted to establish the validity of the superposition of bridge outputs due to thermal and mechanical loads during a heating simulation of Mach 3 flight. The strain-gage bridge outputs due to load cycle at each of the above temperature levels were very repeatable. A number of bridge calibrations were found to change significantly as a function of temperature. The sum of strain-gage bridge outputs due to individually applied thermal and mechanical loads compared well with that due to combined or superimposed loads. The validity of superposition was, therefore, established. 764. Monaghan, R. C.; and Friend, E. L.: Effects of Flaps on Buffet Characteristics and Wing-Rock Onset of an F-8C Airplane at Subsonic and Transonic Speeds. NASA TM X-2873, H-742, August 1973, 73N27905, #. Wind-up-turn maneuvers were performed to establish the values of airplane normal force coefficient for buffet onset, wing-rock onset, and buffet loads with various combinations of leading- and trailing-edge flap deflections. Data were gathered at both subsonic and transonic speeds covering a range from Mach 0.64 to Mach 0.92. Buffet onset and buffet loads were obtained from wingtip acceleration and wing-root bending-moment data, and wing-rock onset was obtained from airplane roll rate data. Buffet onset, wing-rock onset, and buffet loads were similarly affected by the various combinations of leading- and training-edge flaps. Subsonically, the 12 deg leading-edge-flap and trailing-edgeflap combination was most effective in delaying buffet onset, wing-rock onset, and equivalent values of buffet loads to a higher value of airplane normal force coefficient. This was the maximum flap deflection investigated. Transonically, however, the optimum leading-edge flap position was generally less than 12 deg. 765. Gee, S. W.; and McCracken, R. C.: Preliminary Flight Evaluation of a Painted Diamond on a Runway for Visual Indication of Glide Slope. NASA TM X-2849, H-739, August 1973, 73N27027, #. A diamond sized to appear equidimensional when viewed from a 3.6 deg slide slope was painted on the end of a small general aviation airport runway, and a series of flights was made to evaluate its usage as a piloting aid. The pilots could detect and fly reasonably close to the glide slope projected by the diamond. The flight path oscillations that were recorded during approaches using the diamond were not significantly different from the oscillations that were recorded without the diamond; the difference that did exist could be attributed to converging on a known projected glide slope in one case, and flying an unknown, random glide slope in the other. The results indicated that the diamond would be effective as a means of intercepting and controlling a predetermined glide slope. Other advantages of the diamond were positive runway identification and greater aim point visibility. The major disadvantage was a tendency to overconcentrate on the diamond and consequently to neglect cockpit instruments and airport traffic. 766. Berry, D. T.; and Gilyard, G. B.: Airframe/ Propulsion System Interactions - An Important Factor in Supersonic Aircraft Flight Control. AIAA Paper 73-831. Presented at the AIAA Guidance and Control Conference, Key Biscayne, Florida, August 20–22, 1973, 73A40501, #. The demands of supersonic flight have resulted in propulsion system features that have a significant influence on aircraft flight control. Data from a Mach 3 cruise airplane show that airframe/propulsion system interactions can reduce phugoid and dutch-roll damping, increase vehicle sensitivity to atmospheric disturbances, alter the effective static and dynamic stability of the aircraft, and produce moments as strong as aerodynamic controls. In turn, these effects can lead to large aircraft excursions or high pilot workload, or both, and place increased demands on stability augmentation systems and aerodynamic controls. A need to integrate flight control and propulsion control in advanced vehicles is indicated. 767. *Johnson, W. A.; and Rediess, H. A.: Study of Control System Effectiveness in Alleviating Vortex Wake Upsets. AIAA Paper 73-833. Presented at the AIAA Guidance and Control Conference, Key Biscayne, Florida, August 20–22, 1973, 73A38776. The problem of an airplane being upset by encountering the vortex wake of a large transport on takeoff or landing is currently receiving considerable attention. This paper describes the technique and results of a study to assess the effectiveness of automatic control systems in alleviating vortex wake upsets. A six-degree-of-freedom nonlinear digital simulation was used for this purpose. The analysis included establishing the disturbance input due to penetrating a vortex wake from an arbitrary position and angle. Simulations were computed for both a general aviation airplane and a commercial jet transport. Dynamic responses were obtained for the penetrating aircraft with no augmentation and with various command augmentation systems. The results of this preliminary study indicate that it is feasible to use an automatic control system to alleviate vortex encounter upsets. *Systems Technology, Inc., Hawthorne, California. 768. Wolowicz, C. H.; and Yancey, R. B.: Summary of Stability and Control Characteristics of the XB-70 Airplane. NASA TM X-2933, H-781, October 1973, 73N31958, #. The stability and control characteristics of the XB-70 airplane were evaluated for Mach numbers up to 3.0 and altitudes up to 21,300 meters (70,000 feet). The airplane’s inherent longitudinal characteristics proved to be generally satisfactory. In the lateral-directional modes, the airplane was characterized by light wheel forces, low static directional stability beyond approximately 2 deg of sideslip, adverse yaw response to aileron inputs throughout the entire Mach number range, and negative effective dihedral with wingtips full down. At subsonic Mach numbers, with the flight augmentation control system off, the light wheel forces and adverse yaw response to aileron inputs caused the pilots to minimize use of the ailerons. At supersonic Mach numbers, with the augmentation system off, the adverse yaw due to 129 aileron and the negative effective dihedral were conducive to pilot-induced oscillations. 769. Armistead, K. H.; and Webb, L. D.: Flight Calibration Tests of a Nose-Boom-Mounted Fixed Hemispherical Flow-Direction Sensor. NASA TN D-7461, H-779, October 1973, 73N31956, #. Flight calibrations of a fixed hemispherical flow angle-ofattack and angle-of-sideslip sensor were made from Mach numbers of 0.5 to 1.8. Maneuvers were performed by an F-104 airplane at selected altitudes to hemispherical sensor with that from a standard angle-of-attack vane. The hemispherical flow-direction sensor measured differential pressure at two angle-of-attack ports and two angle-ofsideslip ports in diametrically opposed positions. Stagnation pressure was measured at a center port. The results of these tests showed that the calibration curves for the hemispherical flow-direction sensor were linear for angles of attack up to 13 deg. The overall uncertainty in determining angle of attack from these curves was plus or minus 0.35 deg or less. A Mach number position error calibration curve was also obtained for the hemispherical flow-direction sensor. The hemispherical flow-direction sensor exhibited a much larger position error than a standard uncompensated pitot-static probe. 770. Powers, B. G.; and Kier, D. A.: Simulator Evaluation of the Low-Speed Flying Qualities of an Experimental STOL Configuration With an Externally Blown Flap Wing on an Augmentor Wing. NASA TN D-7454, H-780, October 1973, 73N31951, #. The low-speed flying qualities of an experimental STOL configuration were evaluated by using a fixed-base sixdegree-of-freedom simulation. The configuration had either an externally blown flap (EBF) wing or an augmentor wing (AW). The AW configuration was investigated with two tails, one sized for the AW configuration and a larger one sized for the EBF configuration. The emphasis of the study was on the 70-knot approach task. The stability and control characteristics were compared with existing criteria. Several control systems were investigated for the normal four-engine condition and for the engine-out transient condition. Minimum control and stall speeds were determined for both the three- and four-engine operation. 771. Lock, W. P.; Kordes, E. E.; McKay, J. M.; and *Wykes, J. H.: Flight Investigation of a Structural Mode Control System for the XB-70 Aircraft. NASA TN D-7420, H-732, October 1973, 73N31950, #. A flight investigation of a structural mode control system termed identical location of accelerometer and force (ILAF) was conducted on the XB-70-1 airplane. During the first flight tests, the ILAF system encountered localized structural vibration problems requiring a revision of the compensating network. After modification, successful structural mode control that did not adversely affect the rigid body dynamics was demonstrated. The ILAF system was generally more effective in supersonic than subsonic flight, because the conditions for which the system was designed were more nearly satisfied at supersonic speeds. The results of a turbulence encounter at a Mach number of 1.20 and an altitude of 9754 meters indicated that the ILAF system was effective in reducing the vehicle’s response at this flight condition. An analytical study showed that the addition of a small canard to the modal suppression system would greatly improve the automatic control of the higher frequency symmetric modes. *North American Rockwell Corp., Los Angeles, California. 772. Washington, H. P.; and Gibbons, J. T.: Analytical Study of Takeoff and Landing Performance for a Jet STOL Transport Configuration With Full-Span, Externally Blown, Triple-Slotted Flaps. NASA TN D-7441, H-709, October 1973, 73N31939, #. Takeoff and landing performance characteristics and field length requirements were determined analytically for a jet STOL transport configuration with full-span, externally blown, tripleslotted flaps. The configuration had a high wing, high T-tail, and four pod-mounted high-bypass-ratio turbofan engines located under and forward of the wing. One takeoff and three approach and landing flap settings were evaluated. The effects of wing loading, thrust-to-weight ratio, weight, ambient temperature, altitude on takeoff and landing field length requirements are discussed. 773. Sim, A. G.: Results of a Feasibility Study Using the Newton-Raphson Digital Computer Program to Identify Lifting Body Derivatives From Flight Data. NASA TM X-56017, October 1973, 74N11814, #. A brief study was made to assess the applicability of the Newton-Raphson digital computer program as a routine technique for extracting aerodynamic derivatives from flight tests of lifting body types of vehicles. Lateral-direction flight data from flight tests of the HL-10 lifting body research vehicle were utilized. The results in general, show the computer program to be a reliable and expedient means for extracting derivatives for this class of vehicles as a standard procedure. This result was true even when stability augmentation was used. As a result of the study, a credible set of HL-10 lateral-directional derivatives was obtained from flight data. These derivatives are compared with results from wind-tunnel tests. 774. Pyle, J. S.; and Saltzman, E. J.: Review of Drag Measurements From Flight Tests of Manned Aircraft 130 With Comparisons to Wind-Tunnel Predictions. AGARD-CP-124, Paper 26. AGARD Aerodyn. Drag, October 1973, pp. 25-1 to 25-12, 74N14735. In-flight studies of the overall and local components of drag of many types of aircraft were conducted. The primary goal of these studies was to evaluate wind-tunnel and semiempirical prediction methods. Some evaluations are presented in this paper which may be summarized by the following observations: Wind-tunnel predictions of overall vehicle drag can be accurately extrapolated to flight Reynolds numbers, provided that the base drag is removed and the boattail areas on the vehicle are small. The addition of ablated roughness to lifting body configurations causes larger losses in performance and stability than would be expected from the added friction drag due to the roughness. Successful measurements of skin friction have been made in flight to Mach numbers above 4. A reliable inflatable deceleration device was demonstrated in flight which effectively stabilizes and decelerates a lifting aircraft at supersonic speeds. 775. Borek, R. W.: Development of AIFTDS-4000, a Flight-Qualified, Flexible, High-Speed Data Acquisition System. NASA TM X-56018. International Telemetering Conference, Washington, D.C., October 9–11, 1973, 73N32087, #. The NASA flight research center has developed a prototype data acquisition system which integrates an airborne computer with a high-speed pulse code modulation system. The design of the airborne integrated flight test data system (AIFTDS) is the result of experience with airborne pulse code modulation data systems. The AIFTDS-4000 has proved the premise on which it was designed: that the needs and requirements of data acquisition system users can be integrated to produce a highly flexible system that will be more useful than existing systems. 776. Berry, D. T.; and Gilyard, G. B.: Some Stability and Control Aspects of Airframe/Propulsion System Interactions on the YF-12 Airplane. ASME Paper 73-WA/ AERO-4. Presented at the American Society of Mechanical Engineers, Winter Annual Meeting, Detroit, Michigan, November 11–15, 1973, 74A13246, #. Airframe/propulsion system interactions can strongly affect the stability and control of supersonic cruise aircraft. These interactions generate forces and moments similar in magnitude to those produced by the aerodynamic controls, and can cause significant changes in vehicle damping and static stability. This in turn can lead to large aircraft excursions or high pilot workload, or both. For optimum integration of an airframe and its jet propulsion system, these phenomena may have to be taken into account. 777. Lewis, C. E., Jr.; Swaroop, R.; McMurtry, T. C.; *Blakeley, W. R.; and **Masters, R. L.: Landing Performance by Low-Time Private Pilots After the Sudden Loss of Binocular Vision — Cyclops II. Aerospace Medicine, Vol. 44, No. 11, November 1973, pp. 1241–1245, 74A13530. Study of low-time general aviation pilots, who, in a series of spot landings, were suddenly deprived of binocular vision by patching either eye on the downwind leg of a standard, closed traffic pattern. Data collected during these landings were compared with control data from landings flown with normal vision during the same flight. The sequence of patching and the mix of control and monocular landings were randomized to minimize the effect of learning. No decrease in performance was observed during landings with vision restricted to one eye, in fact, performance improved. This observation is reported at a high level of confidence (p less than 0.001). These findings confirm the previous work of Lewis and Krier and have important implications with regard to aeromedical certification standards. *New Mexico University, Albuquerque, New Mexico. **Lovelace Foundation For Medical Education and Research, Albuquerque, New Mexico. 778. Burcham, F. W., Jr.; Holzman, J. K.; and Reukauf, P. J.: Preliminary Results of Flight Tests of the Propulsion System of the YF-12 Airplane at Mach Numbers to 3.0. AIAA Paper 73-1314. Presented at the AIAA and 9th Society of Automotive Engineers Propulsion Conference, Las Vegas, Nevada, November 5–7, 1973, 74A12951, #. Flight tests of the propulsion system of a YF-12 airplane were made which included off-schedule inlet operation and deliberately induced unstarts and compressor stalls. The tests showed inlet/engine compatibility to be good through most of the flight envelope. The position of the terminal shock wave could be determined from throat static pressure profiles or from root-mean-square levels of throat static pressure fluctuations. A digital simulation of the control system showed an oscillation of the forward bypass doors to be caused by hysteresis in the bypass door actuator linkages. 779. *Webb, W. L.; and Reukauf, P. J.: Development of a Turbine Inlet Gas Temperature Measurement and Control System Using a Fluidic Temperature Sensor. AIAA Paper 73-1251. Presented at the 9th AIAA and Society of Automotive Engineers, Propulsion Conference, Las Vegas, Nevada, November 5–7, 1973, 74A11272, #. (See also 811.) *United Aircraft Corp., Pratt and Whitney Aircraft Division, West Palm Beach, Florida. 131 780. Pyle, J. S.; Phelps, J. R.; and Baron, R. S.: Performance of a Ballute Decelerator Towed Behind a Jet Airplane. NASA TM X-56019, H-815, December 1973, 74N14760, #. An F-104B airplane was modified to investigate the drag and stability characteristics of a ballute decelerator in the wake of an asymmetrical airplane. Decelerator deployments were initiated at a Mach number of 1.3 and an altitude of 15,240 meters (50,000 feet) and terminated when the airplane had decelerated to a Mach number of 0.5. The flight tests indicated that the decelerator had a short inflation time with relatively small opening forces. The drag levels attained with the subject decelerator were less than those obtained with other high-speed decelerators behind a symmetrical tow vehicle. The ballute demonstrated good stability characteristics behind the testbed airplane. 781. Sim, A. G.: Flight-Determined Stability and Control Characteristics of the M2-F3 Lifting Body Vehicle. NASA TN D-7511, H-791, December 1973, 74N12534, #. Flight data were obtained over a Mach number range from 0.4 to 1.55 and an angle-of-attack range from –2 deg to 16 deg. Lateral-directional and longitudinal derivatives, reaction control rocket effectiveness, and longitudinal trim information obtained from flight data and wind-tunnel predictions are compared. The effects of power, configuration change, and speed brake are discussed. 782. McMurtry, T. C.; Gee, S. W.; and Barber, M. R.: A Flight Evaluation of Curved Landing Approaches. Society of Experimental Test Pilots, Technical Review, Vol. 11, No. 3, 1973, pp. 5–17, 73A28901. A potential solution to some of the operational problems of STOL aircraft operations in the terminal area lies in the capability of making curved landing approaches under both visual and instrument flight conditions. Tests are described which were conducted with a twin-engine, light weight, general aviation aircraft. The advanced control system mode utilized during the curved approaches was an attitude command control system. Four curved patterns were investigated using a steep glide slope: two display configurations, and two flight control modes. When using the flight director display, curved approaches were not significantly different in difficulty and work load than straight approaches. 783. Peterson, B. A.; Krier, G. E.; and *Jarvis, C. R.: Development and Flight Test of a Digital Fly-By-Wire F-8 Airplane. Society of Experimental Test Pilots, Technical Review, Vol. 11, No. 2, 1973, pp. 57–71, 73A22180. The objective of the F-8 digital fly-by-wire program is to establish a technology base for the implementation of advanced flight control systems. The central element is the Apollo Lunar Guidance Computer (LGC). This versatile computer ran over 2000 hours in support of fly-by-wire without a failure. Difficulties encountered in the first flights were corrected rapidly and simply by changes in the erasable software memory. Control-configured vehicles offer significant weight-saving possibilities. * NASA, Washington, D.C. 784. Iliff, K. W.: Identification and Stochastic Control With Application to Flight Control in Turbulence. UCLAENG-7340 (Ph.D. dissertation), 1973, 74N17383. The problem is dealt with of adaptive control of an aircraft in atmospheric turbulence. The problem is approached by first identifying the unknown coefficients and then applying optimal control theory to the system so determined. The theory developed is general enough to apply to any linear system with unknown coefficients and state noise. The bulk of the development concerns the identification problem and several methods are studied. In particular, what may be called stochastic identification method, taking into account the unknown state noise, is studied. The identification and control theory is first verified on simulated data. It is shown that the methods that accounted for the state noise are adequate where the assumptions hold. The optimal control results agree well with theory in achieving the desired minimization. 132 980549 M2-F3 Lifting Body, Three-View Drawing 1974 Technical Publications 785. Montoya, L. C.; Brauns, D. A.; and Cissell, R. E.: Flight Experience With a Pivoting Traversing BoundaryLayer Probe. NASA TM X-56022, January 1974, 74N16102, #. A pivoting traversing boundary layer probe was evaluated in flight on an F-104 airplane. The evaluation was performed at free stream Mach numbers from 0.8 to 2.0. The unit is described, and operating problems and their solutions are discussed. Conventional boundary layer profiles containing variations in flow angle within the viscous layer are shown for free stream Mach numbers of 0.8, 1.6, and 2.0. Although the unit was not optimized for size and weight, it successfully measured simultaneously flow angularity, probe height, and pitot pressure through the boundary layer. 786. Painter, W. D.; and Sitterle, G. J.: HL-10 Lifting Body Flight Control System Characteristics and Operational Experience. NASA TM X-2956, H-704, January 1974, 74N14753, #. A flight evaluation was made of the mechanical hydraulic flight control system and the electrohydraulic stability augmentation system installed in the HL-10 lifting body research vehicle. Flight tests performed in the speed range from landing to a Mach number of 1.86 and the altitude range from 697 meters (2300 feet) to 27,550 meters (90,300 feet) were supplemented by ground tests to identify and correct structural resonance and limit-cycle problems. Severe limitcycle and control sensitivity problems were encountered during the first flight. Stability augmentation system structural resonance electronic filters were modified to correct the limit-cycle problem. Several changes were made to control stick gearing to solve the control sensitivity problem. Satisfactory controllability was achieved by using a nonlinear system. A limit-cycle problem due to hydraulic fluid contamination was encountered during the first powered flight, but the problem did not recur after preflight operations were improved. 787. Kempel, R. W.; and Manke, J. A.: Flight Evaluation of HL-10 Lifting Body Handling Qualities at Mach Numbers From 0.30 to 1.86. NASA TN D-7537, H-757, January 1974, 74N14535, #. The longitudinal and lateral-directional handling qualities of the HL-10 lifting body vehicle were evaluated in flight at Mach numbers up to 1.86 and altitudes up to approximately 27,450 meters (90,000 feet). In general, the vehicle’s handling qualities were considered to be good. Approximately 91 percent of the pilot ratings were 3.5 or better, and 42.4 percent were 2.0. Handling qualities problems were encountered during the first flight due to problems with the control system and vehicle aerodynamics. Modifications of the flight vehicle corrected all deficiencies, 133 and no other significant handling qualities problems were encountered. 788. Love, J. E.; *Fox, W. J.; and **Wicklund, E. J.: Flight Study of a Vehicle Operational Status and Monitoring System. NASA TN D-7546, H-789, January 1974, 74N13725, #. An analog onboard monitoring system was installed on a YF-12 airplane as the first phase of a program to monitor the engine inlet and portions of the airplane’s electrical and fuel management subsystems in flight. The system provided data which were considered to form a suitable base for diagnostic test logic and decision criteria for the rest of the program. The data were also adequate for the purpose of maintaining the engine inlet and identifying malfunctions within it. The investigation showed that the requirements of an onboard monitoring system should be considered during the original design of the system to be monitored. *Lockheed-California Co., Burbank, California. **Honeywell, Inc., Minneapolis, Minnesota. 789. Putnan, T. W.: Flight Experience With the Decelerating Noise Abatement Approach. NASA TM X-56020, January 1974, 74N12720, #. The noise of older aircraft can be reduced in two principal ways: retrofitting the aircraft with a quiet propulsion system, and changing the flight operational procedures used in flying the aircraft. The former approach has already proved to be expensive, time consuming, and difficult to implement even though low-noise propulsion system technology exists. The latter method seems to hold promise of being less expensive and easier to implement. One operational technique which might reduce the noise beneath the landing approach path is the decelerating approach. This technique requires intercepting the 3 deg approach path at a relatively high speed with the aircraft in the cruise configuration, then reducing the thrust to idle and allowing the aircraft to decelerate along the 3 deg approach path. As the appropriate airspeed is achieved, the landing flaps and landing gear are deployed for a normal flare and landing. Because the engines, which are the predominant noise source on landing approach, are at idle thrust, a significant reduction in the noise beneath the approach path should be realized. 790. Jarvis, C. R.: A Digital Fly-By-Wire Technology Development Program Using an F-8C Test Aircraft. AIAA Paper 74-28, Twelfth AIAA Aerospace Sciences Meeting, Washington, D.C., January 30–February 1, 1974, pp. 11, January 1974, 74A20755, #. A digital fly-by-wire flight control system has been installed in an F-8C test airplane and has undergone extensive ground and flight testing as part of an overall program to develop digital fly-by-wire technology. This is the first airplane to fly with a digital fly-by-wire system as its primary means of control and with no mechanical reversion capability. Fortytwo test flights were made for a total flight time of 57 hours. Six pilots participated in the evaluation. This paper presents an overview of the digital fly-by-wire program and discusses some of the flight-test results. 791. Taillon, N. V.: Flight-Test Investigation of the Aerodynamic Characteristics and Flow Interference Effects About the Aft Fuselage of the F-111A Airplane. NASA TN D-7563, H-717, February 1974, 74N18657, #. Static pressure measurements were made on the aft fuselage of an F-111A airplane to determine local flow characteristics and engine/airframe interaction effects. Data were obtained over the Mach number range from 0.5 to 2.0. Aspiration effects associated with low ejector nozzle expansion ratios reduced the local pressure coefficients particularly on the interfairing but also extending to the trailing edge of the nacelle. The presence of afterbodies also affected the behavior of the air flowing into and about the ejector nozzle. Pressures about the aft fuselage were improved by an increase in primary nozzle area at a supersonic speed. A comparison of wind-tunnel and flight-test results showed generally good agreement, although there was a large disparity in pressure level about the ejector nozzle. However, the shape of the data curves and the local flow behavior were basically similar. 792. Larson, T. J.; and Schweikhard, W. G.: A Simplified Flight-Test Method for Determining Aircraft Takeoff Performance That Includes Effects of Pilot Technique. NASA TN D-7603, H-802, February 1974, 74N16717, #. A method for evaluating aircraft takeoff performance from brake release to air-phase height that requires fewer tests than conventionally required is evaluated with data for the XB-70 airplane. The method defines the effects of pilot technique on takeoff performance quantitatively, including the decrease in acceleration from drag due to lift. For a given takeoff weight and throttle setting, a single takeoff provides enough data to establish a standardizing relationship for the distance from brake release to any point where velocity is appropriate to rotation. The lower rotation rates penalized takeoff performance in terms of ground roll distance; the lowest observed rotation rate required a ground roll distance that was 19 percent longer than the highest. Rotations at the minimum rate also resulted in lift-off velocities that were approximately 5 knots lower than the highest rotation rate at any given liftoff distance. 793. Saltzman, E. J.; and Meyer, R. R., Jr.: Drag Reduction Obtained by Rounding Vertical Corners on a Box-Shaped Ground Vehicle. NASA TM X-56023, March 1974, 74N17703, #. A box-shaped ground vehicle was used to simulate the aerodynamic drag of delivery vans, trucks, and motor homes. 134 A coast-down method was used to define the drag of this vehicle in a configuration with all square corners and a modified configuration with the four vertical corners rounded. The tests ranged in velocity from 30 miles per hour to 65 miles per hour, and Reynolds numbers ranged from 4.4 × 1,000,000 to 1.0 × 10 to the 7th power based on vehicle length. The modified configuration showed a reduction in aerodynamic drag of about 40 percent as compared to the square cornered configuration. 794. Holleman, E. C.: Initial Results From Flight Testing a Large, Remotely Piloted Airplane Model. NASA TM X-56024, March 1974, 74N18671, #. The first four flights of a remotely piloted airplane model showed that a flight envelope can be expanded rapidly and that hazardous flight tests can be conducted safely with good results. The flights also showed that aerodynamic data can be obtained quickly and effectively over a wide range of flight conditions, clear and useful impressions of handling and controllability of configurations can be obtained, and present computer and electronic technology provide the capability to close flight control loops on the ground, thus providing a new method of design and flight test for advanced aircraft. 795. Anon.: Parameter Estimation Techniques and Applications in Aircraft Flight Testing. NASA TN D-7647, H-806, Aircraft Symposium, Edwards, California, April 24–25, 1973, 74N25569, #. Technical papers were presented by selected representatives from industry, universities, and various Air Force, Navy, and NASA installations. The topics covered included the newest developments in identification techniques, the most recent flight-test experience, and the projected potential for the near future. 796. Rediess, H. A.: An Overview of Parameter Estimation Techniques and Applications in Aircraft Flight Testing. NASA TN D-7647, H-806. Parameter Estimation Tech. and Appl. in Aircraft Flight Testing, April 1974, pp. 1–18, (see N74-25569 15-02), 74N25570. Parameter estimation is discussed as it applies to aircraft flight testing, and an overview of the symposium is presented. The evolution of techniques used in flight testing is reviewed briefly, and it is pointed out how the changing character of the aircraft tested and the availability of advanced data systems have promoted this evolution. Recent advances in optimal estimation theory have stimulated widespread interest and activity in parameter estimation. The framework of these advanced techniques is outlined to set the stage for subsequent papers. The session topics are introduced and related to the requirements of flight-test research. 797. Iliff, K. W.: Identification of Aircraft Stability and Control Derivatives in the Presence of Turbulence. NASA TN D-7647, H-806. Parameter Estimation Tech. and Appl. in Aircraft Flight Testing, April 1974, pp. 77–114, (see N74-25569 15-02), 74N25575. A maximum likelihood estimator for a linear system with state and observation noise is developed to determine stability and control derivatives from flight data obtained in the presence of turbulence. The formulation for the longitudinal short-period mode is presented briefly, including a special case that greatly simplifies the problem if the measurement noise on one signal is negligible. The effectiveness and accuracy of the technique are assessed by applying it first to simulated flight data, in which the true parameter values and state noise are known, then to actual flight data obtained in turbulence. The results are compared with data obtained in smooth air and with wind-tunnel data. The complete maximum likelihood estimator, which accounts for both state and observation noise, is shown to give the most accurate estimate of the stability and control derivatives from flight data obtained in turbulence. It is superior to the techniques that ignore state noise and to the simplified method that neglects the measurement noise on the angle-of-attack signal. 798. Gilyard, G. B.: Determination of PropulsionSystem-Induced Forces and Moments of a Mach 3 Cruise Aircraft. NASA TN D-7647, H-806. Parameter Estimation Tech. and Appl. in Aircraft Flight Testing, April 1974, pp. 369–374, (see N74-25569 15-02), 74N25591. During the joint NASA/USAF flight research program with the YF-12 airplane, the Dutch roll damping was found to be much less during automatic inlet operation than during fixed inlet operation at Mach numbers greater than 2.5 and with the yaw stability augmentation system off. It was concluded that the significant reduction in dutch roll damping was due to the forces and moments induced by the variable-geometry features of the inlet. Two stability-derivative extraction techniques were applied to the flight data; the recently developed Newton-Raphson technique and the time vector method. These techniques made it possible to determine the forces and moments generated by spike and bypass door movement. 799. Matheny, N. W.: Flight Investigation of Approach and Flare From Simulated Breakout Altitude of a Subsonic Jet Transport and Comparison With Analytical Models. NASA TN D-7645, H-803, April 1974, 74N19672, #. Satisfactory and optimum flare windows are defined from pilot ratings and comments. Maximum flare normal accelerations, touchdown rates of sink, and total landing maneuver time increments are summarized as a function of approach airspeed margin (with respect to reference airspeed) 135 and flare initiation altitude. The effects of two thrust management techniques are investigated. Comparisons are made with predictions from three analytical models and the results of a simulator study. The approach speed margin was found to have a greater influence on the flare initiation altitude than the absolute airspeed. The optimum airspeed was between the reference airspeed and the reference airspeed plus 10 knots. The optimum flare initiation altitude range for unrestricted landings was from 11 meters to 20 meters (36 feet to 66 feet), and the landing time in the optimum window was 8 seconds. The duration of the landing maneuver increased with increasing flare initiation altitude and with increasing speed margins on the approach. 800. Weirather, L. H.: Transducers. AGARDOGRAPH160-VOL-1. Flight Test Instrumentation Ser., Vol. 1, April 1974, (see N74-25933 15-14), 74N25937. The use of transducers in the measuring channels of flight test instrumentation systems is discussed. Emphasis is placed on transducers with an electrical output. The physical effects used for producing the electrical outputs are defined. Diagrams of the various types of transducers are included to show the operating principles. 801. Reed, R. D.: RPRVs —The First and Future Flights; Remotely Piloted Research Vehicle. Astronautics and Aeronautics, Vol. 12, April 1974, pp. 26–42, 74A26410, #. The merits of the RPRV (remotely piloted research vehicle) concept are discussed, along with its historical background and development culmination in the 3/8-scale F-15. The use of RPRVs is shown to be especially attractive when testing must be done at low cost, or in quick response to demand, or when hazardous testing must assure the safety of proceeding to manned vehicles. F-15 Spin Research Vehicle ECN-4891 802. Schweikhard, W. G.; and Berry, D. T.: Cooperative Airframe/Propulsion Control for Supersonic Cruise Aircraft. SAE Paper 740478, Society of Automotive Engineers, Air Transportation Meeting, Dallas, Texas, April 30–May 2, 1974, 74A34998. Interactions between propulsion systems and flight controls have emerged as a major control problem on supersonic cruise aircraft. This paper describes the nature and causes of these interactions and the approaches to predicting and solving the problem. Integration of propulsion and flight control systems appears to be the most promising solution if the interaction effects can be adequately predicted early in the vehicle design. Significant performance, stability, and control improvements may be realized from a cooperative control system. 803. Deets, D. A.; and Szalai, K. J.: Design and Flight Experience With a Digital Fly-By-Wire Control System in an F-8 Airplane. AGARD CP-137. Advances in Control Systems, May 1974, (see N74-31429 21-02), 74N31450. A digital fly-by-wire flight control system was designed, built, and for the first time flown in an airplane. The system, which uses components from the Apollo guidance system, is installed in an F-8 airplane as the primary control system. A lunar module guidance computer is the central element in the three-axis, single-channel, multimode, digital control system. A triplex electrical analog system which provides unaugmented control of the airplane is the only backup to the digital system. Flight results showed highly successful system operation, although the trim update rate was inadequate for precise trim changes, causing minor concern. The use of a digital system to implement conventional control laws proved to be practical for flight. Logic functions coded as an integral part of the control laws were found to be advantageous. Although software verification required extensive effort, confidence in the software was achieved. 804. Smith, R. H.; and Bauer, C. A.: Atmospheric Effects on the Inlet System of the YF-12 Aircraft. Proceedings, Eleventh National Conference on Environmental Effects on Aircraft and Propulsion Systems, Trenton, New Jersey, U.S. Naval Air Propulsion Test Center, May 21–23, 1974, 1974, 74A39744, #. Flights of a YF-12 airplane were performed over a wide range of operating conditions so that detailed comparisons could be made with data from tests on scale models in NASA ground facilities. Extensive flight instrumentation for inlet performance comparisons provided flight data that also lend insight into supersonic inlet operation during atmospheric turbulence. Pressure and flow direction measurements near the inlet gave results different from conventional accelerometer data normally used for flight determination of turbulence severity. A nonturbulent atmospheric temperature excursion during an XB-70 flight caused inlet duct pressure variations as extreme as those experienced during heavy turbulence on the YF-12 airplane. 805. Loschke, P. C.; Barber, M. R.; Enevoldson, E. K.; and McMurtry, T. C.: Flight Evaluation of Advanced Control Systems and Displays on a General Aviation Airplane. NASA TN D-7703, H-783, June 1974, 74N27499, #. A flight-test program was conducted to determine the effect of advanced flight control systems and displays on the handling qualities of a light twin-engined airplane. A flightdirector display and an attitude-command control system, used separately and in combination, transformed a vehicle with poor handling qualities during ILS approaches in turbulent air into a vehicle with good handling qualities. The attitude-command control system also improved the ride qualities of the airplane. A rate-command control system made only small improvements to the airplane’s ILS handling qualities in turbulence. Both the rate- and the attitudecommand control systems reduced stall warning in the test airplane, increasing the likelihood of inadvertent stalls. The final approach to the point of flare was improved by both the rate- and the attitude-command control systems. However, the small control wheel deflections necessary to flare were unnatural and tended to cause overcontrolling during flare. Airplane handling qualities are summarized for each controlsystem and display configuration. 806. Nugent, J.; and Holzman, J. K.: Flight-Measured Inlet Pressure Transients Accompanying Engine Compressor Surges on the F-111A Airplane. NASA TN D-7696, H-804, June 1974, 74N26251, #. Two F-111A airplanes were subjected to conditions that caused engine compressor surges and accompanying duct hammershock pressure transients. Flight speed ranged from Mach 0.71 to Mach 2.23, and altitude varied from approximately 3200 meters to 14,500 meters. A wide range of compressor pressure ratios was covered. Stabilized freestream, engine, and duct conditions were established before each compressor surge. Dynamic pressure instrumentation at the compressor face and in the duct recorded the pressure transients associated with the surges. Hammershock pressures were analyzed with respect to the stabilized conditions preceding the compressor surges. The hammershock transients caused large pressure rises at the compressor face and in the duct. Hammershock pressure ratios at the compressor face were not affected by free-stream Mach number or altitude but were functions of engine variables, such as compressor pressure ratio. The maximum hammershock pressure ratio of approximately 1.83 occurred at a compressor pressure ratio of approximately 21.7. 807. Lasagna, P. L.; and Putnam, T. W.: Preliminary Measurements of Aircraft Aerodynamic Noise. AIAA Paper 74-572, Seventh AIAA Fluid and Plasma Dynamics 136 Conference, Palo Alto, California, June 17–19, 1974, 74A34332, #. Flight measurements of aerodynamic noise were made on an AeroCommander airplane with engines off and a JetStar airplane with engines at both idle power and completely shut off. The overall sound level for these airplanes in the landing configuration varied as the sixth power of the aircraft velocity. For the JetStar airplane, the overall sound level decreased as the inverse square of the distance in the lateral direction. The aerodynamic noise was approximately 11 decibels below the FAR Part 36 noise level for the JetStar airplane. The landing gear were a significant contributor to aerodynamic noise for both aircraft. technology programs, such as fly-by-wire, digital control, and control configured vehicles; important applications of advanced control systems to such vehicles as the space shuttle orbiter, the Lockheed C-5A, and the Boeing 747; advanced and integrated propulsion control systems; and case studies of the benefits of applying active control technology to transport aircraft. Also included are several papers on the design, testing, and reliability of advanced control systems directed primarily toward the technical specialist. 810. Rediess, Herman A.; Kordes, Eldon E.; and Edwards, John W.: A Remotely Augmented Approach to Flight Testing of Advanced Control Technology. NASA Advanced Control Technology Symposium, Los Angeles, California, July 9–11, 1974. 811. *Webb, W. L.; and Reukauf, P. J.: Development of a Turbine Inlet Gas Temperature Measurement and Control System Using a Fluidic Temperature Sensor. Journal of Aircraft, Vol. 11, No. 7, July 1974, pp. 422–427. (See also 779.) A fluidic turbine engine gas temperature measurement and control system was developed for use on a Pratt & Whitney Aircraft J58 engine. This paper includes the criteria used for material selection, system design, and system performance. It was found that the fluidic temperature sensor had the durability for flight test under conditions existing in the YF-12 airplane. As a result of turbine inlet gas temperatures fluctuations, over-all engine-control system performance cannot be adequately evaluated without a multiple-gas sampling system. *United Aircraft Corp., Pratt and Whitney Aircraft Div., West Palm Beach, Florida. 812. Gilyard, G. B.; Smith, J. W.; and *Falkner, V. L.: Flight Evaluation of a Mach 3 Cruise Longitudinal Autopilot. AIAA Paper 74-910, AIAA Mechanics and Control of Flight Conference, Anaheim, California, August 5–9, 1974, 74A37890, #. At high Mach numbers (approximately 3) and altitudes greater than 70,000 feet, the original altitude and Mach hold modes of the YF-12 autopilot produced aircraft excursions that were erratic or divergent or both. Data analysis and simulator studies showed that static pressure port sensitivity to angle of attack had a detrimental effect on the performance of both altitude and Mach hold modes. Good altitude hold performance was obtained when a high-passed pitch rate feedback was added to compensate for angle-of-attack sensitivity and the altitude error and integral altitude gains were reduced. Good Mach hold performance was obtained with the removal of angle-of-attack sensitivity. *Honeywell, Inc., Minneapolis, Minnesota. C-140 JetStar Airplane ECN-2478 808. Montoya, L. C.; Economu, M. A.; and Cissell, R. E.: Use of a Pitot-Static Probe for Determining Wing Section Drag in Flight at Mach Numbers From 0.5 to Approximately 1.0. NASA TM X-56025, H-844, July 1974, 74N29370, #. The use of a pitot-static probe to determine wing section drag at speeds from Mach 0.5 to approximately 1.0 was evaluated in flight. The probe unit is described and operational problems are discussed. Typical wake profiles and wing section drag coefficients are presented. The data indicate that the pitot-static probe gave reliable results up to speeds of approximately 1.0. 809. Anon.: Advanced Control Technology and Its Potential for Future Transport Aircraft. NASA TM X-70240, Advanced Control Technology Symposium, Los Angeles, California, July 9–11, 1974, 74X10214, #. This document is a compilation of papers prepared for a symposium on advanced control technology sponsored by the National Aeronautics and Space Administration. This symposium focuses national attention on recent advances in control technology and the impact it should have on future transport aircraft. These technical papers present work performed by the Government and industry. The topics covered include recent flight-test results of advanced control 137 813. Layton, G. P.: NASA Flight Research Center Scale F-15 Remotely Piloted Research Vehicle Program. Advancements in Flight Test Engineering; Proceedings of the Fifth Annual Symposium, Anaheim, California, August 7–9, 1974, (see A74-43601 22-02), 74A43603. (See also 814.) The NASA Flight Research Center undertook a remotely piloted research vehicle (RPRV) program with a 3/8-scale model of an F-15 aircraft to determine the usefulness of the RPRV testing technique in high-risk flight testing such as spin testing. The results of the first flights of the program are presented. The program has shown that the RPRV technique, including the use of a digital control system, is a viable method for obtaining flight research data. Also presented are some negative aspects that have been learned about the RPRV technique in terms of model size, command frequency, and launch technique. 814. Layton, G. P.: NASA Flight Research Center Scale F-15 Remotely Piloted Research Vehicle Program. Soc. of Flight Test Engr. Advan. in Flight Test Eng., 1974, (see N75-10910 02-05), 75N10912. (See also 813.) A remotely piloted research vehicle (RPRV) program was conducted with a 3/8-scale model of an F-15 airplane to determine the usefulness of the RPRV testing technique in high risk flight testing such as spin testing. The results of the first flights of the program are presented. The program has shown that the RPRV technique, including the use of a digital control system, is a viable method for obtaining flight research data. Also presented are some negative aspects that have been learned about the RPRV technique in terms of model size, command frequency, and launch technique. 815. Iliff, K. W.: An Aircraft Application of System Identification in the Presence of State Noise. Presented at NATO Advantageous Study Institute, New Directions in Signal Processing in Communications and Control, Darlington, England, August 5–17, 1974, (AD-A001936 AFOSR-74-1756TR), 75N19234, #. (See also 872.) A maximum likelihood estimator for a linear system with state and observation noise is developed to determine unknown aircraft coefficients from flight data in the presence of turbulence (state noise). The formulation of the algorithm is presented briefly. The linear equations for an aircraft in atmospheric turbulence are defined. The effectiveness and accuracy of the technique are assessed by first applying it to simulated flight data, in which the true parameter values are known, then to actual flight data obtained in turbulence. A complete set of aircraft coefficients is obtained as well as an estimate of the turbulence time history. The validity of the estimated state noise and of the estimated coefficients is tested. The feasibility of using the algorithm for defining an adaptive control law to alleviate the effects of turbulence on the aircraft is discussed. 816. Berry, D. T.; and Gilyard, G. B.: Some Stability and Control Aspects of Airframe/Propulsion System Interactions on the YF-12 Airplane. Journal of Engineering for Industry, Transactions of the ASME, August 1974. Airframe/propulsion system interactions can strongly affect the stability and control of supersonic cruise aircraft. These interactions generate forces and moments similar in magnitude to those produced by the aerodynamic controls, and can cause significant changes in vehicle damping and static stability. This in turn can lead to large aircraft excursions or high pilot workload, or both. For optimum integration of an airframe and its jet propulsion system, these phenomena may have to be taken into account. 817. DeMarco, D. M.: A Dynamic Pressure Generator for Checking Complete Pressure Sensing Systems Installed on an Airplane. NASA TM X-56026, September 1974, 74N31923, #. A portable dynamic pressure generator, how it operates, and a test setup on an airplane are described. The generator is capable of providing a sinusoidal pressure having a peak-topeak amplitude of 3.5 N/sq cm (5 psi) at frequencies ranging from 100 hertz to 200 hertz. A typical power spectral density plot of data from actual dynamic pressure fluctuation tests within the air inlet of the YF-12 airplane is presented. 818. Barber, M. R.; and *Tymczyszyn, J. J.: Recent Wake Turbulence Flight Test Programs. Society of Experimental Test Pilots, Technical Review, Vol. 12, No. 2, 1974, pp. 52–68, 75A24805. In early flight tests the size and intensity of the wake vortexes generated by aircraft ranging in size from the Learjet to the C-5A and the B-747 were studied to determine the effects of aircraft configuration, weight, and speed. Early problems were related to vortex marking, the measurement of separation distance, and test techniques. Recent tests conducted with B-747 showed that vortexes were alleviated by reducing the deflection of the outboard flaps. It was found that a more rapid dissipation of the vortex system can be obtained through alterations in the span lift distribution. *FAA, Washington, D.C. 819. Smith, R. H.; and Burcham, F. W., Jr.: Instrumentation for In-Flight Determination of SteadyState and Dynamic Inlet Performance in Supersonic Aircraft. Instrumentation for Airbreathing Propulsion; Proceedings of the Symposium, Monterey, California, September 19–21, 1972. Cambridge, Massachusetts, MIT Press, 1974, pp. 41–58, 74A28286. (See also 726.) Advanced instrumentation and techniques for in-flight measurements of air inlet performance of the XB-70, 138 F-111A, and YF-12 supersonic airplanes were developed and evaluated in flight tests at the NASA Flight Research Center. A compressor face rake with in-flight zeroing capability was flown on the F-111A and found to give excellent steady state as well as high frequency response pressure data. The severe temperature environment of the YF-12 necessitated development of special high temperature transducers. Mounting these transducers to give the required 500-hertz frequency response required some special rake designs. Vibration test requirements necessitated some modifications to the rakes. The transducers and rakes were evaluated in flight tests and were found to function properly. Preliminary data have been obtained from the YF-12 propulsion program in flights that began in May 1972. One example shows the terminal shock wave effects on cowl surface pressures during bypass and spike motions. 820. Green, K. S.; and Putnam, T. W.: Measurements of Sonic Booms Generated by an Airplane Flying at Mach 3.5 and 4.8. NASA TM-X-3126, H-838, October 1974, 74N34486, #. Sonic booms generated by the X-15 airplane flying at Mach numbers of 3.5 and 4.8 were measured. The experimental results agreed within 12 percent with results obtained from theoretical methods. No unusual phenomena related to overpressure were encountered. Scaled data from the X-15 airplane for Mach 4.8 agreed with data for an SR-71 airplane operating at lower Mach numbers and similar altitudes. The simple technique used to scale the data on the basis of airplane lift was satisfactory for comparing X-15 and SR-71 sonic boom signatures. 821. Gilyard, G. B.; and Belte, D.: Flight-Determined Lag of Angle-of-Attack and Angle-of-Sideslip Sensors in the YF-12A Airplane From Analysis of Dynamic Maneuvers. NASA TN D-7819, H-767, October 1974, 74N34460, #. Magnitudes of lags in the pneumatic angle-of-attack and angle-of-sideslip sensor systems of the YF-12A airplane were determined for a variety of flight conditions by analyzing stability and control data. The three analysis techniques used are described. An apparent trend with Mach number for measurements from both of the differentialpressure sensors showed that the lag ranged from approximately 0.15 second at subsonic speed to 0.4 second at Mach 3. Because Mach number was closely related to altitude for the available flight data, the individual effects of Mach number and altitude on the lag could not be separated clearly. However, the results indicated the influence of factors other than simple pneumatic lag. 822. Saltzman, E. J.; Meyer, R. R., Jr.; and Lux, D. P.: Drag Reductions Obtained by Modifying a Box-Shaped Ground Vehicle. NASA TM X-56027, October 1974, 74N34449, #. 139 A box-shaped ground vehicle was used to simulate the aerodynamic drag of high volume transports, that is, delivery vans, trucks, or motor homes. The coast-down technique was used to define the drag of the original vehicle, having all square corners, and several modifications of the vehicle. Test velocities ranged up to 65 miles per hour, which provided maximum Reynolds numbers of 1 times 10 to the 7th power based on vehicle length. One combination of modifications produced a reduction in aerodynamic drag of 61 percent as compared with the original square-cornered vehicle. 823. Wolowicz, C. H.; and Yancey, R. B.: Experimental Determination of Airplane Mass and Inertial Characteristics. NASA TR R-433, H-814, October 1974, 75N10062, #. Current practices are evaluated for experimentally determining airplane center of gravity, moments of inertia, and products of inertia. The techniques discussed are applicable to bodies other than airplanes. In pitching- and rolling-moment-of-inertia investigations with the airplane mounted on and pivoted about knife edges, the nonlinear spring moments that occur at large amplitudes of oscillation can be eliminated by using the proper spring configuration. The single-point suspension double-pendulum technique for obtaining yawing moments of inertia, products of inertia, and the inclination of the principal axis provides accurate results from yaw-mode oscillation data, provided that the swaymode effects are minimized by proper suspension rig design. Rocking-mode effects in the data can be isolated. 824. Schweikhard, W. G.: Test Techniques, Instrumentation, and Data Processing. AGARD Lecture Series No. 72, Distortion Induced Eng. Instability, November 7–15, 1974, (see N75-12954 04-07), 75N12960. Procedures for determining the effects of dynamic distortion on engine stability are analyzed. The test techniques, methods and types of instrumentation, and data processing functions are described. The advantages and limitations of various methods are reported. It is emphasized that ground facility tests are only a simulation of the flight environment, that instrumentation provides only a partial representation of the physical phenomena, and that poorly organized data processing procedures can impede and even distort the final result. 825. Bellman, D. R.; and Kier, D. A.:HiMAT—A New Approach to the Design of Highly Maneuverable Aircraft. SAE Paper 740859, Society of Automotive Engineers, National Aerospace Engineering and Manufacturing Meeting, San Diego, California, October 1–3, 1974, 75A16921. Needed improvements in the maneuvering performance of combat aircraft appear to be possible through the simultaneous application of advances in various disciplines in such a way that they complement one another and magnify the benefits derived. The highly maneuverable aircraft technology (HiMAT) program is being conducted to investigate such multidisciplinary concepts. The program has three phases: preliminary studies, conceptual design studies, and the final design and construction of a test airplane. Work is now in the second phase. The test airplane will be a scaled model flown by a remotely piloted research vehicle technique. This paper outlines the HiMAT program and indicates the types of concepts being considered. 826. Reukauf, P. J.; Schweikhard, W. G.; and *Arnaiz, H.: Flight-Test Techniques for Obtaining Valid Comparisons of Wind-Tunnel and Flight Results From Tests on a YF-12 Mixed-Compression Inlet. AIAA Paper 74-1195, Tenth AIAA Society of Automotive Engineers, Propulsion Conference, San Diego, California, October 21–23, 1974, 75A11301, #. The ability to predict the inlet characteristics of high-speed propulsion systems from wind-tunnel test results is being studied by the NASA Flight Research Center in a flight program on the YF-12 aircraft. The obvious requirement for matching geometry, instrumentation, and test conditions has led to the development of special flight test techniques, hardware, and systems. This paper describes this development, the technical and operational problems encountered and their solutions and the compromises that were found to be necessary. 827. Love, J. E.: Flight Test Results of an Automatic Support System on Board a YF-12A Airplane. Automatic Support Systems for Advanced Maintainability Symposium, San Diego, California, October 30–November 1, 1974, Institute of Electrical and Electronics Engineers, Inc., 1974, pp. 211–220, 75A35272. An automatic support system concept that isolated faults in an existing nonavionics subsystem was flight tested up to a Mach number of 3. The adaptation of the automated support concept to an existing system (the jet engine automatic inlet control system) caused most of the problems one would expect to encounter in other applications. These problems and their solutions are discussed. Criteria for integrating automatic support into the initial design of new subsystems are included in the paper. Cost effectiveness resulted from both the low maintenance of the automated system and the man-hour saving resulting from the real time diagnosis of the monitored subsystem. 828. Larson, T. J.: Compensated and Uncompensated Nose Boom Static Pressures Measured From Two Air Data Systems on a Supersonic Airplane. NASA TM X-3132, H-835, November 1974, 79N77423. Two static-pressure-measuring air data systems that were used on the YF-12 airplane for supersonic flight testing were compared. One system consisted of a nose boom pitot-static 140 probe with two sets of static-pressure orifices designed for static-pressure error compensation, two air data computers, and a photopanel for recording. The other system consisted of an identical nose boom probe and a third set of static-pressure orifices not designed for error compensation, pressure transducers for direct pressure measurements instead of air data computers, and 5 pulse code modulation system for recording. The comparisons showed that the uncompensated static-pressure orifices provided more accurate air data measurements than either set of compensated static-pressure sources. Whereas the uncompensated static-pressure source was relatively insensitive to angle of attack, the compensated sources were characterized by a position error at supersonic speeds that increased with angle of attack and Mach number. Pitot-static measurements acquired by using air data computers that incorporate cams for static error compensation provide reference data that are less accurate than similar measurements made by pressure transducers. 829. Deets, D. A.; and Edwards, J. W.: A Remotely Augmented Vehicle Approach to Flight Testing RPV Control Systems. NASA TM X-56029, H-870, November 1974, 75N10936, #. A remotely augmented vehicle concept for flight testing advanced control systems was developed as an outgrowth of a remotely piloted research vehicle (RPV) program in which control laws are implemented through telemetry uplink and computer which provides the control law computations. Some advantages of this approach are that the cost of one control system facility is spread over a number of RPV programs, and control laws can be changed quickly as required, without changing the flight hardware. The remotely augmented vehicle concept is described, and flight test results from a subscale F-15 program are discussed. Suggestions of how the concept could lead to more effective testing of RPV control system concepts, and how it is applicable to a military RPV reconnaissance mission are given. 830. Ehernberger, L. J.: High Altitude Turbulence Encountered by the Supersonic YF-12A Airplane. Sixth Annual Conference on Aerospace and Aeronautical Meteorology, El Paso, Texas, November 12–15, 1974, pp. 305–312, (see A75-35351 16-47), 75A35409, #. The present work describes the turbulence experienced by the YF-12A airplane on the basis of airplane acceleration data obtained at altitudes above 12.2 km. Data presented include the subjective intensities reported by the air crew, the portion of flight distance in turbulence, the variation of turbulence with season, and the thickness and length of turbulence patches as determined along the flight path. Compared with that experienced by subsonic jets below 12.2 km, turbulence above 12.2 km was mild, but the crew was more sensitive to gust accelerations during supersonic flight at altitudes above 12.2 km than during subsonic flight at lower altitudes. About 6–8% of the distance traveled was in turbulence between 12.2 and 16.8 km, as compared to less than 1% above 18.3 km. High-altitude turbulence increased by a factor of three from summer to winter. Turbulence patches were 0.4 km thick and 10 km long on the average. 831. Montoya, L. C.; and Steers, L. L.: Aerodynamic Drag Reduction Tests on a Full-Scale Tractor-Trailer Combination With Several Add-On Devices. NASA TM X-56028, Reduction of Aerodynamic Drag of Trucks Conference, Pasadena, California, October 10–11, 1974, December 1974, 75N12900, #. (See also 832.) Aerodynamic drag tests were performed on a conventional cab-over-engine tractor with a 45-foot trailer and five commercially available or potentially available add-on devices using the coast-down method. The tests ranged in velocity from approximately 30 miles per hour to 65 miles per hour and included some flow visualization. A smooth, level runway at Edwards Air Force Base was used for the tests, and deceleration measurements were taken with both accelerometers and stopwatches. An evaluation of the drag reduction results obtained with each of the five add-on devices is presented. 832. Montoya, L. C.; and Steers, L. L.: Aerodynamic Drag Reduction Tests on a Full-Scale Tractor-Trailer Combination With Several Add-On Devices. Presented at the Reduction of Aerodynamic Drag of Trucks Conference, Pasadena, California, October 10–11, 1974. Proceedings of the Conference Workshop, (available from the National Science Foundation, RANN Document Center, Washington, D.C.) December 1974, pp. 63–88. (See also 831.) 833. Schweikhard, W. G.; and Montoya, E. J.: Research Instrumentation Requirements for Flight/Wind-Tunnel Tests of the YF-12 Propulsion System and Related Flight Experience. Instrumentation for Airbreathing Propulsion; Proceedings of the Symposium, Monterey, California, September 19–21, 1972, MIT Press, Cambridge Massachusetts, 1974, pp. 19–39, (see A74-28283 12-14), 74A28285. Description of the requirements for a comprehensive flight and wind-tunnel propulsion research program to examine the predictability of inlet performance, evaluate the effects of high-frequency pressure phenomena on inlets, and investigate improved control concepts in order to cope with airframe interactions. This program is unique in that it requires precise similarity of the geometry of the flight vehicle and tunnel modes; the test conditions, including local flow at the inlet; and instrumentation. Although few windtunnel instrumentation problems exist, many problems emerge during flight tests because of the thermal environment. Mach 3 flight temperatures create unique problems with transducers, connectors, and wires. All must be capable of withstanding continuous 1000 F temperatures, as well as the mechanical stresses imposed by vibration and thermal cycling. 1975 Technical Publications 834. Barber, M. R.; Kurkowski, R. L.; Garodz, L. J.; Robinson, G. H.; Smith, H. J.; Jacobsen, R. A.; Stinnett, G. W., Jr.; McMurtry, T. C.; Tymczyszyn, J. J.; and Devereaux, R. L.: Flight Test Investigation of the Vortex Wake Characteristics Behind a Boeing 727 During TwoSegment and Normal ILS Approaches (A Joint NASA/FAA Report). NASA TM X-62398, FAA-NA-151, January 1975, 75N17340, #. Flight tests were performed to evaluate the vortex wake characteristics of a Boeing 727 aircraft during conventional and two-segment instrument landing approaches. Smoke generators were used for vortex marking. The vortex was intentionally intercepted by a Lear Jet and a Piper Comanche aircraft. The vortex location during landing approach was measured using a system of phototheodolites. The tests showed that at a given separation distance there are no readily apparent differences in the upsets resulting from deliberate vortex encounters during the two types of approaches. The effect of the aircraft configuration on the extent and severity of the vortices is discussed. 835. Ehernberger, L. J.; and Love, B. J.: High Altitude Gust Acceleration Environment as Experienced by a Supersonic Airplane. NASA TN D-7868, H-836, January 1975, 75N13791, #. High altitude turbulence experienced at supersonic speeds is described in terms of gust accelerations measured on the YF-12A airplane. The data were obtained during 90 flights at altitudes above 12.2 kilometers (40,000 feet). Subjective turbulence intensity ratings were obtained from air crew members. The air crew often rated given gust accelerations as being more intense during high altitude supersonic flight than during low altitude subsonic flight. The portion of flight distance in turbulence ranged from 6 percent to 8 percent at altitudes between 12.2 kilometers and 16.8 kilometers (40,000 feet and 55,000 feet) to less than 1 percent at altitudes above 18.3 kilometers (60,000 feet). The amount of turbulence varied with season, increasing by a factor of 3 or more from summer to winter. Given values of gust acceleration were less frequent, on the basis of distance traveled, for supersonic flight of the YF-12A airplane at altitudes above 12.2 kilometers (40,000 feet) than for subsonic flight of a jet passenger airplane at altitudes below 12.2 kilometers (40,000 feet). The median thickness of high altitude turbulence patches was less than 400 meters (1300 feet); the median length was less than 16 kilometers (10 miles). The distribution of the patch dimensions tended to be log normal. 836. Rediess, H. A.; and Szalai, K. J.: Status and Trends in Active Control Technology. NASA SP-372, 141 January 1975, 75N29015. pp. 273–322, (see N75-29001 20-01), The emergence of highly reliable fly-by-wire flight control systems makes it possible to consider a strong reliance on automatic control systems in the design optimization of future aircraft. This design philosophy has been referred to as the control configured vehicle approach or the application of active control technology. Several studies and flight tests sponsored by the Air Force and NASA have demonstrated the potential benefits of control configured vehicles and active control technology. The present status and trends of active control technology are reviewed and the impact it will have on aircraft designs, design techniques, and the designer is predicted. 837. Smith, J. W.; and Berry, D. T.: Analysis of Longitudinal Pilot-Induced Oscillation Tendencies of YF-12 Aircraft. NASA TN D-7900, H-805, February 1975, 75N16560, #. Aircraft flight and ground tests and simulator studies were conducted to explore pilot-induced oscillation tendencies. Linear and nonlinear calculations of the integrated flight control system’s characteristics were made to analyze and predict the system’s performance and stability. The investigations showed that the small-amplitude PIO tendency was caused by the interaction of the pilot with a combination of the aircraft’s short-period poles and the structural first bending mode zeros. It was found that the large-amplitude PIOs were triggered by abrupt corrective control actions by the pilot, which caused the stability augmentation system servo to position and rate limit. The saturation in turn caused additional phase lag, further increasing the tendency of the overall system to sustain a PIO. 838. Anon.: Description and Flight Test Results of the NASA F-8 Digital Fly-By-Wire Control System. NASA TN D-7843, H-853. Presented at the NASA Symposium on Advanced Control Technol., Los Angeles, California, July 9–11, 1974, February 1975, 75N18245, #. A NASA program to develop digital fly-by-wire (DFBW) technology for aircraft applications is discussed. Phase I of the program demonstrated the feasibility of using a digital fly-by-wire system for aircraft control through developing and flight testing a single channel system, which used Apollo hardware, in an F-8C airplane. The objective of Phase II of the program is to establish a technology base for designing practical DFBW systems. It will involve developing and flight testing a triplex digital fly-by-wire system using stateof-the-art airborne computers, system hardware, software, and redundancy concepts. The papers included in this report describe the Phase I system and its development and present results from the flight program. Man-rated flight software and the effects of lightning on digital flight control systems are also discussed. 142 839. Jarvis, C. R.: An Overview of NASA’s Digital FlyBy-Wire Technology Development Program. NASA TN D-7843, H-853, Description and Flight Test Results of the NASA F-8 Digital Fly-by-Wire Control System, February 1975, pp. 1–12, (see N75-18245 10-08), 75N18246. (See also 898.) The feasibility of using digital fly-by-wire systems to control aircraft was demonstrated by developing and flight testing a single channel system, which used Apollo hardware, in an F-8C test airplane. This is the first airplane to fly with a digital fly-by-wire system as its primary means of control and with no mechanical reversion capability. The development and flight test of a triplex digital fly-by-wire system, which will serve as an experimental prototype for future operational digital fly-by-wire systems, is underway. 840. Deets, D. A.: Design and Development Experience With a Digital Fly-By-Wire Control System in an F-8C Airplane. NASA TN D-7843, H-853. Description and Flight Test Results of the NASA F-8 Digital Fly-by-Wire Control System, February 1975, pp. 13–40, (see N75-18245 10-08), 75N18247. (See also 899.) To assess the feasibility of a digital fly-by-wire system, the mechanical flight control system of an F-8C airplane was replaced with a digital system and an analog backup system. The Apollo computer was used as the heart of the primary system. This paper discusses the experience gained during the design and development of the system and relates it to active control systems that are anticipated for future civil transport applications. 841. Lock, W. P.; Petersen, W. R.; and *Whitman, G. B.: Mechanization of and Experience With a Triplex Fly-ByWire Backup Control System. NASA TN D-7843, H-853. Description and Flight Test Results of the NASA F-8 Digital Fly-by-Wire Control System, February 1975, pp. 41–72, (see N75-18245 10-08), 75N18248. (See also 900.) A redundant three-axis analog control system was designed and developed to back up a digital fly-by-wire control system for an F-8C airplane. Forty-two flights, involving 58 hours of flight time, were flown by six pilots. The mechanization and operational experience with the backup control system, the problems involved in synchronizing it with the primary system, and the reliability of the system are discussed. The backup control system was dissimilar to the primary system, and it provided satisfactory handling through the flight envelope evaluated. Limited flight tests of a variety of control tasks showed that control was also satisfactory when the backup control system was controlled by a minimumdisplacement (force) side stick. The operational reliability of the F-8 digital fly-by-wire control system was satisfactory, with no unintentional downmodes to the backup control system in flight. The ground and flight reliability of the system’s components is discussed. *Sperry Flight Systems Div., Phoenix, Arizona. 842. *Plumer, J. A.; **Malloy, W. A.; and Craft, J. B.: The Effects of Lightning on Digital Flight Control Systems. NASA TN D-7843, H-853. Description and Flight Test Results of the NASA F-8 Digital Fly-by-Wire Control System, February 1975, pp. 73–92, (see N75-18245 10-08), 75N18249. (See also 905.) Present practices in lightning protection of aircraft deal primarily with the direct effects of lightning, such as structural damage and ignition of fuel vapors. There is increasing evidence of troublesome electromagnetic effects, however, in aircraft employing solid-state microelectronics in critical navigation, instrumentation and control functions. The potential impact of these indirect effects on critical systems such as digital fly-by-wire (DFBW) flight controls has been studied by several recent research programs, including an experimental study of lightning-induced voltages in the NASA F8 DFBW airplane. The results indicate a need for positive steps to be taken during the design of future fly-by-wire systems to minimize the possibility of hazardous effects from lightning. *General Electric Research and Development, Schenectady, New York. **General Motors Corp., Detroit, Michigan. 843. Szalai, K. J.: Flight Test Experience With the F-8 Digital Fly-By-Wire System. NASA TN D-7843, H-853. Description and Flight Test Results of the NASA F-8 Digital Fly-by-Wire Control System, February 1975, pp. 127–180, (see N75-18245 10-08),75N18251. (See also 901.) Flight test results of the F-8 digital fly-by-wire (DFBW) control system are presented and the implications for application to active control technology (ACT) are discussed. The F-8 DFBW system has several of the attributes of proposed ACT systems, so the flight test experience is helpful in assessing the capabilities of those systems. Topics of discussion include the predicted and actual flight performance of the control system, assessments of aircraft flying qualities and other piloting factors, software management and control, and operational experience. 844. Krier, G. E.: A Pilot’s Opinion of the F-8 Digital Fly-By-Wire Airplane. NASA TN D-7843, H-853. Description and Flight Test Results of the NASA F-8 Digital Fly-by-Wire Control System, February 1975, pp. 181–195, (see N75-18245 10-08), 75N18252. (See also 902.) The handling qualities of the F-8 digital fly-by-wire airplane are evaluated by using the Cooper-Harper rating scale. The reasons for the ratings are given, as well as a short description of the flying tasks. It was concluded that the handling qualities of the airplane were good in most situations, although occasional ratings of unsatisfactory were given. 845. Nugent, J.; Couch, L. M.; and Webb, L. D.: Exploratory Wind Tunnel Tests of a Shock-Swallowing 143 Air Data Sensor at a Mach Number of Approximately 1.83. NASA TM X-56030, March 1975, 75N20329, #. The test probe was designed to measure free-stream Mach number and could be incorporated into a conventional airspeed nose boom installation. Tests were conducted in the Langley 4-by 4-foot supersonic pressure tunnel with an approximate angle of attack test range of –5 deg to 15 deg and an approximate angle of sideslip test range of + or –4 deg. The probe incorporated a variable exit area which permitted internal flow. The internal flow caused the bow shock to be swallowed. Mach number was determined with a small axially movable internal total pressure tube and a series of fixed internal static pressure orifices. Mach number error was at a minimum when the total pressure tube was close to the probe tip. For four of the five tips tested, the Mach number error derived by averaging two static pressures measured at horizontally opposed positions near the probe entrance were least sensitive to angle of attack changes. The same orifices were also used to derive parameters that gave indications of flow direction. 846. Steers, S. T.; and Iliff, K. W.: Effects of TimeShifted Data on Flight Determined Stability and Control Derivatives. NASA TN D-7830, H-849, March 1975, 75N18244, #. Flight data were shifted in time by various increments to assess the effects of time shifts on estimates of stability and control derivatives produced by a maximum likelihood estimation method. Derivatives could be extracted from flight data with the maximum likelihood estimation method even if there was a considerable time shift in the data. Time shifts degraded the estimates of the derivatives, but the degradation was in a consistent rather than a random pattern. Time shifts in the control variables caused the most degradation, and the lateral-directional rotary derivatives were affected the most by time shifts in any variable. 847. Putnam, T. W.; Lasagna, P. L.; and *White, K. C.: Measurements and Analysis of Aircraft Airframe Noise. AIAA Paper 75-510, presented at the American Institute of Aeronautics and Astronautics, Second Annual AeroAcoustics Conference, Hampton, Virginia, March 1975, 75A25776, #. (Also Aeroacoustics: STOL Noise; Airframe and Airfoil Noise, Vol. 45, Progress in Astronautics and Aeronautics, March 1975.) Flyover measurements of the airframe noise of AeroCommander, JetStar, CV-990, and B-747 aircraft are presented. Data are shown for both cruise and landing configurations. Correlations between airframe noise and aircraft parameters are developed and presented. The landing approach airframe noise for the test aircraft was approximately 10 EPNdB below present FAA certification requirements. *NASA Ames Research Center, Moffett Field, California. 848. Sanderson, K. C.: A New Flight Test Data System for NASA Aeronautical Flight Research. Proceedings, London Royal Aeronautical Society, Eighth International Aerospace Instrumentation Symposium, Cranfield, Beds., England, March 24–27, 1975, (see A75-28765 12-06), 75A28774, #. The airborne integrated flight test data system (AIFTDS) is described. This system integrates an airborne digital computer with a high-bit-rate pulse code modulation system. Its design was influenced by in-house technical experience with similar modules and by the multiproject environment in which it was expected to operate. The present work describes events leading to the development of the system, reviews factors that influenced the objectives for the system and the resulting design, and describes the elements themselves. Block diagrams supplement the text. 849. Edwards, J. W.; and Deets, D. A.: Development of a Remote Digital Augmentation System and Application to a Remotely Piloted Research Vehicle. NASA TN D-7941, H-854, April 1975, 75N20293, #. A cost-effective approach to flight testing advanced control concepts with remotely piloted vehicles is described. The approach utilizes a ground based digital computer coupled to the remotely piloted vehicle’s motion sensors and control surface actuators through telemetry links to provide high bandwidth feedback control. The system was applied to the control of an unmanned 3/8-scale model of the F-15 airplane. The model was remotely augmented; that is, the F-15 mechanical and control augmentation flight control systems were simulated by the ground-based computer, rather than being in the vehicle itself. The results of flight tests of the model at high angles of attack are discussed. 850. Smith, H. J.: A Flight Test Investigation of the Rolling Moments Induced on a T-37B Airplane in the Wake of a B-747 Airplane. NASA TM X-56031, April 1975, 75N20221, #. A flight test investigation of the B-747 vortex wake characteristics was conducted using a T-37B as a probe aircraft. The primary purpose of the program was the validation of the results of B-747 model tests which predicted significant alleviation of the vortex strength when only the inboard flaps were deflected. Measurements of the vortexinduced rolling moments of the probe aircraft showed that the predicted alleviation did occur. The effects of landing gear extension, increased lift coefficient, idle thrust, and sideslip were investigated, and all had an adverse effect on the alleviated condition as evidenced by increased induced rolling moments of the T-37B probe aircraft. Idle thrust also increased the strength of the B-747 wake vortexes with both inboard and outboard flaps extended. 851. Maine, R. E.; and Iliff, K. W.: A FORTRAN Program for Determining Aircraft Stability and Control 144 Derivatives From Flight Data. NASA TN D-7831, H-856, April 1975, 75N25621, #. A digital computer program written in FORTRAN IV for the estimation of aircraft stability and control derivatives is presented. The program uses a maximum likelihood estimation method, and two associated programs for routine, related data handling are also included. The three programs form a package that can be used by relatively inexperienced personnel to process large amounts of data with a minimum of manpower. This package was used to successfully analyze 1500 maneuvers on 20 aircraft, and is designed to be used without modification on as many types of computers as feasible. Program listings and sample check cases are included. 852. Montoya, L. C.; and Lux, D. P.: Comparisons of Wing Pressure Distribution From Flight Tests of Flush and External Orifices for Mach Numbers From 0.50 to 0.97. NASA TM X-56032, April 1975, 75N22275, #. Wing pressure distributions obtained in flight with flush orifice and external tubing orifice installations for Mach numbers from 0.50 to 0.97 are compared. The procedure used to install the external tubing orifice is discussed. The results indicate that external tubing orifice installations can give useful results. 853. Bennett, D. L.: Evaluation of a Hemispherical Head Flow Direction Sensor for Inlet Duct Measurements. NASA TM X-3232, H-862, May 1975, 75N22277, #. A hemispherical head flow direction sensor was tested in a wind tunnel to evaluate its effectiveness for measuring dynamic duct flow direction angles of plus and minus 27 degrees. The tests were conducted at Reynolds numbers of 3.8 million per meter (1.0 million per foot) and 4.92 million per meter (1.5 million per foot) and at Mach numbers from 0.30 to 0.70. The design criteria for the probe are discussed and the wind tunnel results are presented. Three techniques for deriving the flow angles are described. 854. Iliff, K. W.; and Maine, R. E.: Practical Aspects of Using a Maximum Likelihood Estimator. AGARD CP-172. Methods for Aircraft State and Parameter Identification, May 1975, (see N75-29997 21-01), 75N30013. The application of a maximum likelihood estimator to flight data is discussed and procedures to facilitate routine analysis of a large amount of flight data are proposed. Flight data were used to demonstrate the proposed procedures. Modeling considerations are discussed for the system to be identified, including linear aerodynamics, instrumentation, and data time shifts, and aerodynamic biases for the specific types of maneuvers to be analyzed. Data editing to eliminate common data acquisition problems, and a method of identifying other problems are considered. The need for careful selection of the maneuver or portions of the maneuver to be analyzed is pointed out. Uncertainty levels (analogous to Cramer-Rao bounds) are discussed as a way of recognizing significant new information. 855. Pyle, J. S.; and Steers, L. L.: Flight Determined Lift and Drag Characteristics of an F-8 Airplane Modified With a Supercritical Wing With Comparison to Wind-Tunnel Results. NASA TM X-3250, H-843, June 1975, 79N33159, #. Flight measurements obtained with a TF-8A airplane modified with a supercritical wing are presented for altitudes from 7.6 kilometers (25,000 feet) to 13.7 kilometers (45,000 feet), Mach numbers from 0.6 to 1.2, and Reynolds numbers from 0.8 × 10 to the 7th power to 2.3 × 10 to the 7th power. Flight results for the airplane with and without area-rule fuselage fairings are compared. The techniques used to determine the lift and drag characteristics of the airplane are discussed. Flight data are compared with windtunnel model results, where applicable. 856. Hedgley, D. R., Jr.: An Algorithm and Computer Program to Locate Real Zeros of Real Polynomials. NASA TN D-8009, H-855, June 1975, 75N25651, #. A method for reliably extracting real zeros of real polynomials using an expanded two-point secant and bisection method is formed into an algorithm for a digital computer, and a computer program based on this algorithm is presented. The results obtained with the program show that the proposed method compares favorably with the Laguerre, Newton-Raphson, and Jenkins-Traub methods when the polynomial has all real zeros, and is more efficient when the polynomial has complex zeros. 857. Szalai, K. J.; and Deets, D. A.: F-8 Digital Fly-ByWire Flight Test Results Viewed From an Active Controls Perspective. Impact of Active Control Technologies on Airplane Design, June 1975, (see N75-30027 21-01), 75N30049. The results of the NASA F-8 digital fly-by-wire flight test program are presented, along with the implications for active controls applications. The closed loop performance of the digital control system agreed well with the sampled-data system design predictions. The digital fly-by-wire mechanization also met pilot flying qualities requirements. The advantages of mechanizing the control laws in software became apparent during the flight program and were realized without sacrificing overall system reliability. This required strict software management. The F-8 flight test results are shown to be encouraging in light of the requirements that must be met by control systems for flight-critical active controls applications. 858. Kempel, R. W.; Dana, W. H.; and Sim, A. G.: Flight Evaluation of the M2-F3 Lifting Body Handling Qualities 145 at Mach Numbers From 0.30 to 1.61. NASA TN D-8027, H-852, July 1975, 75N27015, #. Percentage distributions of 423 pilot ratings obtained from 27 flights are used to indicate the general level of handling qualities of the M2-F3 lifting body. Percentage distributions are compared on the basis of longitudinal and lateraldirectional handling qualities, control system, control system status, and piloting task. Ratings of longitudinal handling qualities at low speed were slightly better than those for transonic and supersonic speed. The ratings of lateraldirectional handling qualities were unaffected by speed and configuration. Specific handling qualities problems are discussed in detail, and comparisons are made with pertinent handling qualities criteria. 859. Smith, H. J.: Flight-Determined Stability and Control Derivatives for an Executive Jet Transport. NASA TM X-56034, H-901, July 1975, 76N11105, #. A modified maximum likelihood estimation (MMLE) technique which included a provision for including a priori information about unknown parameters was used to determine the aerodynamic derivatives of the Lockheed JetStar airplane. Two hundred sixty-five maneuvers were performed with the JetStar airplane, which was modified to include direct lift controls, to obtain lateral-directional and longitudinal derivatives. Data were obtained at altitudes of 3048 meters, 6096 meters, and 9144 meters (10,000 feet, 20,000 feet, and 30,000 feet) and over an angle of attack range from approximately 3 deg to 13 deg and a Mach number range from 0.25 to 0.75. Side force generators were installed and tested in 87 maneuvers to determine their effectiveness and their effect on the other derivatives. Lateral-directional data for four flight conditions were analyzed without using a priori information to assess the effect of this feature on the results. The MMLE method generally gave consistent (repeatable) estimates of the derivatives, with the exception of the rolling moment due to yaw rate, which showed large variances. 860. Steers, L. L.; Montoya, L. C.; and Saltzman, E. J.: Aerodynamic Drag Reduction Tests on a Full-Scale Tractor-Trailer Combination and a Representative BoxShaped Ground Vehicle. SAE Paper 750703, presented at the Society of Automotive Engineers, West Coast Meeting, Seattle, Washington, August 11–14, 1975, 76A14479. Aerodynamic drag tests were performed on a tractor-trailer combination and a box-shaped ground vehicle using the coast-down method on a smooth, nearly level runway. The tractor-trailer tests included an investigation of drag reduction add-on devices that are commercially available or under development. The box-shaped vehicle was modified by rounding the corners and sealing the undercarriage. The tests ranged in velocity from approximately 35 miles per hour to 65 miles per hour for the tractor-trailer combination and included fuel consumption measurements and one set of measurements of drive shaft torque. This paper presents the results for both vehicles, showing the effects of the various modifications on the aerodynamic drag. The effects of variation in the aerodynamic drag of the tractor-trailer combination on fuel consumption are also presented. 861. Arnaiz, H. H.: Techniques for Determining Propulsion System Forces for Accurate High Speed Vehicle Drag Measurements in Flight. AIAA Paper 75-964, presented at the AIAA Aircraft Systems and Technology Meeting, Los Angeles, California, August 4–7, 1975, 75A41689, #. As part of a NASA program to evaluate current methods of predicting the performance of large, supersonic airplanes, the drag of the XB-70 airplane was measured accurately in flight at Mach numbers from 0.75 to 2.5. This paper describes the techniques used to determine engine net thrust and the drag forces charged to the propulsion system that were required for the in-flight drag measurements. The accuracy of the measurements and the application of the measurement techniques to aircraft with different propulsion systems are discussed. Examples of results obtained for the XB-70 airplane are presented. 862. Putnam, T. W.: Review of Aircraft Noise Propagation. NASA TM X-56033, September 1975, 75N32119, #. The current state of knowledge about the propagation of aircraft noise was reviewed. The literature on the subject is surveyed and methods for predicting the most important and best understood propagation effects are presented. Available empirical data are examined and the data’s general validity is assessed. The methods used to determine the loss of acoustic energy due to uniform spherical spreading, absorption in a homogeneous atmosphere, and absorption due to ground cover are presented. A procedure for determining ground induced absorption as a function of elevation angle between source and receiver is recommended. Other factors that affect propagation, such as refraction and scattering due to turbulence, which were found to be less important for predicting the propagation of aircraft noise, are also evaluated. 863. Shafer, M. F.: Stability and Control Derivatives of the T-37B Airplane. NASA TM X-56036, September 1975, 76N14137, #. Subsonic stability and control derivatives were determined by a modified maximum likelihood estimator from flight data for the longitudinal and lateral-directional modes of the T-37B airplane. Data from two flights, in which 166 stability and control maneuvers were performed, were used in the determination. The configurations investigated were: zero flaps, gear up; half flaps, gear up; full flaps, gear up; and zero flaps, gear down. 864. Reukauf, P. J.; Burcham, F. W., Jr.; and Holzman, J. K.: Status of a Digital Integrated Propulsion/Flight Control System for the YF-12 Airplane. AIAA Paper 75-1180, presented at the AIAA Institute of Aeronautics and Astronautics and Society of Automotive Engineers, Eleventh Annual Propulsion Conference, Anaheim, California, September 29–October 1, 1975, 76A10252, #. The NASA Flight Research Center is engaged in a program with the YF-12 airplane to study the control of interactions between the airplane and the propulsion system. The existing analog air data computer, autothrottle, autopilot, and inlet control system are to be converted to digital systems by using a general purpose airborne computer and interface unit. First, the existing control laws will be programmed in the digital computer and flight tested. Then new control laws are to be derived from a dynamic propulsion model and a total force and moment aerodynamic model to integrate the systems. These control laws are to be verified in a real time simulation and flight tested. 865. Garodz, L. J.: Flight Test Investigation of the Vortex Wake Characteristics Behind a Boeing 727 During Two-Segment and Normal ILS Approaches. NASA TM X-72908, FAA-NA-75-151, AD-A018366. Joint study with National Aviation Facilities Experimental Center, Atlantic City, New Jersey, October 1975, 76N14046, #. A series of flight tests were performed to evaluate the vortex wake characteristics of a Boeing 727 (B727-200) aircraft during conventional and two-segment ILS approaches. Flights of the B727, equipped with smoke generators for vortex marking, were flown wherein its vortex wake was intentionally encountered by a Lear Jet model 23 (LR-23) or a Piper Twin Comanche (PA-30); and its vortex location during landing approach was measured using a system of ECN-3831 B-727 Airplane With Wingtip Smoke Generators 146 photo-theodolites. The tests showed that at a given separation distance there were no differences in the upsets resulting from deliberate vortex encounters during the two types of approaches. Timed mappings of the position of the landing configuration vortices showed that they tended to descend approximately 91 meters (300 feet) below the flight path of the B727. The flaps of the B727 have a dominant effect on the character of the trailed wake vortex. The clean wing produces a strong, concentrated vortex. As the flaps are lowered, the vortex system becomes more diffuse. Pilot opinion and roll acceleration data indicate that 4.5 nautical miles would be a minimum separation distance at which roll control could be maintained during parallel encounters of the B727’s landing configuration wake by small aircraft. 866. Burcham, F. W., Jr.; Putnam, T. W.; Lasagna, P. L.; and Parish, O. O.: Measured Noise Reductions Resulting From Modified Approach Procedures for Business Jet Aircraft. NASA TM X-56037, November 1975, 76N32973, #. Five business jet airplanes were flown to determine the noise reductions that result from the use of modified approach procedures. The airplanes tested were a Gulfstream 2, JetStar, Hawker Siddeley 125-400, Sabreliner-60 and LearJet-24. Noise measurements were made 3, 5, and 7 nautical miles from the touchdown point. In addition to a standard 3 deg glide slope approach, a 4 deg glide slope approach, a 3 deg glide slope approach in a low-drag configuration, and a twosegment approach were flown. It was found that the 4 deg approach was about 4 EPNdB quieter than the standard 3 deg approach. Noise reductions for the low-drag 3 deg approach varied widely among the airplanes tested, with an average of 8.5 EPNdB on a fleet-weighted basis. The two-segment approach resulted in noise reductions of 7 to 8 EPNdB at 3 and 5 nautical miles from touchdown, but only 3 EPNdB at 7 nautical miles from touchdown when the airplanes were still in level flight prior to glide slope intercept. Pilot ratings showed progressively increasing workload for the 4 deg, lowdrag 3 deg, and two-segment approaches. 867. Gee, S. W.; Wolf, T. D.; and Rezek, T. W.: Passenger Ride Quality Response to an Airborne Simulator Environment. NASA TM X-3295, DOT-TSCOST-75-40. The 1975 Ride Quality Symposium, November 1975, pp. 373–385, (see N76-16754 07-53), 76N16770. The present study was done aboard a special aircraft able to effect translations through the center of gravity with a minimum of pitch and roll. The aircraft was driven through controlled motions by an on-board analog computer. The input signal was selectively filtered Gaussian noise whose power spectra approximated that of natural turbulence. This input, combined with the maneuvering capabilities of this aircraft, resulted in an extremely realistic simulation of turbulent flight. The test flights also included varying bank angles during turns. Subjects were chosen from among NASA Flight Research Center personnel. They were all volunteers, were given physical examinations, and were queried about their attitudes toward flying before final selection. In profile, they were representative of the general flying public. Data from this study include (1) a basis for comparison with previous commercial flights, that is, motion dominated by vertical acceleration, (2) extension to motion dominated by lateral acceleration, and (3) evaluation of various bank angles. 868. Kordes, E. E.; and *Curtis, A. R.: Results of NASTRAN Modal Analyses and Ground Vibration Tests on the YF-12A Airplane. ASME Paper 75-WA/AERO-8, presented at the American Society of Mechanical Engineers, Winter Annual Meeting, Houston, Texas, November 30–December 4, 1975, 76A21855, #. The YF-12A aircraft, a delta-winged vehicle powered by two jet engines, was utilized in an investigation of the structural dynamic characteristics of a large, flexible, supersonic research vehicle. A large NASA structural analysis (NASTRAN) finite-element model was used to compute the ten lowest frequency symmetric and ten lowest frequency antisymmetric modes for the YF-12A aircraft. The results of the analysis were compared with experimental data obtained in a ground vibration test conducted with the completed aircraft. It was found that the finite-element structural model employed provides an adequate prediction of the dynamic behavior of the aircraft structure in the case of basic wing and body modes. *Lockheed-California Co., Burbank, California. 869. Manke, J. A.; and *Love, M. V.: X-24B Flight Test Program. Presented at the Nineteenth Society of Experimental Test Pilots Symposium, Beverly Hills, California, September 24–27, 1975, Society of Experimental Test Pilots, Technical Review, Vol. 12, No. 4, 1975, pp. 129–154, 76A18659. The X-24B is an air launched, rocket powered research aircraft. A number of its design features constitute a tradeoff between aerodynamics and heating considerations. A vehicle description is given and test program objectives are discussed along with operational procedures and aspects of energy management. Attention is also given to X-24B handling 147 qualities, approach and landing, wind tunnel data and simulation, and proposed X-24C vehicle requirements. *USAF, Edwards AFB, California. below the trailing edge plane were averaged for calculating section drag coefficients for flights at low dynamic pressures. PIK-20E Sailplane X-24B Lifting Body EC75-4642 EC91-504-1 870. *Hoffman, E. L.; **Payne, L.; and Carter, A. L.: Fabrication Methods for YF-12 Wing Panels for the Supersonic Cruise Aircraft Research Program. Materials Review ‘75, Proceedings of the Seventh National Technical Conference, Albuquerque, New Mexico, October 14–16, 1975, pp. 68–82, 76A15157. Advanced fabrication and joining processes for titanium and composite materials are being investigated by NASA to develop technology for the Supersonic Cruise Aircraft Research (SCAR) Program. With Lockheed-ADP as the prime contractor, full-scale structural panels are being designed and fabricated to replace an existing integrally stiffened shear panel on the upper wing surface of the NASA YF-12 aircraft. The program involves ground testing and Mach 3 flight testing of full-scale structural panels and laboratory testing of representative structural element specimens. Fabrication methods and test results for weldbrazed and Rohrbond titanium panels are discussed. The fabrication methods being developed for boron/aluminum, Borsic/aluminum, and graphite/polyimide panels are also presented. *NASA Langley Research Center, Hampton, Virginia. **Lockheed-California Co., Sunland, California. 871. Saltzman, E. J.: Use of a Pitot Probe for Determining Wing Section Drag in Flight. NASA CR-145627. Kansas University Proceedings of the NASA, Industry, University, General Aviation Drag Reduction Workshop, 1975, pp. 171–189, (see N76-10997 02-01), 76N11010. A wake traversing probe was used to obtain section drag and wake profile data from the wing of a sailplane. The transducer sensed total pressure defect in the wake as well as freestream total pressure on both sides of the sensing element when the probe moved beyond the wake. Profiles of wake total pressure defects plotted as a function of distance above and 148 872. Iliff, Kenneth W.: An Airplane Application of System Identification in the Presence of State Noise. New Directions in Signal Processing in Communications and Control, NATO Advanced Study Institutes Series, Series E: Applied Sciences - No. 12, 1975. (See also 815.) A maximum likelihood estimator for a linear system with state and observation noise is developed to determine unknown aircraft coefficients from flight data in the presence of turbulence (state noise). The formulation of the algorithm is presented briefly. The linear equations for an aircraft in atmospheric turbulence are defined. The effectiveness and accuracy of the technique are assessed by first applying it to simulated flight data, in which the true parameter values are known, then the actual flight data obtained in turbulence. A complete set of aircraft coefficients is obtained as well as an estimate of the turbulence time history. The validity of the estimated state noise and of the estimated coefficients is tested. The feasibility of using the algorithm for defining an adaptive control law to alleviate the effects of turbulence on the aircraft is discussed. 1976 Technical Publications 873. *Jacobsen, R. T.; *Stewart, R. B.; **Crain, R. W., Jr.; †Rose, G. L.; and Myers, A. F.: A Method for the Selection of a Functional Form for a Thermodynamic Equation of State Using Weighted Linear Least Squares Stepwise Regression. Proceedings, Cryogenic Engineering Conference, Kingston, Ontario, Canada, July 22–25, 1975, pp. 532–537, Plenum Press, New York, New York, 1976, 77A42171. A method was developed for establishing a rational choice of the terms to be included in an equation of state with a large number of adjustable coefficients. The methods presented were developed for use in the determination of an equation of state for oxygen and nitrogen. However, a general application of the methods is possible in studies involving the determination of an optimum polynomial equation for fitting a large number of data points. The data considered in the least squares problem are experimental thermodynamic pressuredensity-temperature data. Attention is given to a description of stepwise multiple regression and the use of stepwise regression in the determination of an equation of state for oxygen and nitrogen. *University of Idaho, Moscow, Idaho. **Washington State University, Pullman, Washington. †McKellip Engineering, Inc., Boise, Idaho. 874. Edwards, J. W.: A FORTRAN Program for the Analysis of Linear Continuous and Sample-Data Systems. NASA TM X-56038, January 1976, 76N18823, #. A FORTRAN digital computer program which performs the general analysis of linearized control systems is described. State variable techniques are used to analyze continuous, discrete, and sampled data systems. Analysis options include the calculation of system eigenvalues, transfer functions, root loci, root contours, frequency responses, power spectra, and transient responses for open- and closed-loop systems. A flexible data input format allows the user to define systems in a variety of representations. Data may be entered by inputting explicit data matrices or matrices constructed in user written subroutines, by specifying transfer function block diagrams, or by using a combination of these methods. 875. Holleman, E. C.: Summary of Flight Tests to Determine the Spin and Controllability Characteristics of a Remotely Piloted, Large-Scale (3/8) Fighter Airplane Model. NASA TN D-8052, H-889, January 1976, 76N17156, #. An unpowered, large, dynamically scaled airplane model was test flown by remote pilot to investigate the stability and controllability of the configuration at high angles of attack. The configuration proved to be departure/spin resistant; however, spins were obtained by using techniques developed on a flight support simulator. Spin modes at high and medium high angles of attack were identified, and recovery techniques were investigated. A flight support simulation of the airplane model mechanized with low speed wind tunnel data over an angle of attack range of ± 90 deg. and an angle of sideslip range of ± 40 deg. provided insight into the effects of altitude, stability, aerodynamic damping, and the operation of the augmented flight control system on spins. Aerodynamic derivatives determined from flight maneuvers were used to correlate model controllability with two proposed departure/spin design criteria. 876. Iliff, K. W.; Maine, R. E.; and Shafer, M. F.: Subsonic Stability and Control Derivatives for an Unpowered, Remotely Piloted 3/8-Scale F-15 Airplane Model Obtained From Flight Test. NASA TN D-8136, H-905, January 1976, 76N15176, #. 149 In response to the interest in airplane configuration characteristics at high angles of attack, an unpowered remotely piloted 3/8-scale F-15 airplane model was flight tested. The subsonic stability and control characteristics of this airplane model over an angle of attack range of –20 to 53 deg are documented. The remotely piloted technique for obtaining flight test data was found to provide adequate stability and control derivatives. The remotely piloted technique provided an opportunity to test the aircraft mathematical model in an angle of attack regime not previously examined in flight test. The variation of most of the derivative estimates with angle of attack was found to be consistent, particularly when the data were supplemented by uncertainty levels. 877. White, K. C.; Lasagna, P. L.; and Putnam, T. W.: Preliminary Measurements of Aircraft Airframe Noise With the NASA CV-990 Aircraft. NASA TM X-73116, A-6506, January 1976, 76N26145, #. Flight tests were conducted in a CV-990 jet transport with engines at idle power to investigate aircraft airframe noise. Test results showed that airframe noise was measured for the aircraft in the landing configuration. The results agreed well with the expected variation with the fifth power of velocity. For the aircraft in the clean configuration, it was concluded that airframe noise was measured only at higher airspeeds with engine idle noise present at lower speeds. The data show that landing gear and flaps make a significant contribution to airframe noise. 878. Hedgley, D. R., Jr.: An Exact Transformation From Geocentric to Geodetic Coordinates for Nonzero Altitudes. NASA TR R-458, H-909, March 1976, 76N19836, #. An exact method for the nonzero altitude transformation from geocentric to geodetic coordinates is derived. The method is mathematically general and should serve as a primary standard. 879. Albers, J. A.: Status of the NASA YF-12 Propulsion Research Program. NASA TM X-56039, H-935, March 1976, 76N19152, #. The YF-12 research program was initiated to establish a technology base for the design of an efficient propulsion system for supersonic cruise aircraft. The major technology areas under investigation in this program are inlet design analysis, propulsion system steady-state performance, propulsion system dynamic performance, inlet and engine control systems, and airframe/propulsion system interactions. The objectives, technical approach, and status of the YF-12 propulsion program are discussed. Also discussed are the results obtained to date by the NASA Ames, Lewis, and Dryden research centers. The expected technical results and proposed future programs are also given. Propulsion system configurations are shown. 880. Sim, Alex G.: A Correlation Between Flight Determined Derivatives and Wind-Tunnel Data for the X-24B Research Aircraft. NASA TM SX-3371, March 1976, (republished as NASA TM-113084, August 1997). Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight by using a modified maximum likelihood estimation method. Data were obtained over a mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5° to 15.7°. Data are presented for a subsonic and transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between flight data and the wind-tunnel predictions is presented and discussed. procedures used for editing the data and for overall analysis are also discussed. 882. Petersen, K. L.: Evaluation of an EnvelopeLimiting Device Using Simulation and Flight Test of a Remotely Piloted Research Vehicle. NASA TN D-8216, H-914, April 1976, 76N21218, #. The operating characteristics of a nonlinear envelopelimiting device were investigated at extreme flight conditions by using a real time digital aircraft spin simulation and flight tests of a scale model remotely piloted research vehicle. A digital mechanization of the F-15 control system, including the stall inhibiter, was used in the simulation and in the control system of the scale model. The operational characteristics of the stall inhibiter and the effects of the stall inhibiter on the spin susceptibility of the airplane were investigated. 883. Kurkowski, R. L.; Barber, M. R.; and Garodz, L. J.: Characteristics of Wake Vortex Generated by a Boeing 727 Jet Transport During Two-Segment and Normal ILS Approach Flight Paths. NASA TN D-8222, A-6208, April 1976, 76N21175, #. A series of flight tests was conducted to evaluate the vortex wake characteristics of a Boeing 727 (B727-200) aircraft during conventional and two-segment ILS approaches. Twelve flights of the B727, which was equipped with smoke generators for vortex marking, were flown and its vortex wake was intentionally encountered by a Lear Jet model 23 (LR-23) and a Piper Twin Comanche (PA-30). Location of the B727 vortex during landing approach was measured using a system of photo-theodolites. The tests showed that at a given separation distance there were no readily apparent differences in the upsets resulting from deliberate vortex position of the landing configuration vortices showed that they tended to descend approximately 91 m (300 ft) below the flight path of the B727. The flaps of the B727 have a dominant effect on the character of the trailed wake vortex. The clean wing produces a strong, concentrated vortex but as the flaps are lowered, the vortex system becomes more diffuse. Pilot opinion and roll acceleration data indicate that 4.5 nmi would be a minimum separation distance at which roll control of light aircraft (less than 5,670 kg (12,500 lb) could be maintained during parallel encounters of the B727’s landing configuration wake. This minimum separation distance is generally in scale with results determined from previous tests of other aircraft using the small roll control criteria. 884. Putnam, T. W.; and Burcham, F. W.: Business Jet Approach Noise Abatement Techniques—Flight Test Results. SAE Paper 760463, presented at the Society of Automotive Engineers, Business Aircraft Meeting, Wichita, Kansas, April 6, 1976, 76A31961. 150 980046 X-24B Lifting Body, Three-View Drawing 881. Iliff, K. W.; and Maine, R. E.: Practical Aspects of a Maximum Likelihood Estimation Method to Extract Stability and Control Derivatives From Flight Data. NASA TN D-8209, H-908, April 1976, 76N23272, #. A maximum likelihood estimation method was applied to flight data and procedures to facilitate the routine analysis of a large amount of flight data were described. Techniques that can be used to obtain stability and control derivatives from aircraft maneuvers that are less than ideal for this purpose are described. The techniques involve detecting and correcting the effects of dependent or nearly dependent variables, structural vibration, data drift, inadequate instrumentation, and difficulties with the data acquisition system and the mathematical model. The use of uncertainty levels and multiple maneuver analysis also proved to be useful in improving the quality of the estimated coefficients. The Operational techniques for reducing approach noise from business jet aircraft were evaluated in flight by measuring the noise generated by five such aircraft during modified approaches. Approaches with 4-deg glide slopes were approximately 4.0 EPNdB quieter than approaches with standard 3-deg glide slopes. Noise reductions for low-drag 3-deg approaches varied widely among the airplanes tested; the fleet-weighted reduction was 8.5 EPNdB. Two-segment approaches resulted in noise reductions of 7.0 EPNdB to 8.5 EPNdB 3 nautical miles and 5 nautical miles from touchdown. Pilot workload increased progressively for the 4-deg, low-drag 3-deg, and two-segment approach. 885. *Lockenour, J. L.; and Layton, G. P.: RPRV Research Focus on HiMAT. Astronautics and Aeronautics, Vol. 14, April 1976, pp. 36–41, 76A25721, #. A review is presented of the F-15 Remotely Piloted Research Vehicle (RPRV) project. The F-15 RPRV is air-launched from a B-52 at 50,000 ft. Following launch a series of research maneuvers are performed during an unpowered descent to a recovery altitude. Another RPRV program considered is the Highly Maneuverable Aircraft Technology (HiMAT) program. This program is designed to use RPRVs to speed the technology transition from wind tunnel to flight and to reduce the cost of aeronautical experiments. It is pointed out that HiMAT will make extensive use of composite materials. *USAF, Flight Dynamics Laboratory, Wright-Patterson AFB, Ohio. 886. Layton, G. P.: A New Experimental Flight Research Technique: The Remotely Piloted Airplane. AGARD CP-187. Flight/Ground Testing Facilities Correlation, April 1976, (see N76-25266 16-09), 76N25287. The results obtained so far with a remotely piloted research vehicle (RPRV) using a 3/8 scale model of an F-15 airplane, to determine the usefulness of the RPRV testing technique in high risk flight testing, including spin testing, were presented. The program showed that the RPRV technique, including the use of a digital control system, is a practical method for obtaining flight research data. The spin, stability, and control data obtained with the 3/8-scale model also showed that predictions based on wind-tunnel tests were generally reasonable. 887. Gee, Shu W.: Model Aircraft Technology Applications at NASA. R/C Modeler Magazine, April 1976. 888. Montoya, Earl J.: Aeronautics. Presented at NASA Symposium on Aeronautics and Space Technology, New Mexico State University, Las Cruces, New Mexico, April 21–23, 1976. This paper compares the NASA organization to the National Football League and the importance to both of training. 889. Sakamoto, G. M.: Aerodynamic Characteristics of a Vane Flow Angularity Sensor System Capable of Measuring Flight Path Accelerations for the Mach Number Range From 0.40 to 2.54. NASA TN D-8242, H-900, May 1976, 76N23257, #. The aerodynamic characteristics of the angle of attack vane and the angle of sideslip vane are summarized. The test conditions ranged in free stream Mach number from 0.40 to 2.54, in angle of attack from –2 deg to 22 deg, in angle of sideslip from –2 deg to 12 deg, and in Reynolds number from 590,000 per meter to 1.8 million per meter. The results of the wind tunnel investigation are compared with results obtained with similar vane configurations. Comparisons with a NACA vane configuration are also made. In addition, wind tunnelderived upwash for the test installation is compared with analytical predictions. 890. *Gillingham, K. K.; and Winter, W. R.: Physiologic and Anti-G Suit Performance Data From YF-16 Flight Tests. Aviation, Space, and Environmental Medicine, Vol. 47, June 1976, pp. 672–673, 76A37075. Biomedical data were collected during high-G portions of 11–16 test flights. Test pilots monitored revealed increased respiratory rate and volume, decreased tidal volume, and increased heart rate at higher G levels, with one pilot exhibiting various cardiac arrhythmias. Anti-G suit inflation and pressurization lags varied inversely with G-onset rate, and suit pressurization slope was near the design value. *USAF, School of Aerospace Medicine, Brooks AFB, Texas. 891. Iliff, K. W.: Maximum Likelihood Estimates of Lift and Drag Characteristics Obtained From Dynamic Aircraft Maneuvers. Proceedings, 3rd AIAA Atmospheric Flight Mechanics Conference, Arlington, Texas, June 7–9, 1976, pp. 137–150, (see A76-36901 17-08), 76A36916, #. A maximum likelihood estimation method for obtaining lift and drag characteristics from dynamic flight maneuvers was investigated. This paper describes the method and compares the estimates of lift and drag obtained by using the method with estimates obtained from wind-tunnel tests and from established methods for obtaining estimates from flight data. In general, the lift and drag coefficients extracted from dynamic flight maneuvers by the maximum likelihood estimation technique are in good agreement with the estimates obtained from the wind-tunnel tests and the other methods. When maneuvers that met the requirements of both flight methods were analyzed, the results of each method were nearly the same. The maximum likelihood estimation technique showed promise in terms of estimating lift and drag characteristics from dynamic flight maneuvers. Further 151 studies should be made to assess the best mathematical model and the most desirable type of dynamic maneuver to get the highest quality results from this technique. 892. Petersen, K. L.: Remotely Piloted Research Vehicle Evaluation of Advanced Control System Effects on Spins. Proceedings, 3rd AIAA Atmospheric Flight Mechanics Conference, Arlington, Texas, June 7–9, 1976, pp. 55–64, (see A76-36901 17-08), 76A36907, #. Special functions of an advanced control system were investigated for effects on spin entries and recoveries utilizing a 3/8-scale model of the F-15 airplane as a remotely piloted research vehicle (RPRV). Telemetry uplinks and downlinks were used with a ground-based digital computer to mechanize the RPRV control system for spin tests in flight. Results from the model RPRV flight tests and from a real time digital spin simulation were used to evaluate the F-15 stall inhibiter and an automatic spin recovery system developed for the RPRV model. 893. Iliff, K. W.: Estimation of Characteristics and Stochastic Control of an Aircraft Flying in Atmospheric Turbulence. Proceedings, 3rd AIAA Atmospheric Flight Mechanics Conference, Arlington, Texas, June 7–9, 1976, pp. 26–38, (see A76-36901 17-08), 76A36905, #. An adaptive control technique to improve the flying qualities of an aircraft in turbulence was investigated. The approach taken was to obtain maximum likelihood estimates of the unknown coefficients of the aircraft system and then, using these estimates along with the separation principle, to define the stochastic optimal control. The maximum likelihood estimation technique that accounted for the effects of turbulence provided good estimates of the unknown coefficients and of the turbulence. The assessment of the stochastic optimal control based on the maximum likelihood estimates showed that the desired effects were attained for the regulator problem of minimizing pitch angle and the tracking problem of requiring normal acceleration to follow the pilot input. 894. Burcham, F. W., Jr.; and *Batterton, P. G.: Flight Experience With a Digital Integrated Propulsion Control System on an F-111E Airplane. AIAA Paper 76-653, presented at 12th AIAA and SAE Annual Propulsion Conference, Palo Alto, California, July 26–29, 1976, 76A42411, #. A digital integrated propulsion control system (IPCS) installed in the left side of an F-111 E aircraft was tested in flight. The F-111 aircraft was selected for the IPCS program because it incorporated a variable geometry inlet and an afterburning turbofan engine and had two engines, one of which could remain in the normal configuration to ensure flight safety. Flight data were compared with results of tests run in an altitude test chamber. The digital system was found 152 to be capable of duplicating the standard engine and inlet control systems. Instabilities such as inlet buzz and afterburner rumble were detected and controlled. The usefulness of an altitude chamber for developing a software and testing hardware was proven. The flexibility of IPCS was demonstrated when an autothrottle, an in-flight thrust calculation, and a coannular noise study capability were added at the end of the flight tests. *NASA Lewis Research Center, Cleveland, Ohio. F-111E IPCS Airplane ECN-4359 895. *Hersh, A. S.; Putnam, T. W.; Lasagna, P. L.; and Burcham, F. W., Jr.: Semi-Empirical Airframe Noise Prediction Model. AIAA Paper 76-527, presented at 3rd AIAA Annual Aero-Acoustics Conference, Palo Alto, California, July 20–23, 1976, (see also NASA TM-56041.) 76A38052, #. (See also 912.) A semi-empirical maximum overall sound pressure level (OASPL) airframe noise model was derived. The noise radiated from aircraft wings and flaps was modeled by using the trailing-edge diffracted quadrupole sound theory derived by Ffowcs Williams and Hall. The noise radiated from the landing gear was modeled by using the acoustic dipole sound theory derived by Curle. The model was successfully correlated with maximum OASPL flyover noise measurements obtained at the NASA Dryden Flight Research Center for three jet aircraft—the Lockheed JetStar, the Convair 990, and the Boeing 747 aircraft. *Hersh Acoustical Engineering, Chatsworth, California. 896. Anon.: Advanced Control Technology and Its Potential for Future Transport Aircraft. NASA TM X-3409, H-904. Presented at a Symposium on Adv. Control Technol., Los Angeles, California, July 9–11, 1974, August 1976, 76N31135, #. 897. *Lange, R. H.; and Deets, D. A.: Study of an ACT Demonstrator With Substantial Performance Improvements Using a Redesigned JetStar. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 3–35, (see N76-31135 22-01), 76N31159. The feasibility was studied of modifying a JetStar airplane into a demonstrator of benefits to be achieved from incorporating active control concepts in the preliminary design of transport type aircraft. Substantial benefits are shown in terms of fuel economy and community noise by virtue of reduction in induced drag through use of a high aspect ratio wing which is made possible by a gust alleviation system. An intermediate configuration was defined which helps to isolate the benefits produced by active controls technology from those due to other configuration variables. *Lockheed-Georgia Company, Marietta, Georgia. 898. Jarvis, C. R.: An Overview of NASA’s Digital FlyBy-Wire Technology Development Program. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 93–103, (see N76-31135 22-01), 76N31140. (See also 839.) The feasibility of using digital fly by wire systems to control aircraft was demonstrated by developing and flight testing a single channel system, which used Apollo hardware, in an F-8C test airplane. This is the first airplane to fly with a digital fly by wire system as its primary means of control and with no mechanical reversion capability. The development and flight test of a triplex digital fly by wire system, which will serve as an experimental prototype for future operational digital fly by wire systems, are underway. 899. Deets, D. A.: Design and Development Experience With a Digital Fly-By-Wire Control System in an F-8C Airplane. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 105–132, (see N76-31135 22-01), 76N31141. (See also 840.) To assess the feasibility of a digital fly by wire system, the mechanical flight control system of an F-8C airplane was replaced with a digital primary system and an analog backup system. The Apollo computer was used as the heart of the primary system. This paper discusses the experience gained during the design and development of the system and relates it to active control systems that are anticipated for future civil transport applications. 900. Lock, W. P.; Petersen, W. R.; and *Whitman, G. B.: Mechanization of and Experience With a Triplex Fly-ByWire Backup Control System. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 133–163, (see N76-31135 22-01), 76N31142. (See also 841.) 153 A redundant three axis analog control system was designed and developed to back up a digital fly by wire control system for an F-8C airplane. The mechanization and operational experience with the backup control system, the problems involved in synchronizing it with the primary system, and the reliability of the system are discussed. The backup control system was dissimilar to the primary system, and it provided satisfactory handling through the flight envelope evaluated. Limited flight tests of a variety of control tasks showed that control was also satisfactory when the backup control system was controlled by a minimum displacement (force) side stick. The operational reliability of the F-8 digital fly by wire control system was satisfactory, with no unintentional downmodes to the backup control system in flight. The ground and flight reliability of the system’s components is discussed. *Sperry Flight Systems Div., Phoenix, Arizona. 901. Szalai, K. J.: Flight Test Experience With the F-8 Digital Fly-By-Wire System. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technology and Its Potential for Future Transport Aircraft, August 1976, pp. 199–252, (see N76-31135 22-01), 76N31144. (See also 843.) Flight test results of the F-8 digital fly by wire control system are presented and the implications for application to active control technology are discussed. The F-8 DFBW system has several of the attributes of proposed ACT systems, so the flight test experience is helpful in assessing the capabilities of those systems. Topics of discussion include the predicted and actual flight performance of the control system, assessments of aircraft flying qualities and other piloting factors, software management and control, and operational experience. 902. Krier, G. E.: A Pilot’s Opinion of the F-8 Digital Fly-By-Wire Airplane. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 253–267, (see N76-31135 22-01), 76N31145. (See also 844.) The handling qualities of the F-8 digital fly by wire airplane are evaluated by using the Cooper-Harper rating scale. The reasons for the ratings are given, as well as a short description of the flying tasks. It was concluded that the handling qualities of the airplane were good in most situations, although occasional ratings of unsatisfactory were given. 903. Berry, D. T.; and Schweikhard, W. G.: Potential Benefits of Propulsion and Flight Control Integration for Supersonic Cruise Vehicles. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 433–452, (see N76-31135 22-01), 76N31152. Typical airframe/propulsion interactions such as Mach/ altitude excursions and inlet unstarts are reviewed. The improvements in airplane performance and flight control that can be achieved by improving the interfaces between propulsion and flight control are estimated. A research program to determine the feasibility of integrating propulsion and flight control is described. This program includes analytical studies and YF-12 flight tests. 904. Anon.: The ACT Transport: Panacea for the 80’s Or Designer’s Illusion (Panel Discussion). NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 805–827, (see N76-31135 22-01), 76N31169. A panel discussion was held which attempted to make an objective and pragmatic assessment of the standing of active control technology. The discussion focused on the standing of active control technology relative to civil air transport applications, the value as opposed to the cost of the projected benefits, the need for research, development, and demonstration, the role of government and industry in developing the technology, the major obstacles to its implementation, and the probable timing of the full utilization of active control technology in commercial transportation. An edited transcription of the prepared statements of the panel members and the subsequent open discussion between the panel and the audience is presented. 905. *Plumer, J. A.; **Malloy, W. A.; and Craft, J. B.: The Effects of Lightning on Digital Flight Control Systems. NASA TM X-3409. NASA Dryden Flight Research Center Advanced Control Technol. and Its Potential for Future Transport Aircraft, August 1976, pp. 989–1008, (see N76-31135 22-01), 76N31176. (See also 842.) Present practices in lightning protection of aircraft deal primarily with the direct effects of lightning, such as structural damage and ignition of fuel vapors. There is increasing evidence of troublesome electromagnetic effects, however, in aircraft employing solid-state microelectronics in critical navigation, instrumentation and control functions. The potential impact of these indirect effects on critical systems such as digital fly by wire (DFBW) flight controls was studied. The results indicate a need for positive steps to be taken during the design of future fly by wire systems to minimize the possibility of hazardous effects from lightning. *General Electric, Corporate Research and Development, Schenectady, New York. **General Motors Corporation, Detroit, Michigan. 906. Szalai, K. J.; *Felleman, P. G.; **Gera, J.; and †Glover, R. D.: Design and Test Experience With a Triply Redundant Digital Fly-By-Wire Control System. AIAA Paper 76-1911. Proceedings, Guidance and Control 154 Conference, San Diego, California, August 16–18, 1976, 76A41491, #. A triplex digital fly-by-wire flight control system was developed and then installed in a NASA F-8C aircraft to provide fail-operative, full authority control. Hardware and software redundancy management techniques were designed to detect and identify failures in the system. Control functions typical of those projected for future actively controlled vehicles were implemented. This paper describes the principal design features of the system, the implementation of computer, sensor, and actuator redundancy management, and the ground test results. An automated test program to verify sensor redundancy management software is also described. *Draper Laboratory, Incorporated, Cambridge, Massachusetts. **NASA Langley Research Center, Hampton, Virginia. †NASA Johnson Space Center, Houston, Texas. 907. Gee, S. W.; *Jenks, G. E.; *Roskam, J.; and **Stone, R. L.: Flight Test Evaluation of a Separate Surface Attitude Command Control System on a Beech 99 Airplane. AIAA Paper 76-1991. Proceedings, Guidance and Control Conference, San Diego, California, August 16–18, 1976, 76A41489. (See also 914.) A joint NASA/university/industry program was conducted to flight evaluate a potentially low cost separate surface implementation of attitude command in a Beech 99 airplane. Saturation of the separate surfaces was the primary cause of many problems during development. Six experienced professional pilots made simulated instrument flight evaluations in light-to-moderate turbulence. They were favorably impressed with the system, particularly with the elimination of control force transients that accompanied configuration changes. For ride quality, quantitative data showed that the attitude command control system resulted in all cases of airplane motion being removed from the uncomfortable ride region. *University of Kansas, Lawrence, Kansas. **Beech Aircraft Corp., Wichita, Kansas. 908. Schweikhard, W. G.; Gilyard, G. B.; *Talbot, J. E.; and *Brown, T. W.: Effects of Atmospheric Conditions on the Operating Characteristics of Supersonic Cruise Aircraft. IAF Paper 76-112, presented at the International Astronautical Federation, Twenty-Seventh International Astronautical Congress, Anaheim, California, October 10–16, 1976, 77A10912, #. Since for maximum range a supersonic transport must cruise near its maximum Mach number, accurate flight control is needed, especially when severe atmospheric transients are encountered. This paper describes atmospheric transients that have been encountered by the XB-70, YF-12, and Concorde aircraft during supersonic flights and the ensuing responses of the aircraft propulsion and flight control systems. It was found that atmospheric conditions affected these supersonic cruise vehicles in much the same way, with minor differences according to the type of propulsion and flight control system. Onboard sensors are sufficiently accurate to provide data on the atmosphere, including turbulence over the route, that are accurate enough for entry in the climatic record and for use as inputs to the control systems. Nominal atmospheric transients can be satisfactorily controlled, but some problems remain for extreme cases. *British Aircraft Company, Bristol, England. 909. Baer, J. L.; Holzman, J. K.; and Burcham, F. W., Jr.: Procedures Used in Flight Tests of an Integrated Propulsion Control System on an F-111E Airplane. SAE Paper 760933, presented at the Society of Automotive Engineers, Aerospace Engineering and Manufacturing Meeting, San Diego, California, November 29–December 2, 1976, 77A28238. A digital integrated propulsion control system (IPCS) was tested on an F-111E airplane. The IPCS provided full authority control of the left inlet and the TF30 afterburning engine. Supersonic test conditions were of primary interest. The operational procedures and maneuvers developed for IPCS evaluation, displays for test monitoring and data acquisition, flight safety, and problems encountered are discussed. The software refinements that made modifications to standard flight test procedures necessary are described. The flexibility of digital control and the ways software was used to overcome hardware deficiencies are discussed. Application of these procedures to a typical IPCS flight is described. 910. Enevoldson, E.: The X-24B, Spot Landing. Soaring, Vol. 40, No. 11, November 1976. 911. Foster, J. D.; and Lasagna, P. L.: Flight-Test Measurement of the Noise Reduction of a Jet Transport Delayed Flap Approach Procedure. NASA TM X-73172, A-6775, December 1976, 77N20078, #. A delayed flap approach procedure was flight tested using the NASA CV-990 airplane to measure and analyze the noise produced beneath the flight path. Three other types of landing approaches were also flight tested to provide a comparison of the noise reduction benefits to the delayed flap approach. The conventional type of approach was used as a baseline to compare the effectiveness of the other approaches. The decelerating approach is a variation of the delayed flap approach. A detailed comparison of the ground perceived noise generated during the approaches is presented. For this comparison, the measured noise data were normalized to compensate for variations in aircraft weight and winds that occurred during the flight tests. The data show that the 155 reduced flap approach offers some noise reduction, while the delayed flap and decelerating approaches offer significant noise reductions over the conventional approach. 912. *Hersh, A. S.; Burcham, F. W., Jr.; Putnam, T. W.; and Lasagna, P. L.: Semiempirical Airframe Noise Prediction Model and Evaluation With Flight Data. NASA TM X-56041, H-951, December 1976, 77N13791, #. (See also 895.) A semiempirical maximum overall sound pressure level (OASPL) airframe noise model was derived. Noise radiated from aircraft wings was modeled on the trailing edge diffractes quadrupole sound theory. The acoustic dipole sound theory was used to model noise from the landing gear. The model was correlated with maximum OASPL flyover noise measurements obtained for three jet aircraft. One third octave band sound pressure level flyover data was correlated and interpreted. *Hersh Acoustical California. Engineering, Westlake Village, 913. Berry, D. T.; and Gilyard, G. B.: A Review of Supersonic Cruise Flight Path Control Experience With the YF-12 Aircraft. NASA SP-416. Aircraft Safety and Operating Problems, 1976, pp. 147–164, (see N77-18081 09-03), 77N18089, #. Flight research with the YF-12 aircraft indicates that solutions to many handling qualities problems of supersonic cruise are at hand. Airframe/propulsion system interactions in the dutch roll mode can be alleviated by the use of passive filters or additional feedback loops in the propulsion and flight control systems. Mach and altitude excursions due to atmospheric temperature fluctuations can be minimized by the use of a cruise autothrottle. Autopilot instabilities in the altitude hold mode have been traced to angle of attacksensitive static ports on the compensated nose boom. For the YF-12, the feedback of high-passed pitch rate to the autopilot resolves this problem. Manual flight path control is significantly improved by the use of an inertial rate of climb display in the cockpit. 914. Gee, S. W.; *Jenks, G. E.; *Roskam, J.; and **Stone, R. L.: Flight Test Evaluation of a Separate Surface Attitude Command Control System on a Beech 99 Airplane. NASA SP-416. Aircraft Safety and Operating Problems, 1976, pp. 121–146, (see N77-18081 09-03), 77N18088, #. (See also 907.) A joint NASA/university/industry program was conducted to flight evaluate a potentially low cost separate surface implementation of attitude command in a Beech 99 airplane. Saturation of the separate surfaces was the primary cause of many problems during development. Six experienced professional pilots who made simulated instrument flight evaluations experienced improvements in airplane handling qualities in the presence of turbulence and a reduction in pilot workload. For ride quality, quantitative data show that the attitude command control system results in all cases of airplane motion being removed from the uncomfortable ride region. *University of Kansas, Lawrence, Kansas. **Beech Aircraft Corp., Wichita, Kansas. 915. Albers, J. A.; and Olinger, F. V.: YF-12 Propulsion Research Program and Results. NASA CP-001. Proceedings of the SCAR Conf., Part 1, 1976, pp. 417–456, (see N77-17996 09-01), 77N18017, #. The objectives and status of the propulsion program, along with the results acquired in the various technology areas, are discussed. The instrumentation requirements for and experience with flight testing the propulsion systems at high supersonic cruise are reported. Propulsion system performance differences between wind tunnel and flight are given. The effects of high frequency flow fluctuations (transients) on the stability of the propulsion system are described, and shock position control is evaluated. 916. Reukauf, P. J.; and Burcham, F. W., Jr.: Propulsion System/Flight Control Integration for Supersonic Aircraft. NASA CP-001. Proceedings of the SCAR Conf., Part 1, 1976, pp. 281–302, (see N77-17996 09-01), 77N18010, #. Digital integrated control systems are studied. Such systems allow minimization of undesirable interactions while maximizing performance at all flight conditions. One such program is the YF-12 cooperative control program. The existing analog air data computer, autothrottle, autopilot, and inlet control systems are converted to digital systems by using a general purpose airborne computer and interface unit. Existing control laws are programmed and tested in flight. Integrated control laws, derived using accurate mathematical models of the airplane and propulsion system in conjunction with modern control techniques, are tested in flight. Analysis indicates that an integrated autothrottle autopilot gives good flight path control and that observers are used to replace failed sensors. 917. Berry, D. T.; Mallick, D. L.; and Gilyard, G. B.: Handling Qualities Aspects of NASA YF-12 Flight Experience. NASA CP-001. Proceedings of the SCAR Conf., Part 1, 1976, pp. 193–214, (see N77-17996 09-01), 77N18007, #. The handling qualities of the YF-12 airplane as observed during NASA research flights over the past five years were reviewed. Aircraft behavior during takeoff, acceleration, climb, cruise, descent, and landing are discussed. Pilot comments on the various flight phases and tasks are presented. Handling qualities parameters such as period, 156 damping, amplitude ratios, roll-yaw coupling, and flight path response sensitivity are compared to existing and proposed handling qualities criteria. The influence of the propulsion systems, stability augmentation, autopilot systems, atmospheric gusts, and temperature changes are also discussed. YF-12 experience correlates well with flying qualities criteria, except for longitudinal short period damping, where existing and proposed criteria appear to be more stringent than necessary. 918. Gillingham, K. K.; and Winter, W. R.: Physiologic and Anti-G Suit Performance Data From YF-16 Flight Tests. NASA TM X-74617, AD-A032357, 1976, 77N77316, #. 919. Donlan, C. J.; and Weil, J.: Characteristics of Swept Wings at High Speeds, 30 January 1952. Collected Works of Charles J. Donlan, 1976, (see N77-29059 20-01), 77N29078, #. Some results of recent swept wing investigations are presented, that were undertaken to determine the effects of thickness and thickness distribution, camber and twist, noseflap deflection, and devices or fixes for improving the wing pitching moment characteristics at high lift coefficients. 920. Montoya, E. J.; and *Faye, A. E., Jr.: NASA Participation in the AMST Program. NASA Langley Res. Center Powered-Lift Aerodyn. and Acoustics, 1976, pp. 465–478, (see N78-24046 15-02), 78N24075, #. The objectives of the NASA Advanced Medium STOL Transport Experiments Program are discussed and several of the NASA experiments currently implemented and conducted on the YC-14 and YC-15 prototype aircraft are described. Emphasis is placed on experiments related to powered lift aerodynamics and acoustics. *NASA Ames Research Center, Moffett Field, California. 921. *Bales, T. T.; *Hoffman, E. L.; *Payne, L.; and Carter, A. L.: Fabrication and Evaluation of Advanced Titanium and Composite Structural Panels. NASA Proceedings of the SCAR Conf., Part. 2, 1976, (see N7718019 09-01), 77N18034, #. Advanced manufacturing methods for titanium and composite material structures are being developed and evaluated. The focus for the manufacturing effort is the fabrication of full-scale structural panels which replace an existing shear panel on the upper wing surface of the NASA YF-12 aircraft. The program involves design, fabrication, ground testing, and Mach 3 flight service of full-scale structural panels and laboratory testing of representative structural element specimens. *Lockheed, California Co., Burbank, California. 1977 Technical Publications 922. Parish, O. O.; and Putnam, T. W.: Equations for the Determination of Humidity From Dewpoint and Psychrometric Data. NASA TN D-8401, H-937, January 1977, 77N16859, #. A general expression based on the Claperon-Clausius differential equation that relates saturation vapor pressure, absolute temperature, and the latent heat of transformation was derived that expresses saturation vapor pressure as a function of absolute temperature. This expression was then used to derive general expressions for vapor pressure, absolute humidity, and relative humidity as functions of either dewpoint and ambient temperature or psychrometric parameters. Constants for all general expressions were then evaluated to give specific expressions in both the international system of units and U.S. customary units for temperatures above and below freezing. 923. Wilner, D. O.: Results of a Remote Multiplexer/Digitizer Unit Accuracy and Environmental Study. NASA TM X-56043, January 1977, 77N15368, #. **AeroVironment, Incorporated, Pasadena, California. †Lockheed Missiles and Space Company, Incorporated, Huntsville, Alabama. ECN-4242 B-747 Airplane With Smoke Generators A remote multiplexer/digitizer unit (RMDU), a part of the airborne integrated flight test data system, was subjected to an accuracy study. The study was designed to show the effects of temperature, altitude, and vibration on the RMDU. The RMDU was subjected to tests at temperatures from –54 C (–65 F) to 71 C (160 F), and the resulting data are presented here, along with a complete analysis of the effects. The methods and means used for obtaining correctable data and correcting the data are also discussed. 924. *Hallock, J. N.; *Burnham, D. C.; **Tombach, I. H.; †Brashears, M. R.; †Zalay, A. D.; and Barber, M. R.: Ground-Based Measurements of the Wake Vortex Characteristics of a B-747 Aircraft in Various Configurations. AIAA Paper 77-9, presented at the American Institute of Aeronautics and Astronautics, Fifteenth Aerospace Sciences Meeting, Los Angeles, California, January 24–26, 1977, 77A19770, #. A Boeing 747 aircraft flew 54 passes at low level over ground-based sensors. Vortex velocities were measured by a laser-Doppler velocimeter, an array of monostatic acoustic sounders, and an array of propeller anemometers. Flow visualization of the wake was achieved using smoke and balloon tracers. Preliminary results were obtained on the initial downwash field, the time for merging of the multiple vortices, the velocity fields, vortex decay, and the effects of spoilers and differential flap settings on the dissipation and structure of vortices. *U.S. Department of Transportation, Transportation Systems Center, Cambridge, Massachusetts. 157 925. Jenkins, Jerald M.; Problems Associated With Attaching Strain Gages to Titanium Alloy Ti-6Al-4V. NASA TM X-56044, February 1977. Weldable strain gages have shown excellent high temperature characteristics for supersonic cruise aircraft application. The spotwelding attachment method, however, has resulted in serious reductions in the fatigue life of titanium alloy (Ti-6Al-4V) fatigue specimens. The reduction is so severe that the use of weldable strain gages on operational aircraft must be prohibited. The cause of the fatigue problem is thought to be a combination of the microstructure changes in the material caused by spotwelding and the presence of the flange of the stain gage. Brazing, plating, and plasma spraying were investigated as substitutes for spotwelding. The attachment of a flangeless gage by plasma spraying provided the most improvement in the fatigue life of the titanium. 926. Burcham, F. W., Jr.; Lasagna, P. L.; and Kurtenbach, F. J.: Static and Flyover Noise Measurements of an Inverted Profile Exhaust Jet. ASME Paper 77-GT-81, presented at the American Society of Mechanical Engineers, Gas Turbine Conference and Products Show, Philadelphia, Pennsylvania, March 27–31, 1977, 77A28592, #. Tests using a TF30 mixed flow afterburning turbofan engine in an F-111 airplane were conducted to study the noise characteristics of an inverted velocity profile jet. Fullauthority digital engine control allowed the inverted profile jet to be compared to a uniform jet of equal thrust statically and in flight. An exhaust velocity survey showed that the ratio of the outer to inner stream velocities was 1.37; therefore, only small noise reductions were expected. At static conditions, the inverted profile jet was approximately 3 decibels quieter than the uniform jet at peak noise angles. During a flyover it was approximately 1 decibel quieter. 927. Edwards, J. W.; *Ashley, H.; and *Breakwell, J. V.: Unsteady Aerodynamic Modeling for Arbitrary Motions. AIAA Paper 77-451, presented at the Eighteenth Structures, Structural Dynamics and Materials Conference, March 21– 23, 1977, and Dynamics Specialist Conference, San Diego, California, March 24–25, 1977, Technical Papers, Vol. B, 1977, (see A77-25778 10-01), 77A25808, #. (See also 932.) A study is presented on the unsteady aerodynamic loads due to arbitrary motions of a thin wing and their adaptation for the calculation of response and true stability of aeroelastic modes. In an Appendix, the use of Laplace transform techniques and the generalized Theodorsen function for twodimensional incompressible flow is reviewed. New applications of the same approach are shown also to yield airloads valid for quite general small motions. Numerical results are given for the two-dimensional supersonic case. Previously proposed approximate methods, starting from simple harmonic unsteady theory, are evaluated by comparison with exact results obtained by the present approach. The Laplace inversion integral is employed to separate the loads into “rational” and “nonrational” parts, of which only the former are involved in aeroelastic stability of the wing. Among other suggestions for further work, it is explained how existing aerodynamic computer programs may be adapted in a fairly straightforward fashion to deal with arbitrary transients. *Stanford University, Stanford, California. 928. Kordes, E. E.: Influence of Structural Dynamics on Vehicle Design—Government View. AIAA Paper 77-438, presented at the Structures, Structural Dynamics and Materials Conference, March 21–23, 1977, and Dynamics Specialist Conference, San Diego, California, March 24–25, 1977, Technical Papers, Vol. B, (A77-25778 10-01), 1977, 77A25798, #. Dynamic design considerations for aerospace vehicles are discussed, taking into account fixed wing aircraft, rotary wing aircraft, and launch, space, and reentry vehicles. It is pointed out that space vehicles have probably had the most significant design problems from the standpoint of structural dynamics, because their large lightweight structures are highly nonlinear. Examples of problems in the case of conventional aircraft include the flutter encountered by high performance military aircraft with external stores. A description is presented of a number of examples which illustrate the direction of present efforts for improving aircraft efficiency. Attention is given to the results of studies 158 on the structural design concepts for the arrow-wing supersonic cruise aircraft configuration and a system study on low-wing-loading, short haul transports. 929. Szalai, K. J.; *Felleman, P. G.; **Gera, J.; and †Glover, R. D.: Design and Test Experience With a Triply Redundant Digital Fly-By-Wire Control System. Integrity in Electron. Flight Control Systems, AGARD AG-224, paper 21, April 1977, (see N77-25055 16-01), 77N25076, #. A triplex digital fly-by-wire flight control system was developed and then installed in a NASA F-8C aircraft to provide fail-operative, full authority control. Hardware and software redundancy management techniques were designed to detect and identify failures in the system. Control functions typical of those projected for future actively controlled vehicles were implemented. The principal design features of the system, the implementation of computer, sensor and actuator redundancy management, and the ground test results are described. An automated test program to verify sensor redundancy management software is also described. *Draper Laboratory, Inc., Cambridge, Massachusetts. **NASA Langley Research Center, Hampton, Virginia. †NASA Johnson Space Center, Houston, Texas. 930. Bartoli, F.: An Advanced Airborne Data Acquisition System. Flight Test Techniques, AGARD CP 223, paper 23, April 1977, (see N77-24107 15-05), 77N24130, #. The development and features of and user experience with an advanced airborne data acquisition system are described. The system consists of as many as 16 high speed pulse code modulation data acquisition units which are integrated with an airborne computer and a cockpit display unit. The data acquisition units may be operated without the computer. Operation without the computer is termed stand-alone operation. Computer integrated operation is intended for large-scale projects, and stand-alone operation is designed for small-scale projects. The cockpit display unit, which is part of the computer operated system, displays computed real time data in engineering units. An example of the cost reduction experienced by a major aircraft company by using the advanced data acquisition system is given. 931. *Newsom, B. D.; *Goldenrath, W. L.; *Sandler, H.; and Winter, W. R.: Tolerance of Females to +Gz Centrifugation Before and After Bedrest. Aviation, Space, and Environmental Medicine, Vol. 48, April 1977, pp. 327–331, 77A30881. Because women may be included as passengers in the proposed Space Shuttle System, experiments were conducted on 12 female subjects aged 24-35 yr. to investigate the +Gz tolerance of women and the possible degradation of this tolerance after a period of weightlessness as simulated by bedrest. Over a 1-week period, each subject was exposed to +Gz levels starting at +2 Gz and increasing by 0.5 Gz increments to a gray-out point. This point was determined by peripheral vision loss with a standard lightbar and by reverse blood flow in the temporal artery. Ultimately, each woman was subjected to three runs at the +3 Gz level for about 55 min long each, separated by 5-min rest periods. Eight subjects with the best tolerance times were selected for 14 days of bedrest in a horizontal position; the other four being ambulatory controls. Tests before bedrest, immediately following, and 5 days later showed that average +Gz tolerance decreased by 67% after bedrest. *NASA Ames Research Center, Moffett Field, California. 932. Edwards, J. W.: Unsteady Aerodynamic Modeling for Arbitrary Motions. AIAA Journal, Vol. 15, April 1977, pp. 593–595, 77A29910, #. (See also 927.) Results indicating that unsteady aerodynamic loads derived under the assumption of simple harmonic motions executed by airfoil or wing can be extended to arbitrary motions are summarized. The generalized Theodorsen (1953) function referable to loads due to simple harmonic oscillations of a wing section in incompressible flow, the Laplace inversion integral for unsteady aerodynamic loads, calculations of root loci of aeroelastic loads, and analysis of generalized compressible transient airloads are discussed. 933. Arnaiz, H. H.: Flight-Measured Lift and Drag Characteristics of a Large, Flexible, High Supersonic Cruise Airplane. NASA TM X-3532, H-913, May 1977, 77N24100, #. Flight measurements of lift, drag, and angle of attack were obtained for the XB-70 airplane, a large, flexible, high supersonic cruise airplane. This airplane had a length of over 57 meters, a takeoff gross mass of over 226,800 kilograms, and a design cruise speed of Mach 3 at an altitude of 21,340 meters. The performance measurements were made at Mach numbers from 0.72 to 3.07 and altitudes from approximately 7620 meters to 21,340 meters. The measurements were made to provide data for evaluating the techniques presently being used to design and predict the performance of aircraft in this category. Such performance characteristics as drag polars, lift-curve slopes, and maximum lift-to-drag ratios were derived from the flight data. The base drag of the airplane, changes in airplane drag with changes in engine power setting at transonic speeds, and the magnitude of the drag components of the propulsion system are also discussed. 934. Montoya, L. C.; and Banner, R. D.: F-8 Supercritical Wing Flight Pressure, Boundary Layer, and Wake Measurements and Comparisons With Wind 159 Tunnel Data. NASA TM X-3544, H-850, June 1977, 77N29098, #. Data for speeds from Mach 0.50 to Mach 0.99 are presented for configurations with and without fuselage area-rule additions, with and without leading-edge vortex generators, and with and without boundary-layer trips on the wing. The wing pressure coefficients are tabulated. Comparisons between the airplane and model data show that higher second velocity peaks occurred on the airplane wing than on the model wing. The differences were attributed to wind tunnel wall interference effects that caused too much rear camber to be designed into the wing. Optimum flow conditions on the outboard wing section occurred at Mach 0.98 at an angle of attack near 4 deg. The measured differences in section drag with and without boundary-layer trips on the wing suggested that a region of laminar flow existed on the outboard wing without trips. 935. Jacobs, P. F.; Flechner, S. G.; and Montoya, L. C.: Effect of Winglets on a First-Generation Jet Transport Wing. 1: Longitudinal Aerodynamic Characteristics of a Semispan Model at Subsonic Speeds. NASA TN D-8473, L-11354, June 1977, 78N20064, #. The effects of winglets and a simple wing-tip extension on the vectors behind the wing tip of a first generation jet transport wing were investigated in the Langley 8-foot transonic pressure tunnel using a semi-span model. The test was conducted at Mach numbers of 0.30, 0.70, 0.75, 0.78, and 0.80. At a Mach number of 0.30, the configurations were tested with combinations of leading- and trailing-edge flaps. 936. Montoya, L. C.; Flechner, S. G.; and Jacobs, P. F.: Effect of Winglets on a First-Generation Jet Transport Wing. 2: Pressure and Spanwise Load Distributions for a Semispan Model at High Subsonic Speeds. NASA TN D-8474, L-11026, July 1977, 78N20065, #. Pressure and spanwise load distributions on a first-generation jet transport semispan model at high subsonic speeds are presented for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wing-tip extension. Selected data are discussed to show the general trends and effects of the various configurations. 937. Montoya, L. C.; Jacobs, P. F.; and Flechner, S. G.: Effect of Winglets on a First-Generation Jet Transport Wing. 3: Pressure and Spanwise Load Distributions for a Semispan Model at Mach 0.30. NASA TN D-8478, L-11370, June 1977, 78N20063, #. Pressure and spanwise load distributions on a first-generation jet transport semispan model at a Mach number of 0.30 are given for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wingtip extension. To simulate second-segment-climb lift conditions, leading- and/or trailing-edge flaps were added to some configurations. 938. Jenkins, J. M.; and Kuhl, A. E.: A Study of the Effect of Radical Load Distributions on Calibrated Strain Gage Load Equations. NASA TM-56047, H-984, July 1977, 77N27430, #. For several decades, calibrated strain gages have been used to measure loads on airplanes. The accuracy of the equations used to relate the strain gage measurements to the applied loads has been based primarily on the results of the load calibration. An approach is presented for studying the effect of widely varying load distributions on strain gage load equations. The computational procedure provides a link between the load calibration and the load to be measured in flight. A matrix approach to equation selection is presented, which is based on equation standard error, load distribution, and influence coefficient plots of the strain gage equations, and is applied to a complex, delta-wing structure. 939. Jenkins, J. M.; Kuhl, A. E.; and Carter, A. L.: The Use of a Simplified Structural Model as an Aid in the Strain Gage Calibration of a Complex Wing. NASA TM-56046, H-959, July 1977, 77N27429, #. The use of a relatively simple structural model to characterize the load responses of strain gages located on various spars of a delta wing is examined. Strains measured during a laboratory load calibration of a wing structure are compared with calculations obtained from a simplified structural analysis model. Calculated and measured influence coefficient plots that show the shear, bending, and torsion characteristics of typical strain gage bridges are presented. Typical influence coefficient plots are shown for several load equations to illustrate the derivation of the equations from the component strain gage bridges. A relatively simple structural model was found to be effective in predicting the general nature of strain distributions and influence coefficient plots. The analytical processes are shown to be an aid in obtaining a good load calibration. The analytical processes cannot, however, be used in lieu of an actual load calibration of an aircraft wing. 940. *Brilliant, H. M.; and Bauer, C. A.: Comparison of Estimated With Measured Maximum Instantaneous Distortion Using Flight Data From an Axisymmetric Mixed Compression Inlet. AIAA Paper 77-876, presented at the American Institute of Aeronautics and Astronautics and Society of Automotive Engineers, Thirteenth Propulsion Conference, Orlando, Florida, July 11–13. 1977, 77A38570, #. 160 YF-12C flight-measured inlet dynamic distortion data are compared with predictions made on the basis of the method reported by Melick et al. (1976). The YF-12C aircraft is a twin engine aircraft capable of speeds above 3. The inlets have a translating spike to control the inlet throat area. A bypass system is used to control the terminal shock of the inlet for operation in the mixed compression mode. The dynamic data were obtained with the aid of 24 high frequency response total pressure sensors. The model of Melick et al. is discussed along with the computer program used to implement the model. It is found that the predictions of maximum instantaneous distortion are within 20 percent of the measured values, which had been obtained at Mach numbers of 1.8, 2.1, 2.5, and 3.0. *U.S. Air Force Academy, Colorado Springs, Colorado. ECN-3516 YF-12C Airplane 941. *Rawlings, K., III; *Cooper, J. M.; and Hughes, D. L.: Dynamic Test Techniques—Concepts and Practices. The Many Disciplines of Flight Test, Proceedings of the Seventh Annual Symposium, Eastsound, Orcas Island, Washington, August 4–6, 1976, Society of Flight Test Engineers, pp. 25-1 to 25-20, (see A77-38003 17-05), 77A38026. An initial investigation of dynamic flight test analysis techniques indicated that a strict, comprehensive forcemoment accounting system would be necessary. An implementation of the longitudinal force-moment accounting system provided excellent results in accounting for small lift/drag and tail deflection changes. Attention is given to gross thrust calculation, instrumentation, maneuvers, and aspects of data correlation. The results of the studies demonstrate that it is possible to generate a lift/drag model which is capable of predicting performance from nearly any maneuver. *USAF, Flight Test Center, Edwards AFB, California. 942. *Johnson, H. J.; and Painter, W. D.: The Development and Flight Test of an Electronic Integrated Propulsion Control System. The Many Disciplines of Flight Test, Proceedings of the Seventh Annual Symposium, Eastsound, Orcas Island, Washington, August 4–6, 1976, Society of Flight Test Engineers, pp. 12-1 to 12-19, (see A77-38003 17-05), 77A38013. Advanced technical features of the electronic integrated propulsion control system (IPCS) and flight evaluation tests of IPCS (F-111E with TF30-P-9 engines as test vehicle) are described. Nine baseline flight tests and 15 IPCS flight tests were conducted. Instrumentation, data acquisition and data processing systems, software maintenance procedures, flight test procedures, flight safety criteria, flight test results, and ground and flight testing of the aircraft system are described. Advantages conferred by IPCS include faster accelerations (both gas generator and afterburner performance), better thrust and flight control, reduced flight idle thrust, reduced engine ground trim, extended service ceiling, automatic stall detection, and stall recovery detection. *Boeing Commercial Airplane Co., Seattle, Washington. 943. Maine, R. E.: Maximum Likelihood Estimation of Aerodynamic Derivatives for an Oblique Wing Aircraft From Flight Data. AIAA Paper 77-1135, presented at the Atmospheric Flight Mechanics Conference, Hollywood, Florida, August 8–10, 1977, pp. 124–133, (see A77-43151 20-08), 77A43166, #. There are several practical problems in using current techniques on 5-degree-of-freedom equations to estimate the stability and control derivatives of oblique wing aircraft from flight data. A technique has been developed to estimate these derivatives by separating the analysis of the longitudinal and lateral-directional motion without neglecting cross-coupling effects. This technique was used on flight data from a remotely piloted oblique wing aircraft. The results demonstrated that the relatively simple approach developed was adequate to obtain high quality estimates of the aerodynamic derivatives of such aircraft. Oblique Wing Research Vehicle ECN-5209 944. Iliff, K. W.; and Maine, R. E.: Further Observations on Maximum Likelihood Estimates of Stability and Control Characteristics Obtained From Flight Data. AIAA Paper 77-1133, presented at the Atmospheric Flight Mechanics Conference, Hollywood, Florida, August 8–10, 1977, Technical Papers, pp. 100–112, (see A77-43151 20-08), 77A43164, #. A maximum likelihood estimation method for flight test data is described. Flight results based on 3000 maneuvers from 30 aircraft on the effect of resolution and sampling rate on the estimates, on understanding the discrepancies previously observed in the magnitude of the Cramer-Rao bounds, on the scale effects on the derivative estimates obtained from dynamic aircraft flight maneuvers, and on the analysis of lateral-directional maneuvers obtained in turbulence, are presented. 945. Edwards, J. W.; *Breakwell, J. V.; and *Bryson, A. E., Jr.: Active Flutter Control Using Generalized Unsteady Aerodynamic Theory. AIAA Guidance and Control Conference, Hollywood, Florida, August 8–10, 1977, Technical Papers, pp. 172–185, 77A42772, #. (See also 970.) This paper describes the application of generalized unsteady aerodynamic theory to the problem of active flutter control. The controllability of flutter modes is investigated. It is shown that the response of aeroelastic systems is composed of a portion due to a rational transform and a portion due to a 161 nonrational transform. The oscillatory response characteristic of flutter is due to the rational portion, and a theorem is given concerning the construction of a linear, finite-dimensional model of this portion of the system. The resulting rational model is unique and does not require state augmentation. Active flutter control designs using optimal regulator synthesis are presented. *Stanford University, Stanford, California. 946. *Hartmann, G.; *Stein, G.; and Petersen, K.: Flight Data Processing With the F-8 Adaptive Algorithm. AIAA Paper 77-1042, AIAA Guidance and Control Conference, Hollywood, Florida, Technical Papers, (see A77-42751 20-35), August 8–10, 1977, pp. 53–60, 77A42758, #. An explicit adaptive control algorithm based on maximum likelihood estimation of parameters has been designed for NASA’s DFBW F-8 aircraft. To avoid iterative calculations, the algorithm uses parallel channels of Kalman filters operating at fixed locations in parameter space. This algorithm has been implemented in NASA/DFRC’s Remotely Augmented Vehicle (RAV) facility. Real-time sensor outputs (rate gyro, accelerometer and surface position) are telemetered to a ground computer which sends new gain values to an on-board system. Ground test data and flight records were used to establish design values of noise statistics and to verify the ground-based adaptive software. The software and its performance evaluation based on flight data are described. *Honeywell, Inc., Minneapolis, Minnesota. 947. Andrews, W. H.; and McMurtry, T. C.: Space Shuttle Orbiter Approach and Landing Program Status. Flight Test Technology, Proceedings of the Eighth Annual Symposium, Washington, D.C., Society of Flight Test Engineers, (see A78-19426 06-01), August 10–12, 1977, pp. 2-1 to 2-14, 78A19428. The approach and landing test (ALT) phase of the Space Shuttle program aims at assessing the Orbiter’s subsonic aerodynamic flight and landing characteristics along with the support equipment, ground facilities, and the approximate hardware and software to be used in the terminal phase of orbital missions. The program also evaluates the performance of the Shuttle carrier aircraft (SCA) as related to the transport of the Orbiter to the launch sites during the stages of Space Shuttle operations. Results are presented for the SCA inert Orbiter flight testing and the program plans up to the completion of the ALT program. Emphasis is placed on testing the airworthiness of the mated B-747 Shuttle carrier aircraft and Orbiter, checkout of the Orbiter systems in captive flight, and launching the Orbiter with and without the tail cone installed. Major program milestones before the first manned orbital flight are summarized in graphic form. ECN77-8608 Space Shuttle Enterprise Launch From B-747 Airplane 948. Andrews, W. H.: Space Shuttle Orbiter Approach and Landing Program Status. AIAA Paper 77-1204, presented at the American Institute of Aeronautics and Astronautics, Aircraft Systems and Technology Meeting, Seattle, Washington, August 22–24, 1977, 77A44314, #. The space shuttle approach and landing test (ALT) program is being conducted in four phases. The first phase, completed in March of 1977, consisted of verifying the airworthiness of the mated B-747 shuttle carrier aircraft and orbiter. The second phase consists of checking the orbiter systems in captive flight. The third phase is to be confined to launching the orbiter with and without the tail cone installed to evaluate the final landing phase of the shuttle operations through the verification of the automatic landing system. The fourth and final phase of the program is to document the mated configuration’s performance relative to the ferry operations to be conducted between the return from the orbiter’s landing sites and launch sites. This paper presents the results of the SCA inert orbiter flight testing and the program plans up to the completion of the ALT program. 949. DeAngelis, V. M.; and Monaghan, R. C.: Buffet Characteristics of the F-8 Supercritical Wing Airplane. NASA TM-56049, H-945, September 1977, 77N32080, #. 162 The buffet characteristics of the F-8 supercritical wing airplane were investigated. Wing structural response was used to determine the buffet characteristics of the wing and these characteristics are compared with wind tunnel model data and the wing flow characteristics at transonic speeds. The wingtip accelerometer was used to determine the buffet onset boundary and to measure the buffet intensity characteristics of the airplane. The effects of moderate trailing edge flap deflections on the buffet onset boundary are presented. The supercritical wing flow characteristics were determined from wind tunnel and flight static pressure measurements and from a dynamic pressure sensor mounted on the flight test airplane in the vicinity of the shock wave that formed on the upper surface of the wing at transonic speeds. The comparison of the airplane’s structural response data to the supercritical flow characteristics includes the effects of a leading edge vortex generator. 950. Megna, Vincent A.; and Szalai, Kenneth J.: MultiFlight Computer Redundancy Management for Digital Fly-By-Wire Aircraft Control. IEEE COMPCON 77, September 1977. 951. Sefic, Walter J.; and Carter, Alan L.: Loads Calibration Experience With a Reentry Wing Structure. Presented at Fall Meeting, Western Regional Strain Gage Committee, Society for Experimental Stress Analysis, September 28, 1977. 952. Jenkins, Jerald M.; and Kuhl, Albert E.: Recent Loads Calibration Experience With a Delta Wing Airplane. Presented at Fall Meeting, Western Regional Strain Gage Committee, Society for Experimental Stress Analysis, September 28, 1977. 953. Tang, Ming H.; and Fields, Roger A.: Analysis of a Loads Calibration of a Hypersonic Cruise Wing Test Structure. Presented at Fall Meeting, Western Regional Strain Gage Committee, Society for Experimental Stress Analysis, September 28, 1977. 954. Steers, L. L.; and Saltzman, E. J.: Reduced Truck Fuel Consumption Through Aerodynamic Design. Journal of Energy, Vol. 1, No. 5, September–October 1977, pp. 312–318, 77A48572, #. Full-scale fuel consumption and drag tests were performed on a conventional cab-over-engine tractor-trailer combination and a version of the same vehicle with significant forebody modifications. The modified configuration had greatly increased radii on all front corners and edges of the tractor and a smooth fairing of the modified tractor top and sides extending to the trailer. Concurrent highway testing of the two configurations showed that the modified design used 20% to 24% less fuel than the baseline configuration at 88.5 km/hr (55 mph) with near-calm wind conditions. Coastdown test results showed that the modified configuration reduced the drag coefficient by 0.43 from the baseline value of 1.17 at 88.5 km/hr (55 mph) in calm wind conditions. ECN-4724 Low-Drag Truck 955. Fulton, F. L., Jr.: Shuttle Carrier Aircraft Flight Tests. Society of Experimental Test Pilots, Twenty-First Annual Symposium, Beverly Hills, California, Society of Experimental Test Pilots, Technical Review, Vol. 13, No. 4, October 12–15, 1977, pp. 191–204, 78A28464. Since the Space Shuttle will need to be transported from its place of assembly to the launch site, a method has been developed whereby the Shuttle rides piggyback on a modified Boeing 747, called the Shuttle carrier aircraft (SCA). This paper describes tests of the SCA in its mated configuration. Tests include flutter, found to decrease when fiberglass and wood fairings were added to the base of each supporting pylon; stability and control, found to be acceptable after damping with control pulses; noise and buffet, found high but acceptable; and climb, in which drag was marked but acceptable with the special rated thrust (SRT) power setting. Simulated launch maneuvers were undertaken at an airspeed of 273 KCAS. Transport of the Shuttle takes place with the 163 Shuttle tail cone on, at a cruise speed of 288 KCAS at an altitude of 22,000 feet. Flight tests were conducted with the YF-12 airplane to examine the airplane’s longitudinal characteristics at a Mach number of approximately 2.9. Phugoid oscillations as well as short period pulses were analyzed with the variable geometry engine inlets in the fixed and the automatic configurations. Stability and control derivatives for the velocity and altitude degrees of freedom and the standard short period derivatives were obtained. Inlet bypass door position was successfully used to represent the total inlet system, and the effect of the inlets on the velocity and altitude derivatives was determined. The phugoid mode of the basic airplane (fixed inlet configuration) had neutral damping, and the height mode was stable. With the addition of the inlets in the automatic configuration, the phugoid mode was slightly divergent and the height mode was divergent with a time to double amplitude of about 114 seconds. The results of the derivative estimation indicated that the change in the height mode characteristics was primarily the result of the change in the longitudinal force derivative with respect to velocity. 958. Jenkins, J. M.; Kuhl, A. E.; and Carter, A. L.: Strain Gage Calibration of a Complex Wing. Journal of Aircraft, Vol. 14, December 1977, pp. 1192–1196, 78A16182, #. Modern complex structural arrangements have complicated the task of measuring flight loads with calibrated strain gages. This paper examines the use of a relatively simple structural model to characterize the load responses of strain gages located on various spars of a delta wing. Strains measured during a laboratory load calibration of a wing structure are compared with calculations obtained from a simplified NASA structural analysis (NASTRAN) model. Calculated and measured influence coefficient plots that show the shear, bending, and torsion characteristics of typical strain-gage bridges are presented. Typical influence coefficient plots are given for several load equations to illustrate the derivation of the equations from the component strain-gage bridges. A relatively simple structural model was found to be effective in predicting the general nature of strain distributions and influence coefficient plots. The analytical processes are shown to be useful in obtaining a good load calibration. The analytical processes cannot, however, be used in lieu of an actual load calibration of an aircraft wing. 959. Iliff, Kenneth W.: Maximum Likelihood Estimation of Lift and Drag From Dynamic Aircraft Maneuvers. Journal of Aircraft, Vol. 14, No. 12, December 1977, pp. 1175–1181. A maximum likelihood estimation method for obtaining lift and drag characteristics from dynamic flight maneuvers was investigated. This paper describes the method and compares the estimates of lift and drag obtained by using the method with estimates obtained from wind-tunnel tests and from established methods for obtaining estimates from flight data. In general, the lift and drag coefficients extracted from 164 980058 Space Shuttle Mated to B-747, Three-View Drawing 956. Brown, S. R.; and Szalai, K. J.: Flight Experience With a Fail-Operational Digital Fly-By-Wire Control System. AIAA Paper 77-1507, presented at the Second Digital Avionics Systems Conference, Los Angeles, California, November 2–4, 1977, Collection of Technical Papers, 1977, pp. 186–199, (see A78-12226 02-04), 78A12253, #. The NASA Dryden Flight Research Center is flight testing a triply redundant digital fly-by-wire (DFBW) control system installed in an F-8 aircraft. The full-time, full-authority system performs three-axis flight control computations, including stability and command augmentation, autopilot functions, failure detection and isolation, and self-test functions. Advanced control law experiments include an active flap mode for ride smoothing and maneuver drag reduction. This paper discusses research being conducted on computer synchronization, fault detection, fault isolation, and recovery from transient faults. The F-8 DFBW system has demonstrated immunity from nuisance fault declarations while quickly identifying truly faulty components. 957. Powers, B. G.: Phugoid Characteristics of a YF-12 Airplane With Variable-Geometry Inlets Obtained in Flight Tests at a Mach Number of 2.9. NASA TP-1107, H-953, December 1977, 78N12100, #. dynamic flight maneuvers by the maximum likelihood estimation technique are in good agreement with the estimates obtained from the wind-tunnel tests and the other methods. When maneuvers that met the requirements of both flight methods were analyzed, the results of each method were nearly the same. The maximum likelihood estimation technique showed promise in terms of estimating lift and drag characteristics from dynamic flight maneuvers. Further studies should be made to assess the best mathematical model and the most desirable type of dynamic maneuver to get the highest quality results from this technique. 960. Smougur, T.; Morgan, T.; Sears, W.; Dana, W.; Enevoldson, E.; Melvin, J.; and *Tays, M.: Joint Testing of the RAF High Altitude Protective Ensemble. SAFE Association, Fifteenth Annual Symposium, Las Vegas, Nevada, December 5–8, 1977, Proceedings, pp. 243–245, (see A79-14401 03-03), 79A14434. The “get-me-down” capability from flight above 50,000 ft for the unencumbering RAF partial pressure clothing for use in F-104 and F-15 aircraft is tested. The equipment assembly tested includes a sleeveless Jerkin pressure vest, a G-suit and an RAF P/Q oronasal mask. The test program consists of six coordinated efforts: laboratory evaluation, orientation/ training of NASA test pilots, quantification of aerodynamic suction effects on cockpit altitude, definition of protective envelope, suit/aircraft integration, and in-flight test and evaluation. It is suggested that the RAF ensemble or equivalent would be the only currently available item that would be acceptable to tactical crews. The Jerkin ensemble appears to meet both the pilot’s physiological and functional requirements. *USAF, Aerospace Medical Div., Holloman AFB, New Mexico. 961. *Jacobsen, R. A.; and Barber, M. R.: Flight Test Techniques for Wake-Vortex Minimization Studies. NASA Wake Vortex Minimization, 1977, pp. 193–220, (see N78-12017 03-02), 78N12022, #. Flight test techniques developed for use in a study of wake turbulence and used recently in flight studies of wake minimization methods are discussed. Flow visualization was developed as a technique for qualitatively assessing minimization methods and is required in flight test procedures for making quantitative measurements. The quantitative techniques are the measurement of the upset dynamics of an aircraft encountering the wake and the measurement of the wake velocity profiles. Descriptions of the instrumentation and the data reduction and correlation methods are given. *NASA Ames Research Center, Moffett Field, California. 962. Barber, M. R.; Hastings, E. C., Jr.; Champine, R. A.; and *Tymczyszyn, J. J.: Vortex Attenuation Flight Experiments. NASA Wake Vortex Minimization, 1977, pp. 369–403, (see N78-12017 03-02), 78N12028, #. Flight tests evaluating the effects of altered span loading, turbulence ingestion, combinations of mass and turbulence ingestion, and combinations of altered span loading turbulence ingestion on trailed wake vortex attenuation were conducted. Span loadings were altered in flight by varying the deflections of the inboard and outboard flaps on a B-747 aircraft. Turbulence ingestion was achieved in flight by mounting splines on a C-54G aircraft. Mass and turbulence ingestion was achieved in flight by varying the thrust on the B-747 aircraft. Combinations of altered span loading and turbulence ingestion were achieved in flight by installing a spoiler on a CV-990 aircraft and by deflecting the existing spoilers on a B-747 aircraft. The characteristics of the attenuated and unattenuated vortexes were determined by probing them with smaller aircraft. Acceptable separation distances for encounters with the attenuated and unattenuated vortexes are presented. *FAA, Los Angeles, California. 963. Albers, J. A.: Inlet Operating Flow Field of the YF-12 Aircraft and Effects of This Flow Field on Inlet Performance. NASA Lewis Research Center Inlet Workshop, 1977, pp. 383–396, (see N86-72197 18-01), 86N72222. 964. Albers, J. A.; and Washington, H. P.: Technique for Determining Inlet Forces and Inlet Airframe Interactions on the F-15 Aircraft. NASA Lewis Research Center Inlet Workshop, 1977, pp. 615–631, (see N86-72197 18-01), 86N72235. ECN-5000 F-15 Airplane 165 965. Edwards, John William: Unsteady Aerodynamic Modeling and Active Aeroelastic Control. AZ 41 E95, NASA Grant NG L-05-020-007, Stanford University, Stanford, California, 1977. motion. Although the number of unknown independent functions is increased to include the state variables, the evaluation of the gradient of the cost function is simplified, resulting in considerable computational savings, thereby making it appear feasible to use the epsilon method for realtime application. 969. Deets, D. A.: Optimal Regulator or Conventional Setup Techniques for a Model Following Simulator Control System. NASA CP-007. NASA Washington Fourth Inter-Center Control Systems Conf.erence, January 1978, pp. 237–252, (see N78-23010 13-99), 78N23020, #. (See also 590.) Optimal regulator technique was compared for determining simulator control system gains with the conventional servo analysis approach. Practical considerations, associated with airborne motion simulation using a model-following system, provided the basis for comparison. The simulation fidelity specifications selected were important in evaluating the relative advantages of the two methods. Frequency responses for a JetStar aircraft following a roll mode model were calculated digitally to illustrate the various cases. A technique for generating forward loop lead in the optimal regulator model-following problem was developed which increases the flexibility of that approach. It appeared to be the only way in which the optimal regulator method could meet the fidelity specifications. 1978 Technical Publications 966. Painter, W. D.; and Caw, L. J.: Design and Physical Characteristics of the Transonic Aircraft Technology (TACT) Research Aircraft. NASA TM-56048, H-976, January 1978, 79N14014, #. The Transonic Aircraft Technology (TACT) research program provided data necessary to verify aerodynamic concepts, such as the supercritical wing, and to gain the confidence required for the application of such technology to advanced high performance aircraft. An F-111A aircraft was employed as the flight test bed to provide full scale data. The data were correlated extensively with predictions based on data obtained from wind tunnel tests. An assessment of the improvement afforded at transonic speeds in drag divergence, maneuvering performance, and airplane handling qualities by the use of the supercritical wing was included in the program. Transonic flight and wind tunnel testing techniques were investigated, and specific research technologies evaluated were also summarized. 967. Ko, W. L.: Finite Element Microscopic Stress Analysis of Cracked Composite Systems. Journal of Composite Materials, Vol. 12, January 1978, pp. 97–115, 78A28851. This paper considers the stress concentration problems of two types of cracked composite systems: (1) a composite system with a broken fiber (a penny-shaped crack problem), and (2) a composite system with a cracked matrix (an annular crack problem). The cracked composite systems are modeled with triangular and trapezoidal ring finite elements. Using NASTRAN (NASA Structural Analysis) finite element computer program, the stress and deformation fields in the cracked composite systems are calculated. The effect of fibermatrix material combination on the stress concentrations and on the crack opening displacements is studied. 968. Taylor, L. W., Jr.; and Smith, H. J.: A New Formulation for the Epsilon Method Applied to the Minimum-Time-to-Climb Problem. NASA CP-007. NASA Washington Fourth Inter-Center Control Systems Conference, January 1978, pp. 423–434, (see N78-23010 13-99), 78N23028, #. Balakrishnan’s epsilon technique is used to compute minimum-time profiles for the F-104 airplane. This technique differs from the classical gradient method in that a quadratic penalty on the error in satisfying the equation of motion is included in the cost function to be minimized as a means of eliminating the requirement of satisfying the equations of 166 980024 JetStar Airplane, Three-View Drawing 970. Edwards, J. W.; Breakwell, J. V.; and Bryson, A. E., Jr.: Active Flutter Control Using Generalized Unsteady Aerodynamic Theory. Journal of Guidance and Control, January–February 1978. Vol 1, No. 1, pp. 32–40. (See also 945.) 971. Powers, B. G.: Analytical Study of Ride Smoothing Benefits of Control System Configurations Optimized for Pilot Handling Qualities. NASA TP-1148, H-922, February 1978, 78N18076, #. An analytical study was conducted to evaluate the relative improvements in aircraft ride qualities that resulted from utilizing several control law configurations that were optimized for pilot handling qualities only. The airplane configuration used was an executive jet transport in the approach configuration. The control law configurations included the basic system, a rate feedback system, three command augmentation systems (rate command, attitude command, and rate command/attitude hold), and a control wheel steering system. Both the longitudinal and lateral directional axes were evaluated. A representative example of each control law configuration was optimized for pilot handling qualities on a fixed base simulator. The root mean square airplane responses to turbulence were calculated, and predictions of ride quality ratings were computed by using three models available in the literature. 972. Meyer, R. R., Jr.: Effect of Winglets on a FirstGeneration Jet Transport Wing. 4: Stability Characteristics for a Full-Span Model at Mach 0.30. NASA TP-1119, L-11705, February 1978, 78N17997, #. The static longitudinal and lateral directional characteristics of a 0.035 scale model of a first generation jet transport were obtained with and without upper winglets. The data were obtained for take off and landing configurations at a free stream Mach number of 0.30. The results generally indicated that upper winglets had favorable effects on the stability characteristics of the aircraft. 973. Sheridan, A. E.; and Grier, S. J.: Drag Reduction Obtained by Modifying a Standard Truck. NASA TM-72846, H-977, February 1978, 78N20457, #. A standard two-axle truck with a box-shaped cargo compartment was tested to determine whether significant reductions in aerodynamic drag could be obtained by modifying the front of the cargo compartment. The coastdown method was used to determine the total drag of the baseline vehicle, which had a square-cornered cargo box, and of several modified configurations. Test velocities ranged from 56.3 to 94.6 kilometers per hour (35 to 60 miles per hour). At 88.5 kilometers per hour (55 miles per hour), the aerodynamic drag reductions obtained with the modified configurations ranged from 8 to 30 percent. 974. Matheny, N. W.; and Gatlin, D. H.: Flight Evaluation of the Transonic Stability and Control Characteristics of an Airplane Incorporating a Supercritical Wing. NASA TP-1167, H-916, February 1978, 78N20140, #. A TF-8A airplane was equipped with a transport type supercritical wing and fuselage fairings to evaluate predicted performance improvements for cruise at transonic speeds. A comparison of aerodynamic derivatives extracted from flight and wind tunnel data showed that static longitudinal stability, effective dihedral, and aileron effectiveness, were higher than predicted. The static directional stability derivative was slower than predicted. The airplane’s handling qualities were acceptable with the stability augmentation system on. The unaugmented airplane exhibited some adverse lateral directional characteristics that involved low Dutch roll damping and low roll control power at high angles of attack and roll control power that was greater than satisfactory for transport aircraft at cruise conditions. Longitudinally, the aircraft exhibited a mild pitchup tendency. Leading edge vortex generators delayed the onset of flow separation, moving the pitchup point to a higher lift coefficient and reducing its severity. 975. Gee, S. W.; Carr, P. C.; Winter, W. R.; and Manke, J. A.: Development of Systems and Techniques for Landing an Aircraft Using Onboard Television. NASA TP-1171, H-973, February 1978, 78N20114, #. A flight program was conducted to develop a landing technique with which a pilot could consistently and safely land a remotely piloted research vehicle (RPRV) without outside visual reference except through television. Otherwise, instrumentation was standard. Such factors as the selection of video parameters, the pilot’s understanding of the television presentation, the pilot’s ground cockpit environment, and the operational procedures for landing were considered. About 30 landings were necessary for a pilot to become sufficiently familiar and competent with the test aircraft to make powered approaches and landings with outside visual references only through television. When steep approaches and landings were made by remote control, the pilot’s workload was extremely high. The test aircraft was used as a simulator for the F-15 RPRV, and as such was considered to be essential to the success of landing the F-15 RPRV. 976. Iliff, K. W.; Maine, R. E.; and Steers, S. T.: FlightDetermined Stability and Control Coefficients of the F-111A Airplane. NASA TM-72851, March 1978, 78N18075, #. A complete set of linear stability and control derivatives of the F-111A airplane was determined with a modified 167 maximum likelihood estimator. The derivatives were determined at wing sweep angles of 26 deg, 35 deg, and 58 deg. The flight conditions included a Mach number range of 0.63 to 1.43 and an angle of attack range of 2 deg to 15 deg. Maneuvers were performed at normal accelerations from 0.9g to 3.8g during steady turns to assess the aeroelastic effects on the stability and control characteristics. The derivatives generally showed consistent trends and reasonable agreement with the wind tunnel estimates. Significant Mach effects were observed for Mach numbers as low as 0.82. No large effects attributable to aeroelasticity were noted. 977. Ko, William L.: Traverse Diffusivity of Dual Phase Composites, Fiber Science and Technology, Vol. II, No. 2, March 1978, pp. 157–162. This paper compares the exact and approximate theories developed for predicting the traverse diffusivity of dual phase composite systems containing a rectangular lattice of uniform parallel circular cylindrical inclusions. Due to the difficulties in obtaining an explicit exact mathematical expression for the diffusivity of the composite system, approximations or mathematical models are usually introduced in the calculation of the diffusivity of the composite. The two most common approximate theories used in estimating the transverse diffusivity (or conductivity) of the composite systems are described. 978. Iliff, Kenneth W.: Identification and Stochastic Control of an Aircraft Flying in Turbulence. Journal of Guidance and Control, Vol. 1, No. 2, March–April 1978. Also published (in Russian) Rocket Technology and Cosmonautics, June 1979, pp. 150–159. An adaptive control technique to improve the flying qualities of an aircraft in turbulence was investigated. The approach taken was to obtain maximum likelihood estimates of the unknown coefficients of the aircraft system and then, using these estimates along with the separation principle, to define the stochastic optimal control. The maximum likelihood estimation technique that accounts for the effects of turbulence provided good estimates of the unknown coefficients and of the turbulence. The assessment of the stochastic optimal control based on the maximum likelihood estimates showed that the desired effects were attained for the regulator problem of minimizing pitch angle and the tracking problem of requiring normal acceleration to follow the pilot input. 979. Hedgley, D. R.: An Efficient Algorithm for Choosing the Degree of a Polynomial to Approximate Discrete Nonoscillatory Data. NASA TM-72854, H-1010, April 1978, 78N21839, #. An efficient algorithm for selecting the degree of a polynomial that defines a curve that best approximates a data set was presented. This algorithm was applied to both oscillatory and nonoscillatory data without loss of generality. 980. Gilyard, G. B.; and Smith, J. W.: Results From Flight and Simulator Studies of a Mach 3 Cruise Longitudinal Autopilot. NASA TP-1180, H-940, April 1978, 78N21160, #. At Mach numbers of approximately 3.0 and altitudes greater than 21,300 meters, the original altitude and Mach hold modes of the YF-12 autopilot produced aircraft excursions that were erratic or divergent, or both. Flight data analysis and simulator studies showed that the sensitivity of the static pressure port to angle of attack had a detrimental effect on the performance of the altitude and Mach hold modes. Good altitude hold performance was obtained when a high passed pitch rate feedback was added to compensate for angle of attack sensitivity and the altitude error and integral altitude gains were reduced. Good Mach hold performance was obtained when the angle of attack sensitivity was removed; however, the ride qualities remained poor. 980055 YF-12A Airplane, Three-View Drawing 981. *Siemers, P. M., III; and Larson, T. J.: The Space Shuttle Orbiter and Aerodynamic Testing. AIAA Paper 78-790, Tenth Annual Aerodynamic Testing Conference, San Diego, California, April 19–21, 1978, Technical Papers, pp. 145–158, (A78-32326 12-09),78A32347, #. The concept of utilizing the Space Shuttle Orbiter as an aerodynamic flight research vehicle is discussed. The orbiter’s planned flight frequency and its complex flight control system provide an unprecedented flight research potential. This paper defines the orbiter’s flight environment and applicable baseline systems, their capabilities and limitations, as well as those instrument systems required to augment the baseline capability. These required systems, 168 which are being developed under NASA’s Orbiter Experiments Program (OEX) are the Aerodynamic Coefficient Identification Package (ACIP), Shuttle Entry Air Data System (SEADS), and the Shuttle Upper Atmosphere Mass Spectrometer (SUMS). Finally, the need for and capability of launching payloads from the orbiter to extend the research potential beyond the orbiter configuration and/or environment is defined. *NASA Langley Research Center, Space Systems Div., Hampton, Virginia. 982. Powers, S. G.: Flight-Measured Pressure Characteristics of Aft-Facing Steps in High Reynolds Number Flow at Mach Numbers of 2.20, 2.50, and 2.80 and Comparison With Other Data. NASA TM-72855, H-956, May 1978, 78N25055, #. The YF-12 airplane was studied to determine the pressure characteristics associated with an aft-facing step in high Reynolds number flow for nominal Mach numbers of 2.20, 2.50, and 2.80. Base pressure coefficients were obtained for three step heights. The surface static pressures ahead of and behind the step were measured for the no-step condition and for each of the step heights. A boundary layer rake was used to determine the local boundary layer conditions. The Reynolds number based on the length of flow ahead of the step was approximately 10 to the 8th power and the ratios of momentum thickness to step height ranged from 0.2 to 1.0. Base pressure coefficients were compared with other available data at similar Mach numbers and at ratios of momentum thickness to step height near 1.0. In addition, the data were compared with base pressure coefficients calculated by a semiempirical prediction method. The base pressure ratios are shown to be a function of Reynolds number based on momentum thickness. Profiles of the surface pressures ahead of and behind the step and the local boundary layer conditions are also presented. 983. Monaghan, R. C.: Flight-Measured Buffet Characteristics of a Supercritical Wing and a Conventional Wing on a Variable-Sweep Airplane. NASA TP-1244, H-991, May 1978, 78N23056, #. Windup-turn maneuvers were performed to assess the buffet characteristics of the F-111A aircraft and the same aircraft with a supercritical wing, which is referred to as the F-111 transonic aircraft technology (TACT) aircraft. Data were gathered at wing sweep angles of 26, 35, and 58 deg for Mach numbers from 0.60 to 0.95. Wingtip accelerometer data were the primary source of buffet information. The analysis was supported by wing strain-gage and pressure data taken in flight, and by oil-flow photographs taken during tests of a wind tunnel model. In the transonic speed range, the overall buffet characteristics of the aircraft having a supercritical wing are significantly improved over those of the aircraft having a conventional wing. 169 F-111 TACT Airplane ECN-3945 984. Jenkins, Jerald M.; Fields, Roger A.; and Sefic, Walter J.: Effect of Elevated Temperature on the Calibrated Strain Gages of the YF-12A Wing. Presented at the Society of Experimental Stress Analysis, Spring Meeting, Wichita, Kansas, May 14–19, 1978. 985. Larson, Terry J.; and Schweikhard, William G.: Use of the Shuttle Entry Air Data Pressure System at Subsonic Speeds. Presented at the Second Biennial Air Data Systems Conference, Colorado Springs, Colorado, May 1–5, 1978. The purpose of this paper is to show that, in limited windtunnel tests of a 0.1-scale shuttle model, SEADS, when combined with auxiliary flush static-pressure measurements aft of the nose cap, can provide accurate air data system. The SEADS pressure data are from tests made in NASA Lewis Research Center’s 10- by 10-Foot Supersonic Wind Tunnel; the auxiliary flush static-pressure data are partly from those tests and partly from tests in the NASA Ames Research Center’s Fourteen-Foot Transonic Wind Tunnel. 986. Reed, R. D.: High-Flying Mini-Sniffer RPV— Mars Bound. Astronautics and Aeronautics, Vol. 16, June 1978, pp. 26–39, 78A38521, #. The Mini-Sniffer is a small unmanned survey aircraft developed by NASA to conduct turbulence and atmospheric pollution measurements from ground level to an altitude of 90,000 ft. Carrying a 25-lb air sampling apparatus, the MiniSniffer typically cruises for one hour at 70,000 ft before being remotely piloted back to earth. A hydrazine monopropellant engine powers the craft, while a PCM telemetering system and a radar transponder provide control functions. Development of a high-performance low-Reynolds-number airfoil could make the research craft suitable for a lowaltitude terrain-following mission on Mars. 988. Nugent, J.; Taillon, N. V.; and *Pendergraft, O. C., Jr.: Status of a Nozzle-Airframe Study of a Highly Maneuverable Fighter. AIAA Paper 78-990, presented at American Institute of Aeronautics and Astronautics and Society of Automotive Engineers, Fourteenth Joint Propulsion Conference, Las Vegas, Nevada, July 25–27, 1978, 78A48470, #. NASA is sponsoring a research program that uses coordinated wind tunnel and flight tests to investigate nozzleairframe flow interactions. The program objective is to compare transonic flight and wind tunnel measurements over a wide Reynolds number range. The paper discusses the progress of the program and the coordination of the wind tunnel and flight tests with regard to program elements, model-airplane differences, instrument locations, and test conditions. The real-time feedback techniques used to obtain steady flight conditions are presented. Available wind tunnel results are presented for the jet effects model showing the influence of the rear-end geometry and test variables on nozzle drag. Available flight results show the effect of the variable inlet ramp angle and angle of attack on fuselage pressures and upper surface boundary layers. Mini-Sniffer RPV ECN-4898 *NASA Langley Research Center, Hampton, Virginia. 989. *Lucas, E. J.; **Fanning, A. E.; and Steers, L. L.: Comparison of Nozzle and Afterbody Surface Pressures From Wind Tunnel and Flight Test of the YF-17 Aircraft. AIAA Paper 78-992, presented at American Institute of Aeronautics and Astronautics and Society of Automotive Engineers, Fourteenth Joint Propulsion Conference, Las Vegas, Nevada, July 25–27, 1978, 78A43540, #. Results are reported from the initial phase of an effort to provide an adequate technical capability to accurately predict the full scale, flight vehicle, nozzle-afterbody performance of future aircraft based on partial scale, wind tunnel testing. The primary emphasis of this initial effort is to assess the current capability and identify the cause of limitations on this capability. A direct comparison of surface pressure data is made between the results from an 0.1-scale model wind tunnel investigation and a full-scale flight test program to evaluate the current subscale testing techniques. These data were acquired at Mach numbers 0.6, 0.8, 0.9, 1.2, and 1.5 on four nozzle configurations at various vehicle pitch attitudes. Support system interference increments were also documented during the wind tunnel investigation. In general, the results presented indicate a good agreement in trend and level of the surface pressures when corrective increments are applied for known effects and surface differences between the two articles under investigation. 987. *Peterson, J. B., Jr.; and Fisher, D. F.: Flight Investigation of Insect Contamination and Its Alleviation. NASA CP-2036-PT-1. NASA CTOL Transport Technol., June 1978, pp. 357–373, (see N78-27046 18-01), 78N27067, #. An investigation of leading edge contamination by insects was conducted with a JetStar airplane instrumented to detect transition on the outboard leading edge flap and equipped with a system to spray the leading edge in flight. The results of airline type flights with the JetStar indicated that insects can contaminate the leading edge during takeoff and climbout. The results also showed that the insects collected on the leading edges at 180 knots did not erode at cruise conditions for a laminar flow control airplane and caused premature transition of the laminar boundary layer. None of the superslick and hydrophobic surfaces tested showed any significant advantages in alleviating the insect contamination problem. While there may be other solutions to the insect contamination problem, the results of these tests with a spray system showed that a continuous water spray while encountering the insects is effective in preventing insect contamination of the leading edges. *NASA Langley Research Center, Hampton, Virginia. 170 *ARO, Inc., Arnold Air Force Station, Tennessee. **USAF, Aero Propulsion Laboratory, Wright-Patterson AFB, Ohio. *Martin Marietta Aerospace, Bethesda, Maryland. **General Dynamics Corp., Pomona, California. 992. Fisher, D. F.; and *Peterson, J. B., Jr.: Flight Experience on the Need and Use of Inflight Leading Edge Washing for a Laminar Flow Airfoil. AIAA Paper 78-1512, presented at American Institute of Aeronautics and Astronautics, Aircraft Systems and Technology Conference, Los Angeles, California, August 21–23, 1978, 78A47947, #. An investigation of leading-edge contamination by insects was conducted at the NASA Dryden Flight Research Center with a JetStar airplane instrumented to detect transition on the outboard leading-edge flap and equipped with a system to wash the leading edge in flight. The results of airline-type flights with the JetStar indicated that insects can contaminate the leading edge during take-off and climbout at large jet airports in the United States. The results also showed that the insects collected on the leading edges at 180 knots did not erode at cruise conditions for a laminar flow control airplane and caused premature transition of the laminar boundary layer. None of the superslick and hydrophobic surfaces tested showed any significant advantages in alleviating the insect contamination problem. While there may be other solutions to the insect contamination problem, the results of these tests with a washer system showed that a continuous water spray while encountering the insects is effective in preventing insect contamination of the leading edges. *NASA Langley Research Center, Hampton, Virginia. 993. Montoya, L. C.; Bikle, P. F.; and Banner, R. D.: Section Drag Coefficients From Pressure Probe Traverses of a Wing Wake at Low Speeds. AIAA Paper 78-1479, presented at the American Institute of Aeronautics and Astronautics, Aircraft Systems and Technology Conference, Los Angeles, California, August 21–23, 1978, 78A47924, #. (See also 1035.) This paper reviews the techniques used to increase data reliability and to minimize certain bias errors during a series of wing profile drag measurements performed in flight on a sailplane airfoil. Unresolved questions concerning errors in the use of total probes in this and other studies are discussed. 994. Sisk, T. R.: A Technique for the Assessment of Fighter Aircraft Precision Controllability. AIAA Paper 78-1364, Atmospheric Flight Mechanics Conference, Palo Alto, California, August 7–9, 1978, Technical Papers, pp. 253–265, (see A78-46526 20-08), 78A46553, #. Today’s emerging fighter aircraft are maneuvering as well at normal accelerations of 7 to 8 g’s as their predecessors did at 4 to 5 g’s. This improved maneuvering capability has significantly expanded their operating envelope and made the EC76-5270 YF-17 Airplane 990. Brownlow, J.: A Statistical Package for Computing Time and Frequency Domain Analysis. NASA TM-56045, H-981, August 1978, 78N29843, #. The spectrum analysis (SPA) program is a general purpose digital computer program designed to aid in data analysis. The program does time and frequency domain statistical analyses as well as some preanalysis data preparation. The capabilities of the SPA program include linear trend removal and/or digital filtering of data, plotting and/or listing of both filtered and unfiltered data, time domain statistical characterization of data, and frequency domain statistical characterization of data. 991. *Morosow, G.; **Dublin, M.; and Kordes, E. E.: Needs and Trends in Structural Dynamics. Astronautics and Aeronautics, Vol. 16, July–August 1978, pp. 90–94, 78A43364, #. The paper discusses dynamic analyses and testing of aerospace vehicles and the application of such analyses and testing to nonaerospace fields. Items covered in the section on dynamic analyses of aerospace vehicles include self-induced and forced oscillatory loads, approaches to dynamic modeling and analysis, nonlinear analyses, and integrated dynamics design and optimization. Items covered in the section on the dynamic testing of aerospace vehicles include integrated test philosophy, test facilities, and ways of improving performance and reducing costs. The nonaerospace applications that are discussed include ground and water transportation, medicine, and nuclear power plants. 171 task of evaluating handling qualities more difficult. This paper describes a technique for assessing the precision controllability of highly maneuverable aircraft, a technique that was developed to evaluate the effects of buffet intensity on gunsight tracking capability and found to be a useful tool for the general assessment of fighter aircraft handling qualities. It has also demonstrated its usefulness for evaluating configuration and advanced flight control system refinements. This technique is believed to have application to future aircraft dynamics and pilot-vehicle interface studies. 995. Maine, R. E.; and Iliff, K. W.: Maximum Likelihood Estimation of Translational Acceleration Derivatives From Flight Data. AIAA Paper 78-1342, Atmospheric Flight Mechanics Conference, Palo Alto, California, August 7–9, 1978, Technical Papers, pp. 121–131, (see A78-46526 20-08), 78A46539, #. (See also 1069.) This paper shows that translational acceleration derivatives, such as pitching moment due to rate of change of angle of attack, can be estimated from flight data with the use of appropriately designed maneuvers. No new development of estimation methodology is necessary to analyze these maneuvers. Flight data from a T-37B airplane were used to verify that rate of change of angle of attack could be estimated from rolling maneuvers. 996. Anon.: YF-12 Experiments Symposium. NASA CP-2054-VOL-1, H-1059. Proceedings of the YF-12 Experiments Symposium, Edwards, California, September 13–15, 1978, 78N32055, #. (See also N78-32056 through N78-32065.) Papers presented by personnel from the Dryden Flight Research Center, the Lewis Research Center, and the Ames Research Center are presented. Topics cover propulsion system performance, inlet time varying distortion, structures, aircraft controls, propulsion controls, and aerodynamics. The reports were based on analytical studies, laboratory experiments, wind tunnel tests, and extensive flight research with two YF-12 airplanes. 997. Kock, B. M.: Overview of the NASA YF-12 Program. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 3–25, (see N78-32055 23-02), 78N32056, #. The history of NASA’s interest in supersonic research and the agency’s contribution to the development of the YF-12 aircraft is reviewed as well as the program designed to use that aircraft as a test bed for supersonic cruise research. Topics cover elements of the program, project organization, and major accomplishments. 998. Jenkins, J. M.; and Kuhl, A. E.: Recent Load Calibrations Experience With the YF-12 Airplane. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 47–72, (see N78-32055 23-02), 78N32057, #. The use of calibrated strain gages to measure wing loads on the YF-12A airplane is discussed as well as structural configurations relative to the thermal environment and resulting thermal stresses. A thermal calibration of the YF-12A is described to illustrate how contaminating thermal effects can be removed from loads equations. The relationship between ground load calibrations and flight measurements is examined for possible errors, and an analytical approach to accommodate such errors is presented. 999. Meyer, R. R., Jr.; and DeAngelis, V. M.: FlightMeasured Aerodynamic Loads on a 0.92 Aspect Ratio Lifting Surface. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 73–91, (see N78-32055 23-02), 78N32058, #. Ventral fin loads, expressed as normal force coefficients, bending moment coefficients, and torque coefficients, were measured during flight tests of a YF-12A airplane. Because of the proximity of the ventral fin to the ailerons, the aerodynamic loads presented were the result of both sideslip loads and aileron crossflow loads. Aerodynamic data obtained from strain gage loads instrumentation and some flight pressure measurements are presented for several Mach numbers ranging from 0.70 to 2.00. Selected wind tunnel data and results of linear theoretical aerodynamic calculations are presented for comparison. 1000. Gilyard, G. B.; and Smith, J. W.: Flight Experience With Altitude Hold and Mach Hold Autopilots on the YF-12 Aircraft at Mach 3. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 97–119, (see N78-30255 23-02), 78N32059, #. The altitude hold mode of the YF-12A airplane was modified to include a high-pass-filtered pitch rate feedback along with optimized inner loop altitude rate proportional and integral gains. An autothrottle control system was also developed to control either Mach number or KEAS at the high-speed flight conditions. Flight tests indicate that, with the modified system, significant improvements are obtained in both altitude and speed control, and the combination of altitude and autothrottle hold modes provides the most stable aircraft platform thus far demonstrated at Mach 3 conditions. 1001. Rezek, T. W.: Pilot Workload Measurement and Experience on Supersonic Cruise Aircraft. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 121–134, (see N78-32055 23-02), 78N32060, #. 172 Aircraft parameters and physiological parameters most indicative of crew workload were investigated. Recommendations were used to form the basis for a continuing study in which variations of the interval between heart beats are used as a measure of nonphysical workload. Preliminary results are presented and current efforts in further defining this physiological measure are outlined. 1002. Ehernberger, L. J.: The YF-12 Gust Velocity Measuring System. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 135–154, (see N78-32055 23-02), 78N32061, #. A true gust velocity measuring system designed to alleviate complications resulting from airframe flexibility and from the high-speed, high-temperature environment of supersonic cruise aircraft was evaluated on a YF-12 airplane. The system uses fixed vanes on which airflow direction changes produce differential pressure variations that are measured. Airframe motions, obtained by postflight integration of recorded angular rate and linear acceleration data, are removed from the flow angle data. An example of turbulence data obtained at high-altitude, supersonic flight conditions is presented and compared with previous high-altitude turbulence measurements obtained with subsonic aircraft and with turbulence criteria contained in both military and civil design specifications for supersonic cruise vehicles. Results of these comparisons indicate that the YF-12 turbulence sample is representative of turbulence present in the supersonic cruise environment. 1003. Powers, S. G.: Flight-Measured Pressure Characteristics of Aft-Facing Steps in Thick Boundary Layer Flow for Transonic and Supersonic Mach Numbers. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 201–226, (see N78-32055 23-02), 78N32063, #. Aft-facing step base pressure flight data were obtained for three step heights for nominal transonic Mach numbers of 0.80, 0.90, and 0.95, and for supersonic Mach numbers of 2.2, 2.5, and 2.8 with a Reynolds number, based on the fuselage length ahead of the step, of about 10 to the 8th power. Surface static pressures were measured ahead of the step, behind the step, and on the step face (base), and a boundary layer rake was used to obtain boundary layer reference conditions. A comparison of the data from the present and previous experiments shows the same trend of increasing base pressure ratio (decreasing drag) with increasing values of momentum thickness to step height ratios. However, the absolute level of these data does not always agree at the supersonic Mach numbers. For momentum thickness to height ratios near 1.0, the differences in the base pressure ratios appear to be primarily a function of Reynolds number based on the momentum thickness. Thus, for Mach numbers above 2, the data analyzed show that the base pressure ratio decreases (drag increases) as Reynolds number based on momentum thickness increases for a given momentum thickness and step height. 1004. Fisher, D. F.: Boundary Layer, Skin Friction, and Boattail Pressure Measurements From the YF-12 Airplane at Mach Numbers Up to 3. NASA CP-2054VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 227–258, (see N78-32055 23-02), 78N32064, #. In-flight measurements of boundary layer and skin friction data were made on YF-12 airplanes for Mach numbers between 2.0 and 3.0. Boattail pressures were also obtained for Mach numbers between 0.7 and 3.0 with Reynolds numbers up to four hundred million. Boundary layer data measured along the lower fuselage centerline indicate local displacement and momentum thicknesses can be much larger than predicted. Skin friction coefficients measured at two of five lower fuselage stations were significantly less than predicted by flat plate theory. The presence of large differences between measured boattail pressure drag and values calculated by a potential flow solution indicates the presence of vortex effects on the upper boattail surface. At both subsonic and supersonic speeds, pressure drag on the longer of two boattail configurations was equal to or less than the pressure drag on the shorter configuration. At subsonic and transonic speeds, the difference in the drag coefficient was on the order of 0.0008 to 0.0010. In the supersonic cruise range, the difference in the drag coefficient was on the order of 0.002. Boattail drag coefficients are based on wing reference area. 1005. Quinn, R. D.; and Gong, L.: In-Flight Compressible Turbulent Boundary Layer Measurements on a Hollow Cylinder at a Mach Number of 3.0. NASA CP-2054-VOL-1. NASA YF-12 Experiments Symposium, Vol. 1, August 1978, pp. 259–286, (see N78-32055 23-02), 78N32065, #. Skin temperatures, shearing forces, surface static pressures, and boundary layer pitot pressures and total temperatures were measured on a hollow cylinder 3.04 meters long and 0.437 meter in diameter mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 and at wall-to-recovery temperature ratios of 0.66 to 0.91. The free stream Reynolds number had a minimal value of 4.2 million per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. Boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor of 1.11 was obtained from the measured heat transfer and skin friction data. The skin friction coefficients predicted by the theory of van Driest were in excellent agreement with the measurements. 173 Theoretical heat transfer coefficients, in the form of Stanton numbers calculated by using a modified Reynolds analogy between skin friction and heat transfer, were compared with measured values. The measured velocity profiles were compared to Coles’ incompressible law-of-the-wall profile. Transonic Aircraft Technology, Lancaster, California, August 15–17, 1978. This paper present the results of a study conducted with the F-111 transonic aircraft technology (TACT) airplane to assess the improvement in maneuverability afforded by a supercritical wing when installed on an F-111A aircraft. The study evaluated the aerodynamic performance, maneuver performance, and precision controllability of both the basic F-111A and the F-111 TACT aircraft over the transonic Mach number region at three wing-sweep positions. The aerodynamic performance evaluation showed that the supercritical wing significantly improved the buffet characteristics of the F-111A airplane at high subsonic and transonic Mach numbers. Wing rock was experienced on both aircraft, with the F-111 TACT airplane having the higher onset boundaries. The maneuver performance evaluation showed the F-111 TACT airplane to have higher drag-rise Mach number as well as improved transonic sustained turn performance. The supercritical wing did not alter the F-111A airplane’s basic precision controllability. The agility analysis demonstrated that the supercritical wing improved the F-111A airplane’s maneuverability. 1008. Maine, R. E.: Aerodynamic Derivatives for an Oblique Wing Aircraft Estimated From Flight Data by Using a Maximum Likelihood Technique. NASA TP-1336, H-1003, October 1978, 78N33054, #. There are several practical problems in using current techniques with five degree of freedom equations to estimate the stability and control derivatives of oblique wing aircraft from flight data. A technique was developed to estimate these derivatives by separating the analysis of the longitudinal and lateral directional motion without neglecting cross coupling effects. Although previously applied to symmetrical aircraft, the technique was not expected to be adequate for oblique wing vehicles. The application of the technique to flight data from a remotely piloted oblique wing aircraft is described. The aircraft instrumentation and data processing were reviewed, with particular emphasis on the digital filtering of the data. A complete set of flight determined stability and control derivative estimates is presented and compared with predictions. The results demonstrated that the relatively simple approach developed was adequate to obtain high quality estimates of the aerodynamic derivatives of such aircraft. 1009. Tang, M. H.; Sefic, W. J.; and Sheldon, R. G.: Comparison of Concurrent Strain Gage- and Pressure Transducer-Measured Flight Loads on a Lifting Reentry Vehicle and Correlation With Wind Tunnel Predictions. NASA TP-1331, H-1035, October 1978, 78N33053, #. Concurrent strain gage and pressure transducer measured flight loads on a lifting reentry vehicle are compared and 174 ECN-4777 YF-12 Airplane With “Coldwall” Experiment 1006. Monaghan, Richard C.: Flight Measured Buffet Characteristics. Presented at the Symposium on Transonic Aircraft Technology, Lancaster, California, August 15–17, 1978. Windup-turn maneuvers were performed to assess the buffet characteristics of the F-111A aircraft and the same aircraft with a supercritical wing, which is referred to as the F-1ll transonic aircraft technology (TACT) aircraft. Data were gathered at wing sweep angles of 26 degree, 35 degree, and 58 degree for Mach numbers from 0.60 to 0.95. Wingtip accelerometer data were the primary source of buffet information. The analysis was supported by wing strain-gage and fuselage accelerometer data. Buffet intensity rise boundaries in the form of plots of aircraft normal-force coefficient as a function of Mach number, as well as individual buffet intensity curves at specific Mach numbers, are presented for each aircraft. In the transonic speed range, the overall buffet characteristics of the aircraft having a supercritical wing are significantly improved over those of the aircraft having a conventional wing. At subsonic speeds or at the aft wing sweep position where the supercritical wing is off design, the two aircraft have similar buffet characteristics. 1007. Sakamoto, Glenn M.; and Friend, Edward L.: Agility Evaluation. Presented at the Symposium on correlated with wind tunnel-predicted loads. Subsonic, transonic, and supersonic aerodynamic loads are presented for the left fin and control surfaces of the X-24B lifting reentry vehicle. Typical left fin pressure distributions are shown. The effects of variations in angle of attack, angle of sideslip, and Mach number on the left fin loads and rudder hinge moments are presented in coefficient form. Also presented are the effects of variations in angle of attack and Mach number on the upper flap, lower flap, and aileron hinge-moment coefficients. The effects of variations in lower flap hinge moments due to changes in lower flap deflection and Mach number are presented in terms of coefficient slopes. 1010. Sim, A. G.; and Curry, R. E.: Flight-Determined Stability and Control Derivatives for the F-111 TACT Research Aircraft. NASA TP-1350, H-1004, October 1978, 79N10068, #. A flight investigation was conducted to provide a stability and control derivative data base for the F-111 transonic aircraft technology research aircraft. Longitudinal and lateral-directional data were obtained as functions of Mach number, angle of attack, and wing sweep. For selected derivatives, the flight results were correlated with derivatives calculated based on vehicle geometry. The validity of the angle of attack measurement was independently verified at a Mach number of 0.70 for angles of attack between 3 and 10 degrees. 1011. Curry, R. E.: Utilization of the Wing-Body Aerodynamic Analysis Program. NASA TM-72856, H-1071, October 1978, 79N10020, #. The analysis program was used to investigate several aircraft characteristics. The studies performed included vehicle stability analysis, determination of upwash angle, identification of nonpotential flow, launch dynamics, and wake vortex upset loads. The techniques and are discussed. When possible, computed results are compared with experimental data. 1012. Ko, W. L.: Effect of Crack Size on the Natural Frequencies of a Cracked Plate. International Journal of Fracture, Vol. 14, October 1978, pp. R273–R275, 79A20012. Results are presented for a finite element modal analysis of a centrally cracked rectangular plate made of linearly elastic material. The objective is to assess the effect of crack size on the natural vibration frequencies of the cracked plate. Only the in-plane vibration modes are studied. The results presented are finite-element mesh-size dependent. Namely, shrinking the mesh size, especially in the crack tip region, 175 would change the magnitudes of the natural frequencies, and the trends would be the same as shown in the present note. 1013. *Hamer, M. J.; and Kurtenbach, F. J.: A Simplified Gross Thrust Computing Technique for an Afterburning Turbofan Engine. Why Flight Test, Proceedings of the Ninth Annual SFTE Symposium, Arlington, Texas, October 4–6, 1978, pp. 22-1 to 22-20, (see A79-50426 22-01), 79A50440. A simplified gross thrust computing technique extended to the F100-PW-100 afterburning turbofan engine is described. The technique uses measured total and static pressures in the engine tailpipe and ambient static pressure to compute gross thrust. Empirically evaluated calibration factors account for three-dimensional effects, the effects of friction and mass transfer, and the effects of simplifying assumptions for solving the equations. Instrumentation requirements and the sensitivity of computed thrust to transducer errors are presented. NASA altitude facility tests on F100 engines (computed thrust versus measured thrust) are presented, and calibration factors obtained on one engine are shown to be applicable to the second engine by comparing the computed gross thrust. It is concluded that this thrust method is potentially suitable for flight test application and engine maintenance on production engines with a minimum amount of instrumentation. *Computing Devices Co., Ottawa, Canada. 1014. Iliff, K. W.; Maine, R. E.; and Montgomery, T. D.: Considerations in the Analysis of Flight Test Maneuvers. Why Flight Test, Proceedings of the Ninth Annual SFTE Symposium, Arlington, Texas, October 4–6, 1978, pp. 10-1 to 10-36, (see A79-50426 22-01), 79A50433. This paper discusses the application of a maximum likelihood estimator to dynamic flight-test data. The information presented is based on the experience in the past twelve years at the NASA Dryden Flight Research Center of estimating stability and control derivatives from over 3,000 maneuvers from 32 aircraft. The overall approach to the analysis of dynamic flight-test data is discussed. Detailed requirements for data and instrumentation are discussed and several examples of the types of problems that may be encountered are presented. 1015. Jenkins, J. M.; Schuster, L. S.; and Carter, A. L.: Correlation of Predicted and Measured Thermal Stresses on a Truss-Type Aircraft Structure. NASA TM-72857, H-1074, November 1978, 79N11995, #. A test structure representing a portion of a hypersonic vehicle was instrumented with strain gages and thermocouples. This test structure was then subjected to laboratory heating representative of supersonic and hypersonic flight conditions. A finite element computer model of this structure was developed using several types of elements with the NASA structural analysis (NASTRAN) computer program. Temperature inputs from the test were used to generate predicted model thermal stresses and these were correlated with the test measurements. 1016. Burcham, F. W., Jr.: Propulsion-Flight Control Integration Technology. AGARD-AG-234, Active Controls in Aircraft Design, November 1978, (see N79-16864 08-08), 79N16872, #. The propulsion-flight control integration technology (PROFIT) concept to be implemented on a high performance supersonic twin-engine aircraft which will make possible the evaluation of a wide variety of integrated control concepts is discussed. The aircraft’s inlet, engine, and flight control systems are to be integrated with a digital computer. The airplane control hardware is to be modified to provide the necessary capability for control research; software will be used to provide flexibility in the control integration capability. The background for flight and propulsion control system development and probable future trends are described. Examples of integrated control research that have application to future aircraft designs are also presented. 1017. Deets, D. A.; and *Crother, C. A.: Highly Maneuverable Aircraft Technology. AGARD AG-234, Active Controls in Aircraft Design, November 1978, (see N79-16864 08-08), 79N16871, #. A remotely piloted research vehicle (RPRV) with active controls designed to develop high maneuverable aircraft technologies (HiMAT) is described. The HiMAT RPRV is the central element in a new method to bring advanced aircraft technologies to a state of readiness. The RPRV is well into the construction phase, with flight test evaluations planned. The closely coupled canard-wing vehicle includes relaxed static stability, direct force control, and a digital active control system. Nonlinearities in the aerodynamics led to unusual demands on the active control systems. For example, the longitudinal static margin is 10-percent negative at low angles of attack, but increases to 30-percent negative at high angles of attack and low Mach numbers. The design procedure followed and experiences encountered as they relate to the active control features are discussed. Emphasis is placed on the aspects most likely to be encountered in the design of a full-scale operational vehicle. In addition, a brief overview of the flight control system features unique to the RPRV operation is presented. *Rockwell International Corp., Los Angeles, California. HiMAT RPRV ECN-14280 1018. Hartmann, G. L.; Stein, G.; Szalai, K. J.; Brown, S. R.; and Petersen, K. L.: F-8 Active Control. AGARD AG-234, paper 6, Active Controls in Aircraft Design, November 1978, (see N79-16864 08-08), 79N16870, #. An advanced flight control research program conducted with a modified F-8C aircraft is described. Key technologies investigated include system redundancy management and active control laws. Two control law packages proposed for flight test are discussed. The first is the control configured vehicle package which incorporates command augmentation, boundary control, ride smoothing, and maneuver flap functions. The second package is an adaptive control law based on a parallel channel maximum likelihood estimation algorithm. The design, implementation, and flight test experience with both sets of control laws are described. 1019. Iliff, K. W.: Estimation of Aerodynamic Characteristics From Dynamic Flight Test Data. AGARD CP-235, Dynamic Stability Parameters, November 1978, (see N79-15061 06-08), 79N15075, #. Significant effort was spent in estimating unknown aircraft coefficients, such as stability and control derivatives from dynamic flight maneuvers. The techniques used to estimate these coefficients are becoming increasingly complex; however, these techniques make it possible to obtain estimates of coefficients that in the past were nearly impossible to obtain. A survey of the investigations that were undertaken to obtain estimates of coefficients from dynamic flight maneuvers is presented. One method, the maximum likelihood estimation technique, is described briefly and some of the successful applications of the technique are presented. Possible techniques for analyzing responses obtained in the stall/spin regime are discussed. Recent data obtained in the stall/spin flight regime are presented along with a discussion of how some basic results can be obtained with simple analysis techniques. 176 1020. Szalai, K. J.; Jarvis, C. R.; Krier, G. E.; *Megna, V. A.; *Brock, L. D.; and *O’Donnell, R. N.: Digital Fly-ByWire Flight Control Validation Experience. NASA TM-72860, R-1164, H-1080, December 1978, 79N14109, #. The experience gained in digital fly-by-wire technology through a flight test program being conducted by the NASA Dryden Flight Research Center in an F-8C aircraft is described. The system requirements are outlined, along with the requirements for flight qualification. The system is described, including the hardware components, the aircraft installation, and the system operation. The flight qualification experience is emphasized. The qualification process included the theoretical validation of the basic design, laboratory testing of the hardware and software elements, systems level testing, and flight testing. The most productive testing was performed on an iron bird aircraft, which used the actual electronic and hydraulic hardware and a simulation of the F-8 characteristics to provide the flight environment. The iron bird was used for sensor and system redundancy management testing, failure modes and effects testing, and stress testing in many cases with the pilot in the loop. The flight test program confirmed the quality of the validation process by achieving 50 flights without a known undetected failure and with no false alarms. *Charles Draper Laboratory, Cambridge, Massachusetts, Inc. 1021. *Burnham, D. C.; *Hallock, J. N.; **Tombach, I. H.; †Brashears, M. R.; and Barber, M. R.: Ground-Based Measurements of the Wake Vortex Characteristics of a B-747 Aircraft in Various Configurations. NASA TM-80474, AD-A067588, December 1978, 79N26016, #. A Boeing 747 aircraft flew 54 passes at low altitude over ground based sensors. Vortex velocities were measured by a laser Doppler velocimeter, an array of monostatic acoustic sounders, and an array of propeller anemometers. Flow visualization of the wake was achieved using smoke and balloon tracers and was recorded photographically. Data were obtained on vortex velocity fields, vortex decay, and the effects of spoilers and differential flap settings on the dissipation and structure of the vortices. *Transportation Systems Center, Cambridge, Massachusetts. **AeroVironment, Inc., Pasadena, California. †Lockheed Missiles and Space Co., Huntsville, Alabama. 1022. Friend, E. L.; and Sakamoto, G. M.: Flight Comparison of the Transonic Agility of the F-111A Airplane and the F-111 Supercritical Wing Airplane. NASA TP-1368, H-985, December 1978, 79N13056, #. A flight research program was conducted to investigate the improvements in maneuverability of an F-111A airplane equipped with a supercritical wing. In this configuration the aircraft is known as the F-111 TACT (transonic aircraft technology) airplane. The variable-wing-sweep feature 177 permitted an evaluation of the supercritical wing in many configurations. The primary emphasis was placed on the transonic Mach number region, which is considered to be the principal air combat arena for fighter aircraft. An agility study was undertaken to assess the maneuverability of the F-111A aircraft with a supercritical wing at both design and off-design conditions. The evaluation included an assessment of aerodynamic and maneuver performance in conjunction with an evaluation of precision controllability during tailchase gunsight tracking tasks. 1023. Burcham, F. W., Jr.; Lasagna, P. L.; and *Oas, S. C.: Measurements and Predictions of Flyover and Static Noise of a TF30 Afterburning Turbofan Engine. NASA TP-1372, H-1017, December 1978, 79N13045, #. The noise of the TF30 afterburning turbofan engine in an F-111 airplane was determined from static (ground) and flyover tests. A survey was made to measure the exhaust temperature and velocity profiles for a range of power settings. Comparisons were made between predicted and measured jet mixing, internal, and shock noise. It was found that the noise produced at static conditions was dominated by jet mixing noise, and was adequately predicted by current methods. The noise produced during flyovers exhibited large contributions from internally generated noise in the forward arc. For flyovers with the engine at nonafterburning power, the internal noise, shock noise, and jet mixing noise were accurately predicted. During flyovers with afterburning power settings, however, additional internal noise believed to be due to the afterburning process was evident; its level was as much as 8 decibels above the nonafterburning internal noise. Power settings that produced exhausts with inverted velocity profiles appeared to be slightly less noisy than power settings of equal thrust that produced uniform exhaust velocity profiles both in flight and in static testing. *Boeing Co. Airplane Co., Seattle, Washington. 1024. Kurtenbach, F. J.: Comparison of Calculated and Altitude-Facility-Measured Thrust and Airflow of Two Prototype F100 Turbofan Engines. NASA TP-1373, H-1015, December 1978, 79N13044, #. A comparison is made of the facility performance data for the two engines with an engine performance model, and it provides corrections that can be applied to the model so that it represents the test engines accurately over the flight envelope. Test conditions ranged from Mach numbers of 0.80 to 2.00 and altitudes from 4020 meters to 15,240 meters. Two distortion screens were used to determine the effect of distortion on airflow. Reynolds number effects were also determined. Engine hysteresis is documented, as is an attempt to determine engine degradation. The calibrated engine model had a twice standard deviation accuracy of approximately 1.24 percent for corrected airflow and 2.38 percent for gross thrust. 1025. Brenner, M. J.; Iliff, K. W.; and Whitman, R. K.: Effect of Sampling Rate and Record Length on the Determination of Stability and Control Derivatives. NASA TM-72858, H-1077, December 1978, 79N12096, #. Flight data from five aircraft were used to assess the effects of sampling rate and record length reductions on estimates of stability and control derivatives produced by a maximum likelihood estimation method. Derivatives could be extracted from flight data with the maximum likelihood estimation method even if there were considerable reductions in sampling rate and/or record length. Small amplitude pulse maneuvers showed greater degradation of the derivative maneuvers than large amplitude pulse maneuvers when these reductions were made. Reducing the sampling rate was found to be more desirable than reducing the record length as a method of lessening the total computation time required without greatly degrading the quantity of the estimates. 1026. *Siegel, W. H.; Fields, R. A.; and **Easley, J. T.: Experimental Investigation of the Buckling Characteristics of a Beaded Skin Panel for a Hypersonic Aircraft—Including Comparisons With Finite Element and Classical Analyses. ASME Paper 78-WA/AERO-3, presented at the American Society of Mechanical Engineers, Winter Annual Meeting, San Francisco, California, December 10–15, 1978, 79A19717, #. Results of a compression test of a beaded panel intended for a proposed hypersonic aircraft are presented. The panel was tested to failure at room temperature to determine its buckling characteristics, in particular, to study the buckling caused by pure compression. The boundary conditions of the panel simulated as nearly as possible a wing mounted condition. Strain, out-of-plane deflection, and load data were measured, and elastic buckling strength as well as mode shapes of the panel were determined. Application of the moire technique is described. *University of California, Berkley, California and Lawrence Livermore Laboratory, Livermore, California. **University of Kansas, Lawrence, Kansas. 1027. Montoya, L. C.; *Flechner, S. G.; and *Jacobs, P. F.: Effect of an Alternate Winglet on the Pressure and Spanwise Load Distributions of a First Generation Jet Transport Wing. NASA TM-78786, L-12519, December 1978, 79N14012, #. Pressure and spanwise load distributions on a first-generation jet transport semispan model at subsonic speeds are presented. The wind tunnel data were measured for the wing with and without an alternate winglet. The results show that the winglet affected outboard wing pressure distributions and increased the spanwise loads near the tip. *NASA Langley Research Center, Hampton, Virginia. 178 1028. Szalai, Kenneth J.: The F-8 Digital Fly-By-Wire Program. A Status Report to the Winter Meeting of the ASME, December 13, 1978. 1029. Gee, S. W.; and Brown, Samuel R.: Flight Tests of a Radio-Controlled Airplane Mode With a Free-Wing, Free-Canard Configuration. NASA TM-72853, H-1008, 1978, 78N18042, #. Flight characteristics, controllability, and potential operating problems were investigated in a radio-controlled airplane model in which the wing is so attached to the fuselage that it is free to pivot about a spanwise axis forward of its aerodynamic center and is subject only to aerodynamic pitching moments imposed by lift and drag forces and a control surface. A simple technique of flying the test vehicle in formation with a pickup truck was used to obtain trim data. The test vehicle was flown through a series of maneuvers designed to permit evaluation of certain characteristics by observation. The free-wing free-canard concept was determined to be workable. Stall/spin characteristics were considered to be excellent, and no effect on longitudinal stability was observed when center of gravity changes were made. Several problems were encountered during the early stages of flight testing, such as aerodynamic lockup of the free canard and excessive control sensitivity. Lack of onboard instrumentation precluded any conclusions about gust alleviation or ride qualities. 1030. *Dunham, R. E., Jr.; Barber, M. R.; and Croom, D. R.: Wake Vortex Technology. NASA CP-2036-PT-2. NASA CTOL Transport Technology, 1978, pp. 757–771, (see N7829046 20-01), 78N29055, #. A brief overview of the highlights of NASA’s wake vortex minimization program is presented. The significant results of this program are summarized as follows: (1) it is technically feasible to reduce significantly the rolling upset created on a trailing aircraft; (2) the basic principles or methods by which reduction in the vortex strength can be achieved have been identified; and (3) an analytical capability for investigating aircraft vortex wakes has been developed. *NASA Langley Research Center, Hampton, Virginia. 1031. Ko, W. L.: An Orthotropic Sandwich Plate Containing a Part-Through Crack Under Mixed Mode Deformation. Engineering Fracture Mechanics, Vol. 10, No. 1, 1978, pp. 15–23, 78A23567. 1032. Fields, R. A.: Dryden Flight Research Center Hot Structures Research. Recent Advan. in Structures for Hypersonic Flight, NASA CP-2065, Part 2, 1978, pp. 707–750, (see N79-21435 12-39), 79N21441, #. The facilities, testing techniques, and design methods are described for NASA Dryden Flight Research Center. High temperature strain gage technology, realistic flight hardware fabrication, and structural analysis are discussed. A considerable amount of experimental work on hot structure concepts for hypersonic vehicles was performed; all the work is not complete, and there are still problem areas that need to be resolved. 1033. Shideler, J. L.; Fields, R. A.; and Reardon, L. F.: Tests of Beaded and Tubular Structural Panels. Recent Advan. in Structures for Hypersonic Flight, NASA CP-2065, Part 2, 1978, pp. 538–576, (see N79-21435 12-39), 79N21436, #. Two efficient concepts built from curved elements were identified, and a data base for tubular panels was developed. The tubular panel failure modes were understood, and the data base for these panels indicated that their performance can be predicted. The concepts are currently being tested in a realistic builtup structure; 157 room temperature tests and 67 hot tests were made with no structural failures, although all of these tests were not at the design load of the structure. 1034. Lux, D. P.: In-Flight Three-Dimensional Boundary Layer and Wake Measurements From a Swept Supercritical Wing. NASA CP-2045-PT-2. NASA Langley Res. Center Advanced Technology Airfoil Res., Vol. 1, Part 2, 1978, pp. 643–655, (see N79-19989 11-01), 79N20002, #. Three-dimensional boundary layer and wake velocity profiles were measured in flight on the supercritical wing of the F-111 transonic aircraft technology aircraft. These data, along with pressure distributions, were obtained to establish a data base with which data obtained by three-dimensional analytical techniques could be correlated. Only a brief summary of the total data base is given. The data presented represented one chord station at a wing leading-edge sweep angle of 26 deg. They cover an angle of attack range from 6 degs to 9 degs at free-stream Mach numbers from 0.85 to 0.90. A brief discussion of the techniques used to obtain the boundary layer and wake profiles is included. 1035. Montoya, L. C.; *Bikle, P. F.; and Banner, R. D.: Section Drag Coefficients From Pressure Probe Transverses of a Wing Wake at Low Speeds. NASA CP-2045-PT-2. NASA Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Part 2, 1978, pp. 601–621, (see N79-19989 11-01), 79N20000, #. (See also 993.) An in-flight wing wake section drag investigation was conducted using traversing pitot and static probes. The primary objective was to develop measurement techniques and improve the accuracy of in-flight wing profile drag measurements for low values of dynamic pressure and Reynolds number. Data were obtained on a sailplane for speeds from about 40 knots to 125 knots at chord Reynolds numbers between 1,000,000 and 3,000,000. Tests were conducted with zero flap deflection, deflected flaps, and various degrees of surface roughness, and for smooth and rough atmospheric conditions. Several techniques were used 179 to increase data reliability and to minimize certain bias errors. A discussion of the effects of a total pressure probe in a pressure gradient, and the effects of discrete turbulence levels, on the data presented and other experimental results is also included. *Soaring Society of America, Santa Monica, California. 1979 Technical Publications 1036. Rediess, H. A.: Avionics and Controls Research and Technology. NASA CP-2061, January 1979, 79N15898. The workshop provided a forum for industry and universities to discuss the state-of-the-art, identify the technology needs and opportunities, and describe the role of NASA in avionics and controls research. 1037. Webb, L. D.; Whitmore, S. A.; and *Janssen, R. L.: Preliminary Flight and Wind Tunnel Comparisons of the Inlet/Airframe Interaction of the F-15 Airplane. AIAA Paper 79-0102, presented at the American Institute of Aeronautics and Astronautics, Seventeenth Annual Aerospace Sciences Meeting, New Orleans, Louisiana, January 15–17, 1979, 79A23513, #. Preliminary flight and wind tunnel comparison data are presented for the F-15 inlet/airframe interactions program. Test conditions and instrumentation for both the model and the aircraft are described. Flight and wind tunnel inlet drag data (for a 0-deg angle of attack and Mach numbers of 0.6, 0.9, and 1.2), derived by using nearly identical pressure integration equations, are compared. The effects of a movable cowl, movable ramps, and other system components on pressure flow fields along the airframe are discussed. Excellent agreement between wind tunnel and flight pressure-integrated drags is found at all three Mach numbers. The wind tunnel data show good agreement for pressureintegrated and force-balance-measured inlet drag, except at Mach 0.6. Flight-measured pressure-integrated inlet lift is lower than that measured in the wind tunnel. 1038. *Stevens, C. H.; *Spong, E. D.; Nugent, J.; and **Neumann, H. E.: Reynolds Number, Scale, and Frequency Content Effects on F-15 Inlet Instantaneous Distortion. AIAA Paper 79-0104, presented at the American Institute of Aeronautics and Astronautics, Seventeenth Annual Aerospace Sciences Meeting, New Orleans, Louisiana, January 15–17, 1979, 79A19533, #. An inlet instantaneous distortion study program sponsored by NASA was recently completed using an F-15 fighter aircraft. Peak distortion data from subscale inlet model wind tunnel tests are shown to be representative of full-scale flight test peak distortion. The effects on peak distortion are investigated for engine presence, Reynolds number, scale and frequency content. Data are presented which show that: (1) the effect of engine presence on total pressure recovery, peak distortion, and turbulence is small but favorable, (2) increasing the Reynolds number increases total pressure recovery, decreases peak distortion, and decreases turbulence, and (3) increasing the filter cutoff frequency increases the calculated values of both peak distortion and turbulence. *McDonnell Douglas Corp., St. Louis, Missouri. **NASA Lewis Research Center, Cleveland, Ohio. 1039. Rawlings, K., III; Cooper, J. M.; and Hughes, D. L.: Dynamic Test Techniques—Concepts and Practices. Society of Flight Test Engineers, Journal, Vol. 1, January 1979, pp. 10–20, 79A50164. The concepts involved in dynamic performance testing represent a philosophy of developing a single, coherent performance model using thrust and drag modeling. It is shown that if all thrust drag interactions, maneuver rates, and instrumentation are taken into consideration, it is possible to generate a lift/drag model which is capable of predicting performance from nearly any maneuver. Although this capability is dependent on the ability to calculate gross thrust adequately, the engines can be calibrated through static thrust runs and the thrust computation procedures verified in flight. The capability of generating a consistent lift/drag model in considerably less time than conventional performance methods is demonstrated. 1040. Hedgley, D. R.: A Characterization of the Real Zeros of a Particular Transcendental Function. NASA TP-1420, H-1065, March 1979, 79N18679, #. The real zeros of the transcendental function y = ax + becx are characterized, and the results should alleviate the difficulty of determining their existence, location, and number. 1041. Jenkins, J. M.; Fields, R. A.; and Sefic, W. J.: Elevated-Temperature Effects on Strain Gages on the YF-12A Wing. Presented at the Society for Experimental Stress Analysis, Spring Meeting, Wichita, Kansas, May 14–18, 1978, Experimental Mechanics, Vol. 19, March 1979, pp. 81–86, 79A26400. A general study is made of the effects of structural heating on calibrated-strain-gage load measurements on the wing of a supersonic airplane. The primary emphasis is on temperatureinduced effects as they relate to slope changes and thermal shifts of the applied load/strain relationships. These effects are studied by using the YF-12A airplane, a structural computer model, and subsequent analyses. Such topics as the thermal environment of the structure, the variation of load paths at elevated temperature, the thermal response characteristics of load equations, elevated-temperature loadmeasurement approaches, the thermal calibration of wings, 180 and the correlation of strains are discussed. Ways are suggested to measure loads with calibrated strain gages in the supersonic environment. 1042. Iliff, K. W.; Maine, R. E.; and Montgomery, T. D.: Important Factors in the Maximum Likelihood Analysis of Flight Test Maneuvers. NASA TP-1459, H-1076, April 1979, 79N22113, #. The information presented is based on the experience in the past 12 years at the NASA Dryden Flight Research Center of estimating stability and control derivatives from over 3500 maneuvers from 32 aircraft. The overall approach to the analysis of dynamic flight test data is outlined. General requirements for data and instrumentation are discussed and several examples of the types of problems that may be encountered are presented. 1043. Jenkins, J. M.: Correlation of Predicted and Measured Thermal Stresses on an Advanced Aircraft Structure With Similar Materials. NASA TM-72862, H-1086, April 1979, 79N20989, #. A laboratory heating test simulating hypersonic heating was conducted on a heat-sink type structure to provide basic thermal stress measurements. Six NASTRAN models utilizing various combinations of bar, shear panel, membrane, and plate elements were used to develop calculated thermal stresses. Thermal stresses were also calculated using a beam model. For a given temperature distribution there was very little variation in NASTRAN calculated thermal stresses when element types were interchanged for a given grid system. Thermal stresses calculated for the beam model compared similarly to the values obtained for the NASTRAN models. Calculated thermal stresses compared generally well to laboratory measured thermal stresses. A discrepancy of significance occurred between the measured and predicted thermal stresses in the skin areas. A minor anomaly in the laboratory skin heating uniformity resulted in inadequate temperature input data for the structural models. 1044. Tanner, R. R.; and Montgomery, T. D.: Stability and Control Derivative Estimates Obtained From Flight Data for the Beech 99 Aircraft. NASA TM-72863, H-1081, April 1979, 79N20134, #. Lateral-directional and longitudinal stability and control derivatives were determined from flight data by using a maximum likelihood estimator for the Beech 99 airplane. Data were obtained with the aircraft in the cruise configuration and with one-third flap deflection. The estimated derivatives show good agreement with the predictions of the manufacturer. 1045. Burcham, F. W., Jr.: Measurements and Predictions of Flyover and Static Noise of an Afterburning Turbofan Engine in an F-111 Airplane. AIAA Paper 79-7018, Proceedings of the Fourth International Symposium on Air Breathing Engines, Orlando, Florida, April 1–6, 1979, pp. 133–145, (see A79-29376 11-07), 79A29391, #. The noise of the TF30 afterburning turbofan engine in an F-111 airplane was determined from static (ground) and flyover tests. Exhaust temperatures and velocity profiles were measured for a range of power settings. Comparisons were made between predicted and measured jet mixing, internal, and shock noise. It was found that the noise produced at static conditions was dominated by jet mixing noise, and was adequately predicted by current methods. The noise produced during flyovers exhibited large contributions from internally generated noise in the forward arc. For flyovers with the engine at nonafterburning power, the internal noise, shock noise, and jet mixing noise were accurately predicted. During flyovers with afterburning power settings, however, additional internal noise believed to be due to the afterburning process was evident; its level was as much as 8 decibels above the nonafterburning internal noise. 1046. Edwards, J. W.: Applications of Laplace Transform Methods to Airfoil Motion and Stability Calculations. AIAA Paper 79-0772, Technical Papers on Structures and Materials, Twentieth Structures, Structural Dynamics, and Materials Conference, St. Louis, Missouri, April 4–6, 1979, pp. 465–481, (see A79-29002 11-39), 79A29050, #. This paper reviews the development of generalized unsteady aerodynamic theory and presents a derivation of the generalized Possion integral equation. Numerical calculations resolve questions concerning subsonic indicial lift functions and demonstrate the generation of Kutta waves at high values of reduced frequency, subsonic Mach number, or both. The use of rational function approximations of unsteady aerodynamic loads in aeroelastic stability calculations is reviewed, and a reformulation of the matrix Pade approximation technique is given. Numerical examples of flutter boundary calculations for a wing which is to be flight tested are given. Finally, a simplified aerodynamic model of transonic flow is used to study the stability of an airfoil exposed to supersonic and subsonic flow regions. 1047. Ko, W. L.: Elastic Constants for Superplastically Formed/Diffusion-Bonded Sandwich Structures. AIAA Paper 79-0756, Technical Papers on Dynamics and Loads, Twentieth Structures, Structural Dynamics, and Materials Conference, St. Louis, Missouri, April 4–6, 1979, pp. 188–207, (see A79-28251 10-39), 79A28271, #. (See also 1102.) Formulae and the associated graphs are presented for contrasting the effective elastic constants for a superplastically formed/diffusion-bonded (SPF/DB) corrugated sandwich core and a honeycomb sandwich core. 181 The results used in the comparison of the structural properties of the two types of sandwich cores are under conditions of equal sandwich density. It was found that the stiffness in the thickness direction of the optimum SPF/DB corrugated core (i.e., triangular truss core) was lower than that of the honeycomb core, and that the former had higher transverse shear stiffness than the latter. 1048. *Davis, R. E.; *Champine, R. A.; and Ehernberger, L. J.: Meteorological and Operational Aspects of 46 Clear Air Turbulence Sampling Missions With an Instrument B-57B Aircraft. Volume 1: Program Summary. NASA TM-80044, May 1979, 79N25667, #. The results of 46 clear air turbulence (CAT) probing missions conducted with an extensively instrumented B-57B aircraft are summarized. Turbulence samples were obtained under diverse conditions including mountain waves, jet streams, upper level fronts and troughs, and low altitude mechanical and thermal turbulence. CAT was encouraged on 20 flights comprising 77 data runs. In all, approximately 4335 km were flown in light turbulence, 1415 km in moderate turbulence, and 255 km in severe turbulence during the program. The flight planning, operations, and turbulence forecasting aspects conducted with the B-57B aircraft are presented. *NASA Langley Research Center, Hampton, Virginia. B-57B Airplane ECN-21064 1049. Sisk, T. R.; and Matheny, N. W.: Precision Controllability of the F-15 Airplane. NASA TM-72861, H-1073, May 1979, 79N23979, #. A flying qualities evaluation conducted on a preproduction F-15 airplane permitted an assessment to be made of its precision controllability in the high subsonic and low transonic flight regime over the allowable angle of attack range. Precision controllability, or gunsight tracking, studies were conducted in windup turn maneuvers with the gunsight in the caged pipper mode and depressed 70 mils. This evaluation showed the F-15 airplane to experience severe buffet and mild-to-moderate wing rock at the higher angles of attack. It showed the F-15 airplane radial tracking precision to vary from approximately 6 to 20 mils over the load factor range tested. Tracking in the presence of wing rock essentially doubled the radial tracking error generated at the lower angles of attack. The stability augmentation system affected the tracking precision of the F-15 airplane more than it did that of previous aircraft studied. 1051. Saltzman, E. J.: Reductions in Vehicle Fuel Consumption Due to Refinements in Aerodynamic Design. Learning To Use Our Environment; Proceedings of the Twenty-Fifth Annual Technical Meeting, Seattle, Washington, April 30–May 2, 1979, pp. 63–68, Institute of Environmental Sciences, Prospect, Illinois (see A79-50326 22-42), 79A50335. Over-the-highway fuel consumption and coastdown drag tests were performed on cab-over-engine, van type trailer trucks and modifications of these vehicles incorporating refinements in aerodynamic design. In addition, 1/25-scale models of these configurations, and derivatives of these configurations were tested in a wind tunnel to determine the effects of wind on the magnitude of the benefits that aerodynamic refinements can provide. The results of these tests are presented for a vehicle incorporating major redesign features and for a relatively simple add-on modification. These results include projected fuel savings on the basis of annual savings per vehicle year as well as probable nationwide fuel savings. 1052. Borek, Robert W.: Practical Considerations in the Selection of Electrical Connectors for Aircraft Instrumentation Systems. AGARD Short Course on Flight Test Instrumentation, Cranfield, England, May 7–18, 1979. 1053. Iliff, Kenneth W.; and Maine, Richard E.: Observation on Maximum Likelihood Estimation of Aerodynamic Characteristics From Flight Data. Journal of Guidance and Control, Vol. 2, No. 3, May–June 1979. (Also in Rocket Technology and Cosmonautics, September 1979, pp. 152–160, [in Russian].) 980059 F-15 Aircraft, Three-View Drawing. 1050. Stoll, F.; Tremback, J. W.; and Arnaiz, H. H.: Effect of Number of Probes and Their Orientation on the Calculation of Several Compressor Face Distortion Descriptors. NASA TM-72859, H-1070, May 1979, 79N23087, #. A study was performed to determine the effects of the number and position of total pressure probes on the calculation of five compressor face distortion descriptors. This study used three sets of 320 steady state total pressure measurements that were obtained with a special rotating rake apparatus in wind tunnel tests of a mixed-compression inlet. The inlet was a one third scale model of the inlet on a YF-12 airplane, and it was tested in the wind tunnel at representative flight conditions at Mach numbers above 2.0. The study shows that large errors resulted in the calculation of the distortion descriptors even with a number of probes that were considered adequate in the past. There were errors as large as 30 and -50 percent in several distortion descriptors for a configuration consisting of eight rakes with five equal-area-weighted probes on each rake. This paper discusses the application of a maximum likelihood estimation method in flight test data. The results are based on 11 years’ experience of estimating stability and control derivatives from 3000 maneuvers from 30 aircraft. Flight results are presented from recent studies on understanding the discrepancies previously observed in the magnitude of the Cramer-Rao bounds, on the scale effects on the derivative estimates obtained from dynamic aircraft flight maneuvers, and on the analysis of lateral-directional maneuvers obtained in turbulence. 1054. *Ashworth, G. R.; Putnam, T. W.; Dana, W. H.; Enevoldson, E. K.; and Winter, W. R.: Flight Test Evaluation of an RAF High Altitude Partial Pressure Protective Assembly. NASA TM-72864, June 1979, 79N24654, #. A partial pressure suit was evaluated during tests in an F-104 and F-15 as a protective garment for emergency descents. The garment is an pressure jerkin and modified anti-g suit combined with an oronasal mask. The garment can be donned and doffed at the aircraft to minimize thermal buildup. The oronasal mask was favored by the pilots due to its immobility 182 on the face during high g-loading. The garment was chosen to provide optimum dexterity for the pilot, which is not available in a full pressure suit, while protecting the pilot at altitudes up to 18,288 meters, during a cabin decompression, and subsequent aircraft descent. During cabin decompressions in the F-104 and F-15, cabin pressure altitude was measured at various aircraft angles of attack, Mach numbers, and altitudes to determine the effect of the aerodynamic slipstream on the cabin altitude. *System Development Corp., Edwards, California. 1055. Jenkins, J. M.: Correlation of Predicted and Measured Thermal Stresses on an Advanced Aircraft Structure With Dissimilar Materials. NASA TM-72865, H-1092, June 1979, 79N27088, #. Additional information was added to a growing data base from which estimates of finite element model complexities can be made with respect to thermal stress analysis. The manner in which temperatures were smeared to the finite element grid points was examined from the point of view of the impact on thermal stress calculations. The general comparison of calculated and measured thermal stresses is quite good and there is little doubt that the finite element approach provided by NASTRAN results in correct thermal stress calculations. Discrepancies did exist between measured and calculated values in the skin and the skin/frame junctures. The problems with predicting skin thermal stress were attributed to inadequate temperature inputs to the structural model rather than modeling insufficiencies. The discrepancies occurring at the skin/frame juncture were most likely due to insufficient modeling elements rather than temperature problems. 1056. Kurtenbach, F. J.: Evaluation of a Simplified Gross Thrust Calculation Technique Using Two Prototype F100 Turbofan Engines in an Altitude Facility. NASA TP-1482, H-1061, June 1979, 79N26057, #. The technique which relies on afterburner duct pressure measurements and empirical corrections to an ideal one dimensional flow analysis to determine thrust is presented. A comparison of the calculated and facility measured thrust values is reported. The simplified model with the engine manufacturer’s gas generator model are compared. The evaluation was conducted over a range of Mach numbers from 0.80 to 2.00 and at altitudes from 4020 meters to 15,240 meters. The effects of variations in inlet total temperature from standard day conditions were explored. Engine conditions were varied from those normally scheduled for flight. The technique was found to be accurate to a twice standard deviation of 2.89 percent, with accuracy a strong function of afterburner duct pressure difference. 1057. Wilner, D. O.: AIFTDS Stand-Alone RMDU Flight Test Report. NASA TM-72866, H-1099, July 1979, 79N28167, #. 183 The remote multiplexer/digitizer unit for the airborne integrated flight test data system was subjected to a flight test environment in order to study its dynamic response and that of its associated instrumentation circuitry during an actual flight test. The shielding schemes and instrumentation used are described and the data obtained are analyzed. 1058. Rediess, H. A.; and McIver, D. E.: Avionics and Controls Research and Technology. NASA CP-2061, L-12498, 1979, 79N15898, #. The workshop provided a forum for industry and universities to discuss the state-of-the-art, identify the technology needs and opportunities, and describe the role of NASA in avionics and controls research. 1059. Wolowicz, C. H.; Brown, J. S., Jr.; and Gilbert, W. P.: Similitude Requirements and Scaling Relationships as Applied to Model Testing. NASA TP-1435, H-1022, August 1979, 79N30176, #. The similitude requirements for the most general test conditions are presented. These similitude requirements are considered in relation to the scaling relationships, test technique, test conditions (including supersonic flow), and test objectives. Particular emphasis is placed on satisfying the various similitude requirements for incompressible and compressible flow conditions. For free flying models tests, the test velocities for incompressible flow are scaled from Froude number similitude requirements and those for compressible flow are scaled from Mach number similitude requirements. The limitations of various test techniques are indicated, with emphasis on the free flying model. 1060. Petersen, K. L.: Flight Control Systems Development of Highly Maneuverable Aircraft Technology/HiMAT/Vehicle. AIAA Paper 79-1789, presented at the American Institute of Aeronautics and Astronautics, Aircraft Systems and Technology Meeting, New York, New York, August 20–22, 1979, 79A47878, #. The highly maneuverable aircraft technology (HiMAT) program was conceived to demonstrate advanced technology concepts through scaled-aircraft flight tests using a remotely piloted technique. Closed-loop primary flight control is performed from a ground-based cockpit, utilizing a digital computer and up/down telemetry links. A backup flight control system for emergency operation resides in an onboard computer. The onboard systems are designed to provide failoperational capabilities and utilize two microcomputers, dual uplink receiver/decoders, and redundant hydraulic actuation and power systems. This paper discusses the design and validation of the primary and backup digital flight control systems as well as the unique pilot and specialized systems interfaces. 1061. *Hartmann, G.; *Stein, G.; and Powers, B.: Flight Test Experience With an Adaptive Control System Using a Maximum Likelihood Parameter Estimation Technique. AIAA Paper 79-1702, Collection of Technical Papers, Guidance and Control Conference, Boulder, Colorado, August 6–8, 1979, (see A79-45351 19-12), 79A45357, #. The flight test performance of an adaptive control system for the F-8 DFBW aircraft is summarized. The adaptive system is based on explicit identification of surface effectiveness parameters which are used for gain scheduling in a command augmentation system. Performance of this control law under various design parameter variations is presented. These include variations in test signal level, sample rate, and identification channel structure. Flight performance closely matches analysis and simulation predictions from previous references. *Honeywell, Inc., Minneapolis, Minnesota. 1062. *Lorincz, D. J.; and Friend, E. L.: Water Tunnel Visualization of the Vortex Flows of the F-15. AIAA Paper 79-1649, Collection of Technical Papers, Atmospheric Flight Mechanics Conference for Future Space Systems, Boulder, Colorado, August 6–8, 1979, (see A79-45302 19-01), 79A45325, #. Flow visualization studies were conducted in a diagnostic water tunnel to provide details of the wing, glove, and forebody vortex flow fields of the F-15 aircraft over a range of angles of attack and sideslip. Both the formation and breakdown of the vortex flow as a function of angle of attack and sideslip are detailed for the basic aircraft configuration. Additional tests showed that the wing upper surface vortex flows were sensitive to variations in an inlet mass flow ratio and an inlet cowl cross-sectional shape, were tested in addition to the basic forebody. Asymmetric forebody vortices were observed at zero sideslip and high angles of attack on each forebody. A large nose boom was added to each of the three forebodies, and it was observed that the turbulent wake shed from the boom disrupted the forebody vortices. *Northrop Corp., Hawthorne, California. 1063. Smith, J. W.: Analysis of a Lateral Pilot-Induced Oscillation Experienced on the First Flight of the YF-16 Aircraft. NASA TM-72867, September 1979, 79N31220, #. In order to compare and assess potential improvements, two control systems were modeled; the original first flight or prototype aircraft system, and a modification of the prototype system, which essentially reduced the overall gain for the takeoff and landing phase. In general, the overall system gain reduction of the modified flight control system was sufficient to avoid lateral pilot-induced oscillation tendencies. Lowering the system gain reduced the tendency to rate saturate, which resulted in correspondingly higher critical pilot gains for the same control input. 184 1064. Walker, H. J.: Performance Evaluation Method for Dissimilar Aircraft Designs. NASA RP-1042, H-1064, September 1979, 79N30139, #. A rationale is presented for using the square of the wingspan rather than the wing reference area as a basis for nondimensional comparisons of the aerodynamic and performance characteristics of aircraft that differ substantially in planform and loading. Working relationships are developed and illustrated through application to several categories of aircraft covering a range of Mach numbers from 0.60 to 2.00. For each application, direct comparisons of drag polars, lift-to-drag ratios, and maneuverability are shown for both nondimensional systems. The inaccuracies that may arise in the determination of aerodynamic efficiency based on reference area are noted. Span loading is introduced independently in comparing the combined effects of loading and aerodynamic efficiency on overall performance. Performance comparisons are made for the NACA research aircraft, lifting bodies, century-series fighter aircraft, F-111A aircraft with conventional and supercritical wings, and a group of supersonic aircraft including the B-58 and XB-70 bomber aircraft. An idealized configuration is included in each category to serve as a standard for comparing overall efficiency. 1065. Jenkins, J. M.: Criteria for Representing Circular Arc and Sine Wave Spar Webs by Non-Curved Elements. NASA TM-72869, H-1106, October 1979, 79N33499, #. The basic problem of how to simply represent a curved web of a spar in a finite element structural model was addressed. The ratio of flat web to curved web axial deformations and longitudinal rotations were calculated using NASTRAN models. Multiplying factors were developed from these calculations for various web thicknesses. These multiplying factors can be applied directly to the area and moment of inertia inputs of the finite element model. This allows the thermal stress relieving configurations of sine wave and circular arc webs to be simply accounted for in finite element structural models. 1066. Jenkins, J. M.: Effect of Element Density on the NASTRAN Calculated Mechanical and Thermal Stresses of a Spar. NASA TM-72868, H-1104, October 1979, 79N32151, #. A NASTRAN model of a spar was examined to determine the sensitivity of calculated axial thermal stresses and bending stresses to changes in element density of the model. The thermal stresses calculated with three different element densities resulted in drastically differing values. The position of the constraint also significantly affected the value of the calculated thermal stresses. Mechanical stresses calculated from an applied loading were insensitive to element density. 1067. Cooper, D. W.; and *James, R.: Shuttle Orbiter Radar Cross-Sectional Analysis. NASA TM-72870, H-1095, October 1979, 80N10276, #. Theoretical and model simulation studies on signal to noise levels and shuttle radar cross section are described. Premission system calibrations, system configuration, and postmission system calibration of the tracking radars are described. Conversion of target range, azimuth, and elevation into radar centered east north vertical position coordinates are evaluated. The location of the impinging rf energy with respect to the target vehicles body axis triad is calculated. Cross section correlation between the two radars is presented. *James and Associates, Lancaster, California. 1068. *Muirhead, V. U.; and Saltzman, E. J.: Reduction of Aerodynamic Drag and Fuel Consumption for Tractor-Trailer Vehicles. Journal of Energy, Vol. 3, No. 5, September–October 1979, pp. 279–284, 80A16948, #. Wind-tunnel tests were performed on a scale model of a cabover-engine tractor-trailer vehicle and several modifications of the model. Results from two of the model configurations were compared with full-scale drag data obtained from similar configurations during coast-down tests. Reductions in fuel consumption derived from these tests are presented in terms of fuel quantity and dollar savings per vehicle year, based on an annual driving distance of 160,900 km (100,000 mi). The projected savings varied from 13,001 (3435) to 25,848 (6829) liters (gallons) per year which translated to economic savings from $3435 to about $6829 per vehicle year for an operating speed of 88.5 km/h (55 mph) and wind speeds near the national average of 15.3 km/h (9.5 mph). The estimated cumulative fuel savings for the entire U.S. fleet of cab-over-engine tractor, van-type trailer combinations ranged from 4.18 million kl (26.3 million bbl) per year for a low-drag configuration to approximately twice that amount for a more advanced configuration. *Kansas University, Lawrence, Kansas. 1069. Maine, Richard E.; and Iliff, Kenneth W.: Maximum Likelihood Estimation of Translational Acceleration Derivatives From Flight Data. Journal of Aircraft, Vol. 16, No. 10, pp. 674–679, October 1979. (See also 995.) This paper shows that translational acceleration derivatives, such as pitching moment due to rate of change of angle of attack, Cm , can be estimated from flight data with the use ˙ α of appropriately design maneuvers. No new development of estimation methodology is necessary to analyze these maneuvers. Flight data from a T-37B airplane were used to verify that Cm , could be estimated from rolling maneuvers. ˙ α 1070. Steers, L. L.: Flight-Measured Afterbody Pressure Coefficients From an Airplane Having Twin Side-By-Side Jet Engines for Mach Numbers From 0.6 to 1.6. NASA TP-1549, H-1066, November 1979, 80N11035, #. Afterbody pressure distribution data were obtained in flight from an airplane having twin side-by-side jet exhausts. The data were obtained in level flight at Mach numbers from 0.60 to 1.60 and at elevated load factors for Mach numbers of 0.60, 0.90, and 1.20. The test altitude varied from 2300 meters (7500 feet) to 15,200 meters (50,000 feet) over a speed range that provided a matrix of constant Mach number and constant unit Reynolds number test conditions. The results of the fullscale flight afterbody pressure distribution program are presented in the form of plotted pressure distributions and tabulated pressure coefficients with Mach number, angle of attack, engine nozzle pressure ratio, and unit Reynolds number as controlled parameters. 1071. Iliff, K. W.: Aircraft Identification Experience. AGARD Lecture Series No. 104 on Parameter Identification, AGARD LS-104, paper 6, November 1979, (see N80-19094 10-05), 80N19100, #. Important aspects of estimating the unknown coefficients of the aircraft equations of motion from dynamic flight data are presented. The primary topic is the application of the maximum likelihood estimation technique. Basic considerations that must be addressed in the estimation of stability and control derivatives from conventional flight maneuvers are discussed. Some complex areas of estimation (such as estimation in the presence of atmospheric turbulence, estimation of acceleration derivatives, and analysis of maneuvers where both kinematic and aerodynamic coupling are present) are also discussed. 1072. Swaroop, R.; Brownlow, J. D.; and Winter, William R.: Extreme Mean and Its Applications. NASA TM-81346, December 1979, 80N13863, #. Extreme value statistics obtained from normally distributed data are considered. An extreme mean is defined as the mean of p-th probability truncated normal distribution. An unbiased estimate of this extreme mean and its large sample distribution are derived. The distribution of this estimate even for very large samples is found to be nonnormal. Further, as the sample size increases, the variance of the unbiased estimate converges to the Cramer-Rao lower bound. The computer program used to obtain the density and distribution functions of the standardized unbiased estimate, and the confidence intervals of the extreme mean for any data are included for ready application. An example is included to demonstrate the usefulness of extreme mean application. 185 1073. Sefic, W. J.: Friction Characteristic of Steel Skids Equipped With Skegs on a Lakebed Surface. NASA TM-81347, H-1111, December 1979, 80N13027, #. The coefficient of friction was determined for steel skids with and without skegs. The addition of a 1.27 centimeter deep skeg caused the coefficient of friction to increase from an average value of .36 to .53, a 47 percent increase over the flat skid. The addition of a .64 centimeter deep skeg increased the friction coefficient from .36 to .46, a 16 percent increase over the flat skid. Comparisons are made with data for similar test conditions obtained during the X-15 program. 1074. Carr, P. C.; and Gilbert, W. P.: Effects of Fuselage Forebody Geometry on Low-Speed Lateral-Directional Characteristics of Twin-Tail Fighter Model at High Angles of Attack. NASA TP-1592, L-13270, December 1979, 80N13002, #. Low-speed, static wind-tunnel tests were conducted to explore the effects of fighter fuselage forebody geometry on lateral-directional characteristics at high angles of attack and to provide data for general design procedures. Effects of eight different forebody configurations and several add-on devices (e.g., nose strakes, boundary-layer trip wires, and nose booms) were investigated. Tests showed that forebody design features such as fineness ratio, cross-sectional shape, and add-on devices can have a significant influence on both lateral-directional and longitudinal aerodynamic stability. Several of the forebodies produced both lateral-directional symmetry and strong favorable changes in lateral-directional stability. However, the same results also indicated that such forebody designs can produce significant reductions in longitudinal stability near maximum lift and can significantly change the influence of other configuration variables. The addition of devices to highly tailored forebody designs also can significantly degrade the stability improvements provided by the clean forebody. force-stiffness data did not result in any predictions of buckling failure. It was, therefore, concluded that the panels were conservatively designed as a result of design constraints and assumptions of panel eccentricities. The analysis programs calculated strains and stresses competently. Comparisons between calculated and measured structural deflections showed good agreement. The test program offered a positive demonstration of the beaded panel concept subjected to room-temperature load conditions. 1076. Lasagna, P. L.; Mackall, K. G.; Burcham, F. W., Jr.; and Putnam, T. W.: Landing Approach Airframe Noise Measurements and Analysis. NASA TP-1602, January 1980, 80N15028, #. Flyover measurements of the airframe noise produced by the AeroCommander, JetStar, CV-990, and B-747 airplanes are presented for various landing approach configurations. Empirical and semiempirical techniques are presented to correlate the measured airframe noise with airplane design and aerodynamic parameters. Airframe noise for the jetpowered airplanes in the clean configuration (flaps and gear retracted) was found to be adequately represented by a function of airplane weight and the fifth power of airspeed. Results show the airframe noise for all four aircraft in the landing configuration (flaps extended and gear down) also varied with the fifth power of airspeed, but this noise level could not be represented by the addition of a constant to the equation for clean-configuration airframe noise. 1077. *Dougherty, N. S., Jr.; and *Fisher, D. F.: Boundary-Layer Transition on a 10-Deg Cone—Wind Tunnel/Flight Correlation. AIAA Paper 80-0154. Presented at AIAA 18th Aerospace Sciences Meeting, Pasadena, California, January 14–16, 1980, 80A22737, #. Boundary-layer transition location measurements were made on a 10-deg sharp cone in 23 wind tunnels of the US and Europe and in flight. The data were acquired at subsonic, transonic, and supersonic Mach numbers over a range of unit Reynolds numbers to obtain an improved understanding of wind tunnel flow quality influence. Cone surface microphone measurements showed Tollmien-Schlichting waves present. Transition location defined by pitot probe measurements showed transition Reynolds number to be correlatable to cone surface disturbance amplitude within ±20 percent for the majority of tunnel and flight data. *ARO, Inc., Arnold Air Force Station, Tennessee. 1078. *Bangert, L. H.; Burcham, F. W., Jr.; and Mackall, K. G.: YF-12 Inlet Suppression of Compressor Noise—First Results. AIAA Paper 80-0099. Presented at AIAA 18th Aerospace Sciences Meeting, Pasadena, California, January 14–16, 1980, 80A34537, #. An aeroacoustic test program was performed with a YF-12 aircraft at aerodynamic data that could determine the cause of 186 1980 Technical Publications 1075. Fields, R. A.; Reardon, L. F.; and Siegel, W. H.: Loading Tests of a Wing Structure for a Hypersonic Aircraft. NASA TP-1596, H-1046, January 1980, 80N15068, #. Room-temperature loading tests were conducted on a wing structure designed with a beaded panel concept for a Mach 8 hypersonic research airplane. Strain, stress, and deflection data were compared with the results of three finite-element structural analysis computer programs and with design data. The test program data were used to evaluate the structural concept and the methods of analysis used in the design. A force stiffness technique was utilized in conjunction with load conditions which produced various combinations of panel shear and compression loading to determine the failure envelope of the buckling critical beaded panels The inlet noise suppression observed earlier. The first results of the test program are presented here. There was no indication that the flow was close to choking. The data indicated significant reduction in sound pressure level (SPL) across the strut and bypass region at frequencies near the blade passing. Far-field data showed that the maximum sound pressure level near the blade-passing frequency was at zero degrees from the inlet centerline. *Lockheed-California Co., Burbank, California. 1079. Plant, T. J.; Nugent, J.; and Davis, R. A.: FlightMeasured Effects of Boattail Angle and Mach Number on the Nozzle Afterbody Flow of a Twin-Jet Fighter. AIAA Paper 80-0110. Presented at AIAA 18th Aerospace Sciences Meeting, Pasadena, California, January 14–16, 1980, 80A23009, #. The paper presents the flight-measured nozzle afterbody surface pressures and engine exhaust nozzle pressure-area integrated axial force coefficients on a twin-jet fighter for varying boattail angles. The objective of the tests was to contribute to a full-scale flight data base applicable to the nozzle afterbody drag of advanced tactical fighter concepts. The data were acquired during the NASA F-15 Propulsion/Airframe Interactions Flight Research Program. Nozzle boattail angles from 7.7 deg to 18.1 deg were investigated. Results are presented for cruise angle of attack at Mach numbers from 0.6 to 2.0 at altitudes from 20,000 to 45,000 feet. The data show the nozzle axial force coefficients to be a strong function of nozzle boattail angle and Mach number. 1080. Ko, W. L.: Structural Properties of Superplastically Formed/Diffusion-Bonded Orthogonally Corrugated Core Sandwich Plates. AIAA Paper 80-0304. Presented at AIAA 18th Aerospace Sciences Meeting, Pasadena, California, January 14–16, 1980, 80A18305, #. (See also 1196.) This paper describes a new superplastically formed/diffusion-bonded (SPF/DB) orthogonally corrugated sandwich structure, and presents formulae and the associated plots for evaluating the effective elastic constants for the core of this new sandwich structure. Comparison of structural properties of this new sandwich structure with the conventional honeycomb core sandwich structure was made under the condition of equal sandwich density. It was found that the SPF/DB orthogonally corrugated sandwich core has higher transverse shear stiffness than the conventional honeycomb sandwich core. However, the former has lower stiffness in the sandwich core thickness direction than the latter. 1081. Maine, R. E.; and Iliff, K. W.: Estimation of the Accuracy of Dynamic Flight-Determined Coefficients. AIAA Paper 80-0171. Presented at AIAA 18th Aerospace 187 Sciences Meeting, Pasadena, California, January 14–16, 1980, 80A17700, #. This paper discusses means of assessing the accuracy of maximum likelihood parameter estimates obtained from dynamic flight data. The commonly used analytical predictors of accuracy are compared from both statistical and simplified geometric standpoints. Emphasizing practical considerations, such as modeling error, the accuracy predictions are evaluated with real and simulated data. Improved computations of the Cramer-Rao bound to correct large discrepancies caused by colored noise and modeling error are presented. This corrected Cramer-Rao bound is the best available analytical predictor of accuracy. Engineering judgment, aided by such analytical tools, is the final arbiter of accuracy estimation. 1082. Arnaiz, H. H.; *Peterson, J. B., Jr.; and **Daugherty, J. C.: Wind-Tunnel/Flight Correlation Study of Aerodynamic Characteristics of a Large Flexible Supersonic Cruise Airplane (XB-70-1). 3: A Comparison Between Characteristics Predicted From Wind-Tunnel Measurements and Those Measured in Flight. NASA TP-1516, H-1079, March 1980, 80N17986, #. A program was undertaken by NASA to evaluate the accuracy of a method for predicting the aerodynamic characteristics of large supersonic cruise airplanes. This program compared predicted and flight-measured lift, drag, angle of attack, and control surface deflection for the XB-70-1 airplane for 14 flight conditions with a Mach number range from 0.76 to 2.56. The predictions were derived from the wind-tunnel test data of a 0.03-scale model of the XB-70-1 airplane fabricated to represent the aeroelastically deformed shape at a 2.5 Mach number cruise condition. Corrections for shape variations at the other Mach numbers were included in the prediction. For most cases, differences between predicted and measured values were within the accuracy of the comparison. However, there were significant differences at transonic Mach numbers. At a Mach number of 1.06 differences were as large as 27 percent in the drag coefficients and 20 deg in the elevator deflections. A brief analysis indicated that a significant part of the difference between drag coefficients was due to the incorrect prediction of the control surface deflection required to trim the airplane. *NASA Langley Research Center, Hampton, Virginia. **NASA Ames Research Center, Moffett Field, California. 1083. McWithey, R. R.; Royster, D. M.; and Ko, W. L.: Compression Panel Studies for Supersonic Cruise Vehicles. NASA TP-1617, L-13525, NAS 1.60:1617, March 1980, 83N28099, #. Results of analytical and experimental studies are summarized for titanium, boron fiber reinforced aluminum matrix composite, Borsic fiber reinforced aluminum matrix composite, and titanium sheathed Borsic fiber reinforced aluminum matrix composite stiffened panels. The results indicate that stiffened panels with continuous joints (i.e., brazed, diffusion bonded or adhesive bonded joints) are more structurally efficient than geometrically similar panels with discrete joints (i.e., spotwelded or bolted joints). In addition, results for various types of fiber reinforced aluminum matrix stiffened panels indicate that titanium sheathed Borsic fiber reinforced aluminum matrix composite panels are the most structurally efficient. Analytical results are also presented for graphite fiber reinforced polyimide matrix composite stiffened panels and superplastically formed and diffusion bonded titanium sandwich panels. 1084. Ehernberger, L. J.: Clear Air Turbulence: Historical Comments. NASA CP-2139, FAA-RD-80-67. NASA Marshall Space Flight Center Proceedings Fourth Annual Workshop on Meteorological and Environmental Inputs to Aviation Systems, March 1980 (see N81-14555 05-47), pp. 71–81, 81N14562, #. The basic reference material for gust design criteria are cited. The status of clear air turbulence meteorology (forecasting and detection) is discussed. The directions of further research technology is indicated. 1085. Smith, J. W.; and Edwards, J. W.: Design of a Nonlinear Adaptive Filter for Suppression of Shuttle Pilot-Induced Oscillation Tendencies. NASA TM-81349, H-1119, April 1980, 80N21355, #. Analysis of a longitudinal pilot-induced oscillation (PIO) experienced just prior to touchdown on the final flight of the space shuttle’s approach landing tests indicated that the source of the problem was a combination of poor basic handling qualities aggravated by time delays through the digital flight control computer and rate limiting of the elevator actuators due to high pilot gain. A nonlinear PIO suppression (PIOS) filter was designed and developed to alleviate the vehicle’s PIO tendencies by reducing the gain in the command path. From analytical and simulator studies it was shown that the PIOS filter, in an adaptive fashion, can attenuate the command path gain without adding phase lag to the system. With the pitch attitude loop of a simulated shuttle model closed, the PIOS filter increased the gain margin by a factor of about two. 1086. Ko, W. L.: Elastic Constants for Superplastically Formed/Diffusion-Bonded Corrugated Sandwich Core. NASA TP-1562, H-1094, May 1980, 80N23677, #. Formulas and associated graphs for evaluating the effective elastic constants for a superplastically formed/diffusion bonded (SPF/DB) corrugated sandwich core, are presented. A comparison of structural stiffnesses of the sandwich core and a honeycomb core under conditions of equal sandwich core density was made. The stiffness in the thickness direction of the optimum SPF/DB corrugated core (that is, triangular truss core) is lower than that of the honeycomb 188 core, and that the former has higher transverse shear stiffness than the latter. 1087. Sisk, T. R.; and Matheny, N. W.: Precision Controllability of the YF-17 Airplane. NASA TP-1677, H-1089, May 1980, 80N23327, #. A flying qualities evaluation conducted on the YF-17 airplane permitted assessment of its precision controllability in the transonic flight regime over the allowable angle of attack range. The precision controllability (tailchase tracking) study was conducted in constant-g and windup turn tracking maneuvers with the command augmentation system (CAS) on, automatic maneuver flaps, and the caged pipper gunsight depressed 70 mils. This study showed that the YF-17 airplane tracks essentially as well at 7 g’s to 8 g’s as earlier fighters did at 4 g’s to 5 g’s before they encountered wing rock. The pilots considered the YF-17 airplane one of the best tracking airplanes they had flown. Wing rock at the higher angles of attack degraded tracking precision, and lack of control harmony made precision controllability more difficult. The revised automatic maneuver flap schedule incorporated in the airplane at the time of the tests did not appear to be optimum. The largest tracking errors and greatest pilot workload occurred at high normal load factors at low angles of attack. The pilots reported that the high-g maneuvers caused some tunnel vision and that they found it difficult to think clearly after repeated maneuvers. 1088. Larson, T. J.; Flechner, S. G.; and Siemers, P. M., III: Wind Tunnel Investigation of an All Flush Orifice Air Data System for a Large Subsonic Aircraft. NASA TP-1642, H-1085, May 1980, 80N23304, #. The results of a wind tunnel investigation on an all flush orifice air data system for use on a KC-135A aircraft are presented. The investigation was performed to determine the applicability of fixed all flush orifice air data systems that use only aircraft surfaces for orifices on the nose of the model (in a configuration similar to that of the shuttle entry air data system) provided the measurements required for the determination of stagnation pressure, angle of attack, and angle of sideslip. For the measurement of static pressure, additional flush orifices in positions on the sides of the fuselage corresponding to those in a standard pitot-static system were required. An acceptable but less accurate system, consisting of orifices only on the nose of the model, is defined and discussed. 1089. Swaroop, R.; Brownlow, J. D.; Ashworth, G. R.; and Winter, W. R.: Bivariate Normal, Conditional and Rectangular Probabilities: A Computer Program With Applications. NASA TM-81350, May 1980, 80N24099, #. Some results for the bivariate normal distribution analysis are presented. Computer programs for conditional normal probabilities, marginal probabilities, as well as joint probabilities for rectangular regions are given: routines for computing fractile points and distribution functions are also presented. Some examples from a closed circuit television experiment are included. 1090. Ko, W. L.: Elastic Stability of Superplastically Formed/Diffusion-Bonded Orthogonally Corrugated Core Sandwich Plates. AIAA Paper 80-0683. Presented at 21st AIAA Structures, Structural Dynamics and Materials Conference, Seattle, Washington, May 12–14, 1980, pp. 167–176, 80A35005, #. The paper concerns the elastic buckling behavior of a newly developed superplastically formed/diffusion-bonded (SPF/DB) orthogonally corrugated core sandwich plate. Uniaxial buckling loads were calculated for this type of sandwich plate with simply supported edges by using orthotropic sandwich plate theory. The buckling behavior of this sandwich plate was then compared with that of an SPF/DB unidirectionally corrugated core sandwich plate under conditions of equal structural density. It was found that the buckling load for the former was considerably higher than that of the latter. 1091. Reed, R. D.: Flight Research Techniques Utilizing Remotely Piloted Research Vehicles. AGARD LS-108, Paper 8, Aircraft Assessment and Acceptance Testing, (ISBN-02-835-0266-3), May 1980, (see N80-31329 22-01), AD-A088530, 80N31337, #. The use of the remotely piloted research vehicle (RPRV) in aeronautical research is surveyed. The flight test experience that has been acquired with several types of RPRV’s including those with a pilot in the loop is emphasized. The approaches utilized range from the simplest and least expensive of vehicles, such as the Minisniffer, to the sophisticated highly maneuverable aircraft technology (HiMAT) RPRV. The advantages and disadvantages of RPRV’s are discussed, as well as safety considerations. The ground rules set early in a program can profoundly affect program cost effectiveness and timeliness. 1092. Bender, G. L.; Arnaiz, H.; Ottomeyer, D.; Woratschek, R.; Higgins, L.; and Tulloch, J. S.: Preliminary Airworthiness Evaluation AH-1S Helicopter With Ogee Tip Shape Rotor Blades. AD-A089625, USAAEFA-77-25, May 1980, 81N10061, #. The United States Army Aviation Engineering Flight Activity conducted a Preliminary Airworthiness Evaluation of the AH-1S helicopter with OGEE tip-shape main rotor blades to determine if any improvement in performance or handling qualities resulted from replacing the K747 blades. Additionally, the acoustics signature of the OGEE blades were measured by the US Army Research and Technology Laboratories (Aeromechanics Lab). Tests were conducted at Edwards Air Force Base (elevation 2302 feet) and Coyote Flats (elevation 9980 feet), California from 1 November 1979 through 8 April 1980. Forty-five test flights were flown for a 189 total of 36.6 productive hours (63.2 total hours). Both hover and level flight performance were degraded by installation of OGEE tip-shape main rotor blades. Low-speed handling qualities were unaffected by the OGEE blades. Other handling qualities tests were not accomplished. Results of acoustics tests will be reported by the laboratories under a separate cover. 1093. Larson, T. J.; and *Siemers, P. M., III: Subsonic Tests of an All-Flush-Pressure-Orifice Air Data System. NASA TP-1871, H-1122. Presented at the 1980 Air Data Systems Conference, Colorado, Springs, Colorado, May 1980, June 1981, 81N26144, #. The use of an all-flush-pressure-orifice array as a subsonic air data system was evaluated in flight and wind tunnel tests. Two orifice configurations were investigated. Both used orifices arranged in a cruciform pattern on the airplane nose. One configuration also used orifices on the sides of the fuselage for a source of static pressure. The all-nose-orifice configuration was similar to the shuttle entry air data system (SEADS). The flight data were obtained with a KC-135A airplane. The wind tunnel data were acquired with a 0.035-scale model of the KC-135A airplane. With proper calibration, several orifices on the vertical centerline of the vehicle’s nose were found to be satisfactory for the determination of total pressure and angle of attack. Angle of sideslip could be accurately determined from pressure measurements made on the horizontal centerline of the aircraft. Orifice pairs were also found that provided pressure ratio relationships suitable for the determination of Mach number. The accuracy that can be expected for the air data determined with SEADS during subsonic orbiter flight is indicated. *NASA Langley Research Center, Hampton, Virginia. 1094. Ko, W. L.: Comparison of Structural Behavior of Superplastically Formed/Diffusion-Bonded Sandwich Structures and Honeycomb Core Sandwich Structures. NASA TM-81348, June 1980, 80N23685, #. A superplasticity formed/diffusion-bonded (SPF/DB) orthogonally corrugated core sandwich structure is discussed and its structural behavior is compared to that of a conventional honeycomb core sandwich structure. The stiffness and buckling characteristics of the two types of sandwich structures are compared under conditions of equal structural density. It is shown that under certain conditions, the SPF/DB orthogonally corrugated core sandwich structure is slightly more efficient than the optimum honeycomb core (square-cell core) sandwich structure. However, under different conditions, this effect can be reversed. 1095. Ko, W. L.: Mode I Fracture Behavior of Nonlinear Materials. International Journal of Fracture, Vol. 16, June 1980, pp. 207–219, 80A39939, #. A finite element analysis is presented for the Mode I fracture behavior of cracked plates (stationary crack) made of different nonlinear materials (elastoplastic or elastic locking). The assumed stress-strain behavior of the nonlinear materials was piecewise linear. For each given “stationary” crack size, the corresponding critical remote tensile stress was calculated based on the maximum crack-tip stress failure criterion. It was found that in the log-log plots of the critical remote stress versus critical crack length, the fracture data of the piecewise linear materials obey a “stepwise-linear-inverse-square-root” fracture law. It is also shown how the fracture data of the piecewise linear materials can be fitted by proper piecewise graphical shifting of the classical “inverse-square-root” fracture curve for the linearly elastic materials. 1096. Myers, A. F.; and Sheets, S. G.: Qualification of HiMAT Flight Systems. Proceedings, Seventh Annual Technical Symposium of the Association for Unmanned Vehicle Systems, Dayton, Ohio, June 16–18, 1980, pp. 1–10, 81A22603, #. The highly maneuverable aircraft technology (HiMAT) remotely piloted research vehicle is discussed with emphasis on the advanced composite and metallic structures, digital fly-by-wire controls, and digitally implemented integrated propulsion control systems. Techniques used to qualify the systems for flight are examined. Computation and simulation of the HiMAT system are investigated in relation to CyberVarian simulation. The techniques used in flight qualification are complicated by ground based flight critical systems and severe onboard volume constraints imposed by the scale design. 1097. Bennett, George; Enevoldson, Einar; Gera, Joe; and Patton, Jim: Pilot Evaluation of Sailplane Handling Qualities. Technical Soaring, Vol. 5, No. 4, June 1980, pp. 3–14. 1098. Gilyard, G. B.; and Burken, J. J.: Development and Flight Test Results of an Autothrottle Control System at Mach 3 Cruise. NASA TP-1621, H-1090, July 1980, 80N26328, #. Flight test results obtained with the original Mach hold autopilot designed the YF-12C airplane which uses elevator control and a newly developed Mach hold system having an autothrottle integrated with an altitude hold autopilot system are presented. The autothrottle tests demonstrate good speed control at high Mach numbers and high altitudes while simultaneously maintaining control over altitude and good ride qualities. The autothrottle system was designed to control either Mach number or knots equivalent airspeed (KEAS). Excellent control of Mach number or KEAS was obtained with the autothrottle system when combined with altitude hold. Ride qualities were significantly better than with the conventional Mach hold system. 1099. *McRae, D. S.; Fisher, D. F.; and **Peake, D. J.: A Computational and Experimental Study of High 190 Reynolds Number Viscous/Inviscid Interaction About a Cone at High Angle of Attack. AIAA Paper 80-1422. Presented at 13th AIAA Fluid and Plasma Dynamics Conference, Snowmass, Colorado, July 14–16 1980, 80A44492, #. The flow over a 5 deg problem for the flow over aircraft forebodies. A computational method utilizing the conically symmetric Navier-Stokes equations is used to obtain theoretical flow results which are compared with experimental data from the Ames Research Center 6- by 6-Foot Wind Tunnel and with results from a cone model sting mounted on an F-15 aircraft. The computed results agree well with the wind-tunnel data but less well with the flight data. Modification of the algebraic turbulence model was necessary to reflect an apparent lower turbulence level in flight than was present in the wind tunnel. *USAF, Flight Dynamics Laboratory, Wright-Patterson AFB, Ohio. **NASA Ames Research Center, Moffett Field, California. 1100. Pool, A.; Sanderson, K. C.; and *Ferrell, K. R.: Helicopter Flight Test Instrumentation. AGARD-AG160-VOL-10, AD-A089909, July 1980, 80N33406. The helicopter characteristics with which instrumentation must contend with are discussed. Typical tests that are conducted are outlined. Major aircraft components and systems which may be instrumented are listed and suggestions are made for sensors, locations, and installation. Instrumentation requirements are summarized. A sample instrumentation management technique is also presented. *Army Aviation Research and Development Command, Edwards AFB, California. 1101. *Wuest, W.; Pool, A.; and Sanderson, K. C.: Pressure and Flow Measurement. AGARD-AG-160VOL-11, AD-A090961, July 1980, 80N33407. The evolution of flight test instrumentation systems during the last decade reflects the radical changes of electronic measuring techniques. Nevertheless the basic principles of measurement methods are essentially unchanged and the sensors for flow and pressure measurements have experienced only slight changes. The fundamentals of flow and pressure measurements are explained from the viewpoint of flight test instrumentation. An overview of modern instrumentation is given with important applications to altitude measurement, vertical and horizontal speed measurement, boundary layer, wake and engine flow measurement. The scope of this manual is to give self consistent information on the different techniques and systems and to give references for a more detailed study of special techniques. *DFVLR, Goettingen, West Germany. 1102. Ko, William L.: Elastic Constants for Superplastically Formed/Diffusion-Bonded Sandwich Structures. AIAA Journal, Vol. 18, No. 8, pp. 986–987, August 1980. (See also 1047.) Formulas for evaluating the effective elastic constants for a superplastically formed/diffusion-bonded (SPF/DB) unidirectionally corrugated sore sandwich structure like that shown in Fig. 1 are presented. This structure is formed by diffusion bonding three superplastic alloy sheets (two face sheets and one core sheet) in the selected areas and the superplastically expanding the multiple sheet pack inside a die cavity by using gas pressure. Thus, this structure is slightly different from the conventional corrugated-core sandwich structure because the corrugation leg does not have uniform thickness. Because of superplastic expansion, the diagonal segment of the corrugation leg is always thinner than the flat segment (crest or trough) of the corrugation leg, which has a thickness that is nearly the pre-expansion thickness. Thus, the results given by Ref. 2 for a corrugatedcore sandwich plate cannot be used in the present structure without considerable modification. 1103. Balakrishnan, A. V.; and Edwards, J. W.: Calculation of the Transient Motion of Elastic Airfoils Forced by Control Surface Motion and Gusts. NASA TM-81351, H-1125, August 1980, 80N32329, #. The time-domain equations of motion of elastic airfoil sections forced by control surface motions and gusts were developed for the case of incompressible flow. Extensive use was made of special functions related to the inverse transform of Theodorsen’s function. Approximations for the special cases of zero stream velocity, small time, large and time are given. A numerical solution technique for the solution of the general case is given. Examples of the exact transient response of an airfoil are presented. 1104. *Nguyen, L. T.; *Gilbert, W. P.; Gera, J.; Iliff, K. W.; and Enevoldson, E. K.: Application of High-Alpha Control System Concepts to a Variable-Sweep Fighter Airplane. AIAA Paper 80-1582. AIAA Atmospheric Flight Mechanics Conference, Danvers, Massachusetts, August 11–13, 1980, 80A50098, #. The use of control system design to enhance high-angle-ofattack flying qualities and departure/spin resistance has become an accepted and widely used approach for modern fighter aircraft. NASA and the Navy are currently conducting a joint research program to investigate the application of this technology to the F-14. The paper discusses the results of this program within the context of its contributions to advancing high-alpha control system technology. General topics covered include (1) analysis and design tools, (2) control system design approach, and (3) flight test approach and results. *NASA Langley Research Center, Hampton, Virginia. 191 1105. *Hoyt, C. E.; Kempel, R. W.; and Larson, R. R.: Backup Flight Control System for a Highly Maneuverable Remotely Piloted Research Vehicle. AIAA Paper 80-1761. Presented at AIAA Guidance and Control Conference, Danvers, Massachusetts, August 11–13, 1980, pp. 283–289, 80A45548, #. NASA is currently conducting flight tests of a remotely piloted subscale advanced fighter configuration as part of the Highly Maneuverable Aircraft Technology (HiMAT) program. This paper describes the initial development, user modification, and flight test experience of a back-up control system (BCS) contained within one of two onboard microprocessors. The development of the BCS proceeded in two distinct steps: the initial contractor development of control laws and logic to satisfy BCS design objectives, and user modifications required to satisfy operational requirements. A brief resume of flight qualification procedures and pilot comments is presented. *Teledyne Ryan Aeronautical, San Diego, California. 1106. *Travassos, R. H.; *Gupta, N. K.; Iliff, K. W.; and Maine, R.: Determination of an Oblique Wing Aircraft’s Aerodynamic Characteristics. AIAA Paper 80-1630. Presented at AIAA Atmospheric Flight Mechanics Conference, Danvers, Massachusetts, August 11–13, 1980. pp. 608–618, 80A45918, #. In this paper, the integration of wind tunnel and flight test procedures are studied for specifying aerodynamic model forms. A procedure is described which employs a stepwise regression method to systematically determine model structures and F-ratio statistics to rank the importance of each aerodynamic coefficient within a given model. Application of this technique and wind tunnel procedures to an oblique-wing aircraft indicate that the aircraft’s measured and estimated response are in good agreement at both small and large wing skew angles. *Systems Control, Inc., Palo Alto, California. 1107. Shafer, M. F.: Low Order Equivalent Models of Highly Augmented Aircraft Determined From Flight Data Using Maximum Likelihood Estimation. AIAA Paper 80-1627. Presented at AIAA Atmospheric Flight Mechanics Conference, Danvers, Massachusetts, August 11–13, 1980, pp. 572–582, 80A45915, #. (See also 1220.) This paper presents the results of a study of the feasibility of using low order equivalent mathematical models of a highly augmented aircraft, the F-8 digital fly-by-wire (DFBW), for flying qualities research. Increasingly complex models were formulated and evaluated using flight data and maximum likelihood estimation techniques. The aircraft actuator was modeled alone first. Next the equivalent derivatives were used to model the longitudinal unaugmented F-8 DFBW aircraft dynamics. The most complex model incorporated a pure time shift of the pilot input, a first order lag, and the basic longitudinal airframe model. This same model was then implemented for the F-8 DFBW aircraft in a highly augmented mode. Excellent matching of the dynamics resulted for this model, indicating that low order equivalent models which are good representations of the highly augmented F-8 DFBW aircraft can be formulated with these methods. 1108. Berry, D. T.; Powers, B. G.; Szalai, K. J.; and Wilson, R. J.: A Summary of an In-Flight Evaluation of Control System Pure Time Delays During Landing Using the F-8 DFBW Airplane. AIAA Paper 80-1626. Presented at AIAA Atmospheric Flight Mechanics Conference, Danvers, Massachusetts, August 11–13, 1980, pp. 561–571, 80A45914, #. (See also 1197.) An in-flight investigation of the effect of pure time delays on low L/D space shuttle type landing tasks was undertaken. The results indicate that the sensitivity of the pilot ratings to changes in pure time delay in pitch is strongly affected by the task and only slightly affected by changes in control system augmentation mode. Low L/D spot landings from a lateral offset were twice as sensitive to pure time delay as normal low L/D landings. For comparison purposes, formation flying was also investigated, and was found to be less sensitive to time delay than the landing tasks. 1109. Maine, R. E.; and Iliff, K. W.: Formulation and Implementation of a Practical Algorithm for Parameter Estimation With Process and Measurement Noise. AIAA Paper 80-1603. Presented at AIAA Atmospheric Flight Mechanics Conference, Danvers, Massachusetts, August 11–13, 1980, pp. 397–411, 80A45896, #. (See also 1184, 1209.) A new formulation is proposed for the problem of parameter estimation of dynamic systems with both process and measurement noise. The formulation gives estimates that are maximum likelihood asymptotically in time. The means used to overcome the difficulties encountered by previous formulations are discussed. It is then shown how the proposed formulation can be efficiently implemented in a computer program. A computer program using the proposed formulation is available in a form suitable for routine application. Examples with simulated and real data are given to illustrate that the program works well. 1110. Powers, B. G.: Experience With an Adaptive Stick-Gain Algorithm to Reduce Pilot-InducedOscillation Tendencies. AIAA Paper 80-1571. Presented at Atmospheric Flight Mechanics Conference, Danvers, Massachusetts, August 11–13, 1980, pp. 142–154, 80A45870, # As part of a program to improve the approach and landing characteristics of the Space Shuttle, the NASA Dryden Flight 192 Research Center has developed an adaptive algorithm that varies the longitudinal stick gearing to reduce the Shuttle’s tendency for pilot-induced oscillation (PIO). This paper describes the algorithm, which is known as the PIO suppresser, and discusses some of the tradeoffs involved in optimizing the system. The results of fixed-base, movingbase, and in-flight simulations of the PIO suppresser are presented. 1111. Iliff, K. W.: Stall/Spin Flight Results for the Remotely Piloted Spin Research Vehicle. AIAA Paper 80-1563. Presented at AIAA Atmospheric Flight Mechanics Conference, Danvers, Massachusetts, August 11–13, 1980, pp. 62–75, 80A45862, #. The unmanned, remotely piloted, unpowered, spin research vehicle was used to evaluate the effects of the nose boom and of a wind tunnel-designed nose strake on the vehicle’s stall/spin characteristics. The flight-determined directional stability derivatives and the attempted spin entries indicated that the vehicle with a nose strake had increased resistance to departure and spin. The acquisition of high quality steady spin data for this vehicle was made possible by the remotely piloted technique. The zero control smooth spin modes were found to be highly repeatable for a given configuration and to vary with forebody configuration. Several spin recovery techniques, including a nose parachute, are also evaluated. 1112. Myers, A.: Simulation Use in the Development and Validation of HiMAT Flight Software. AGARD CP-272, Paper 22. AGARD Advan. in Guidance and Control Systems Using Digital Tech., (ISBN-92-835-0247-7), August 1980, AD-A076146, 80N14039, #. The use of real time simulation in the development and validation of flight software for the highly maneuverable aircraft technology (HiMAT) remotely piloted research vehicle is described. Four simulations are interfaced with varying amounts of actual flight hardware to produce dynamic system operation. 1113. Larson, T. J.; and *Siemers, P. M., III: Use of Nose Cap and Fuselage Pressure Orifices for Determination of Air Data for Space Shuttle Orbiter Below Supersonic Speeds. NASA TP-1643, H-1096, September 1980, 80N32389, #. Wind tunnel pressure measurements were acquired from orifices on a 0.1 scale forebody model of the space shuttle orbiter that were arranged in a preliminary configuration of the shuttle entry air data system (SEADS). Pressures from those and auxiliary orifices were evaluated for their ability to provide air data at subsonic and transonic speeds. The orifices were on the vehicle’s nose cap and on the sides of the forebody forward of the cabin. The investigation covered a Mach number range of 0.25 to 1.40 and an angle of attack range from 4 deg. to 18 deg. An air data system consisting of nose cap and forebody fuselage orifices constitutes a complete and accurate air data system at subsonic and transonic speeds. For Mach numbers less than 0.80 orifices confined to the nose cap can be used as a complete and accurate air data system. Air data systems that use only flush pressure orifices can be used to determine basic air data on other aircraft at subsonic and transonic speeds. *NASA Langley Research Center, Hampton, Virginia. 1114. Tang, M. H.: A Modified T-Value Method for Selection of Strain Gages for Measuring Loads. NASA TM-85464. Presented at Western Regional Gage Comm. Soc. for Expt. Stress Analysis, China, Lake, California, September 16–17, 1980. 1980, 84N75770. 1115. Szalai, K. J.; Larson, R. R.; and Glover, R. D.: Flight Experience With Flight Control Redundancy Management. AGARD LS 109, Paper 8. Fault Tolerance Design and Redundancy Management Tech., (ISBN92-835-0274-4), September 1980, (see N81-11266 02-31), (AD-A090849, 81N11274, #. Flight experience with both current and advanced redundancy management schemes was gained in recent flight research programs using the F-8 digital fly by wire aircraft. The flight performance of fault detection, isolation, and reconfiguration (FDIR) methods for sensors, computers, and actuators is reviewed. Results of induced failures as well as of actual random failures are discussed. Deficiencies in modeling and implementation techniques are also discussed. The paper also presents comparison off multisensor tracking in smooth air, in turbulence, during large maneuvers, and during maneuvers typical of those of large commercial transport aircraft. The results of flight tests of an advanced analytic redundancy management algorithm are compared with the performance of a contemporary algorithm in terms of time to detection, false alarms, and missed alarms. The performance of computer redundancy management in both iron bird and flight tests is also presented. 1116. Andrews, W. H.; Sim, A. G.; Monaghan, R. C.; Felt, L. R.; McMurtry, T. C.; and *Smith, R. C.: AD-1 Oblique Wing Aircraft Program. SAE Paper 801180. Society of Automotive Engineers, Aerospace Congress and Exposition, Los Angeles, California, October 13–16, 1980, 81A34193. The oblique wing concept for super- and subsonic transport was assessed by analysis and wind tunnel radio control model and remotely piloted vehicle testing. A one-sixth scale wind tunnel model and a low speed manned oblique wing research airplane (AD-1) were developed. Model wind tunnel test data on dynamic structural response characteristics were used in a simulator to develop the control system. The airplane is of simple design with fiber glass skin, weight of approximately 2100 lbs and speeds of up to 175 knots at altitudes up to 15,000 ft. Flight testing will investigate handling and flying qualities, oblique wing flight control characteristics, aeroelastic wing design and will compare actual with predicted aerodynamic characteristics. Nineteen flights were made at 12,000 to 13,000 feet with speeds of 100-160 knots. Flutter clearance as a function of wing sweep angle is now under investigation. *NASA Ames Research Center, Moffett Field, California ECN-13302B AD-1 Oblique Wing Airplane 1117. Maine, R. E.; and Iliff, K. W.: User’s Manual for MMLE3, a General FORTRAN Program for Maximum Likelihood Parameter Estimation. NASA TP-1563, H-1084, November 1980, 81N12744, #. A user’s manual for the FORTRAN IV computer program MMLE3 is described. It is a maximum likelihood parameter estimation program capable of handling general bilinear dynamic equations of arbitrary order with measurement noise and/or state noise (process noise). The theory and use of the program is described. The basic MMLE3 program is quite general and, therefore, applicable to a wide variety of problems. The basic program can interact with a set of user written problem specific routines to simplify the use of the program on specific systems. A set of user routines for the aircraft stability and control derivative estimation problem is provided with the program. 1118. Quinn, R. D.; and Gong, L.: In-Flight BoundaryLayer Measurements on a Hollow Cylinder at a Mach Number of 3.0. NASA TP-1764, H-1101, November 1980, 81N12361, #. Skin temperatures, shear forces, surface static pressures, boundary layer pitot pressures, and boundary layer total temperatures were measured on the external surface of a 193 hollow cylinder that was 3.04 meters long and 0.437 meter in diameter and was mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 (a local Mach number of 2.9) and at wall to recovery temperature ratios of 0.66 to 0.91. The local Reynolds number had a nominal value of 4,300,000 per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. In addition, boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor was obtained from the heat transfer and skin friction measurements. The measured data are compared with several boundary layer prediction methods. 1119. Tang, M. H.; and Sheldon, R. G.: A Modified T-Value Method for Selection of Strain Gages for Measuring Loads on a Low Aspect Ratio Wing. NASA TP-1748, H-1108, November 1980, 81N12066, #. A technique which may be useful for selecting strain gages for use in load equations is described. The technique is an adaptation of the previously used T-value method and is applied to a multispar structure. The technique, called the modified T-value method, is used to reduce the number of strain gages used in a load equation from twelve to two. A parallel reduction is made by calculating relative equation accuracies from three applied load distributions. The equations developed from the modified T-value method proved to be accurate more consistently than the T-value method. 1120. Jenkins, J. M.: The Effect of Thermal Stresses on the Integrity of Three Built-Up Aircraft Structures. NASA TM-81352, H-1138, November 1980, 81N12064, #. A Mach 6 flight was simulated in order to examine heating effects on three frame/skin specimens. The specimens included: a titanium truss frame with a lockalloy skin; a stainless steel z-frame with a lockalloy skin; and a titanium z-frame with a lockalloy skin. Thermal stresses and temperature were measured on these specimens for the purpose of examining their efficiency, performance, and integrity. Measured thermal stresses were examined with respect to material yield strengths, buckling criteria, structural weight, and geometric locations. Principal thermal stresses were studied from the standpoint of uniaxial stress assumptions. Measured thermal stresses were compared to predicted values. 1121. *Runyan, L. J.; and Steers, L. L.: Boundary Layer Stability Analysis of a Natural Laminar Flow Glove on the F-111 TACT Airplane. Viscous Flow Drag Reduction Symposium, Dallas, Texas, Technical Papers, November 7–8, 1979, pp. 17–32, 81A26503, #. A natural laminar flow airfoil has been developed as a part of the aircraft energy efficiency program. A NASA flight program incorporating this airfoil into partial wing gloves on the F-111 TACT airplane was scheduled to start in May, 1980. In support of this research effort, an extensive boundary layer stability analysis of the partial glove has been conducted. The results of that analysis show the expected effects of wing leading-edge sweep angle, Reynolds number, and compressibility on boundary layer stability and transition. These results indicate that it should be possible to attain on the order of 60% laminar flow on the upper surface and 50% laminar flow on the lower surface for sweep angles of at least 20 deg, chord Reynolds numbers of 25 × 10 to the 6th and Mach numbers from 0.81 to 0.85. *Boeing Commercial Airplane Co., Seattle, Washington. 1122. Anon.: Research and Technology Accomplishments. NASA TM-102949, NAS 1.15:102949, 1980, 90N70764. 1123. Andrews, W. H.: The Oblique Wing-Research Aircraft. Society of Experimental Test Pilots Tech. Rev., Vol. 15, No. 1, 1980 (see N80-33337 24-01), pp. 4–5, 80N33338. The AD-1 airplane was designed as a low cost, low speed manned research tool to evaluate the flying qualities of the oblique wing concept. The airplane is constructed primarily of foam and fiberglass and incorporates simplicity in terms of the onboard systems. There are no hydraulics, the control system is cable and torque tube, and the electrical systems consist of engine driven generators which power the battery for engine start, cockpit gages, trim motors, and the onboard data system. The propulsion systems consist of two Microturbo TRS-18 engines sea level trust rated at 220 pounds. The airplane weighs approximately 2100 pounds and has a performance potential in the range of 200 knots and an altitude of 15,000 feet. 1981 Technical Publications 1124. Kurtenbach, F. J.; and Burcham, F. W., Jr.: Flight Evaluation of a Simplified Gross Thrust Calculation Technique Using an F100 Turbofan Engine in an F-15 Airplane. NASA TP-1782, H-1118, January 1981, 81N15000, #. A simplified gross thrust calculation technique was evaluated in flight tests on an F-15 aircraft using prototype F100-PW-100 engines. The technique relies on afterburner duct pressure measurements and empirical corrections to an ideal one-dimensional analysis to determine thrust. In-flight gross thrust calculated by the simplified method is compared to gross thrust calculated by the engine manufacturer’s gas 194 generator model. The evaluation was conducted at Mach numbers from 0.6 to 1.5 and at altitudes from 6000 meters to 13,700 meters. The flight evaluation shows that the simplified gross thrust method and the gas generator method agreed within plus or minus 3 percent. The discrepancies between the data generally fell within an uncertainty band derived from instrumentation errors and recording system resolution. 1125. *Peake, D. J.; Fisher, D. F.; and **McRae, D. S.: Flight Experiments With a Slender Cone at Angle of Attack. AIAA Paper 81-0337, Presented at the 19th AIAA Aerospace Sciences Meeting, St. Louis, Missouri, January 1981, 81A20761, #. The three-dimensional leeward separation about a 5 deg semi-angle cone at an 11 deg angle of attack was investigated in flight, in the wind tunnel, and by numerical computations. The test conditions were Mach numbers of 0.6, 1.5, and 1.8 at Reynolds numbers between 7 and 10 million based on freestream conditions and a 30-inch wetted length or surface. The surface conditions measured included mean static and fluctuating pressures; skin friction magnitudes and separation line positions were obtained using obstacle blocks. The mean static pressures from flight and wind tunnel were in good agreement. The computed results gave the same distributions, but were slightly more positive in magnitude. The experimentally measured primary and secondary separation line locations compared closely with computed results. There were substantial differences in level and in trend between the surface root-mean-square pressure fluctuations obtained in flight and in the wind tunnel, due, it is thought, to a relatively high acoustic disturbance level in the tunnel compared with the quiescent conditions in flight. *3-D Flowz, Inc., Moffett Field, California. **USAF, Moffett Field, California. 1126. Ehernberger, L. J.; and Guttman, N. B.: Climatological Characteristics of High Altitude Wind Shear and Lapse Rate Layers. NASA TM-81353, H-1132, February 1981, 81N18608, #. Indications of the climatological distribution of wind shear and temperature lapse and inversion rates as observed by rawinsonde measurements over the western United States are recorded. Frequencies of the strongest shear, lapse rates, and inversion layer strengths were observed for a 1 year period of record and were tabulated for the lower troposphere, the upper troposphere, and five altitude intervals in the lower stratosphere. Selected bivariate frequencies were also tabulated. Strong wind shears, lapse rates, and inversion are observed less frequently as altitude increases from 175 millibars to 20 millibars. On a seasonal basis the 195 frequencies were higher in winter than in summer except for minor stratospheric wind reversal in the spring and fall. 1127. Monaghan, R. C.: Description of the HiMAT Tailored Composite Structure and Laboratory Measured Vehicle Shape Under Load. NASA TM-81354, H-1144, February 1981, 81N18047, #. The aeroelastically tailored outer wing and canard of the highly maneuverable aircraft technology (HiMAT) vehicle are closely examined and a general description of the overall structure of the vehicle is provided. Test data in the form of laboratory measured twist under load and predicted twist from the HiMAT NASTRAN structural design program are compared. The results of this comparison indicate that the measured twist is generally less than the NASTRAN predicted twist. These discrepancies in twist predictions are attributed, at least in part, to the inability of current analytical composite materials programs to provide sufficiently accurate properties of matrix dominated laminates for input into structural programs such as NASTRAN. 1128. Jarvis, C. R.; and Szalai, K. J.: Ground and Flight Test Experience With a Triple Redundant Digital Fly by Wire Control System. NASA CP-2172. NASA Langley Research Center Advan. Aerodyn. and Active Controls, February 1981, (see N81 19001 10-01), pp. 67–84, 81N19006, #. A triplex digital fly by wire flight control system was developed and installed in an F-8C aircraft to provide fail operative, full authority control. Hardware and software redundancy management techniques were designed to detect and identify failures in the system. Control functions typical of those projected for future actively controlled vehicles were implemented. 1129. Steers, L. L.: Natural Laminar Flow Flight Experiment. NASA CP-2172. NASA Langley Research Center Advan. Aerodyn. and Active Controls, February 1981, (see N81-19001 10-01), pp. 135–144, 81N19010, #. A supercritical airfoil section was designed with favorable pressure gradients on both the upper and lower surfaces. Wind tunnel tests were conducted in the Langley 8 Foot Transonic Pressure Tunnel. The outer wing panels of the F-111 TACT airplane were modified to incorporate partial span test gloves having the natural laminar, flow profile. Instrumentation was installed to provide surface pressure data as well as to determine transition location and boundary layer characteristics. The flight experiment encompassed 19 flights conducted with and without transition fixed at several locations for wing leading edge sweep angles which varied from 10 to 26 at Mach numbers from 0.80 to 0.85 and altitudes of 7620 meters and 9144 meters. Preliminary results indicate that a large portion of the test chord experienced laminar flow. 1130. Montoya, L. C.: KC-135 Winglet Flight Results. NASA CP-2172. NASA Langley Research Center Advan. Aerodyn. and Active Controls, February 1981, (see N81-19001 10-01), pp. 145–156, 81N19011, #. Three KC-135 winglet configurations were flight tested for cant/incidence angles of 15 deg/–4 deg, 15 deg/–2 deg, and 0 deg/–4 deg, as well as the basic wing. The flight results for the 15 deg/–4 deg and basic wing configurations confirm the wind tunnel predicted 7% incremental decrease in total drag at cruise conditions. The 15 deg/–4 configuration flight measured wing and winglet pressure distributions, loads, stability and control, flutter, and buffet also correlate well with predicted values. The only unexpected flight results as compared with analytical predictions is a flutter speed decrease for the 0 deg/–4 deg configuration. The 15 deg/–2 deg configuration results show essentially the same incremental drag reduction as the 15 deg/–4 deg configuration; however, the flight loads are approximately 30% higher for the 15 deg/–2 deg configuration. The drag data for the 0 deg/–4 deg configuration show only a flight drag reduction. and a number of technological advances applicable to future fighter aircraft were demonstrated. The aircraft control system uses airborne and ground-based computers which communicate via uplink and downlink telemetry. Antenna radiation patterns are normally much less than ideal for continuous reception or transmission for all aircraft attitudes. After flight qualification and testing on other aircraft, a frequency diversity concept and an antenna diversity concept were implemented on the HiMAT vehicle to obtain omnidirectional telemetry coverage. 1132. Maine, R. E.: User’s Manual for SYNC: A FORTRAN Program for Merging and TimeSynchronizing Data. NASA TM-81355, March 1981, 81N19793, #. The FORTRAN 77 computer program SYNC for merging and time synchronizing data is described. The program SYNC reads one or more input files which contain either synchronous data frames or time-tagged data points, which can be compressed. The program decompresses and time synchronizes the data, correcting for any channel time skews. Interpolation and hold last value synchronization algorithms are available. The output from SYNC is a file of time synchronized data frames at any requested sample rate. 1133. *Weaver, E. A.; Ehernberger, L. J.; **Gary, B. L.; †Kurkowski, R. L.; ††Kuhn, P. M.; and ††Stearns, L. P.: The 1979 Clear Air Turbulence Flight Test Program. NASA Langley Research Center, CP-2178, The 1980 Aircraft Safety and Operating Problems, Part 1, March 1981, (see N81-19035 10-03), pp. 293–311, 81N19050, #. The flight experiments for clear air turbulence (CAT) detection and measurement concepts are described. The test were conducted over the western part of the United States during the winter season of 1979 aboard NASA’s Galileo 2 flying laboratory. A carbon dioxide pulsed Doppler lidar and an infrared radiometer were tested for the remote detection and measurement of CAT. Two microwave radiometers were evaluated for their ability to provide encounter warning and altitude avoidance information. *NASA Marshall Space Flight Center, Huntsville, Alabama. **NASA Jet Propulsion Laboratory, Pasadena, California. †NASA Ames Research Center, Moffett Field, California. ††National Oceanic and Atmospheric Administration, Department of Commerce, Washington, D. C. 1134. *Chambers, J. R.; and Iliff, K. W.: Estimation of Dynamic Stability Parameters From Drop Model Flight KC-135 Airplane With Winglets ECN-11484 1131. Harney, P. F.: Diversity Techniques for Omnidirectional Telemetry Coverage of the HiMAT Research Vehicle. NASA TP-1830, H-1133, March 1981, 81N20074, #. The highly maneuverable aircraft technology (HiMAT) remotely piloted research vehicle (RPRV) was flight tested 196 Tests. AGARD LS-114. Presented at the NATO, AGARD Lecture Series on Dynamic Stability Parameters, Moffett Field, California, March 2–5, 1981, and Rhode-SaintGenese, Belgium, March 16–19, 1981, 81A35551, #. (See also 1141.) A recent NASA application of a remotely-piloted drop model to studies of the high angle-of-attack and spinning characteristics of a fighter configuration has provided an opportunity to evaluate and develop parameter estimation methods for the complex aerodynamic environment associated with high angles of attack. The paper discusses the overall drop model operation including descriptions of the model, instrumentation, launch and recovery operations, piloting concept, and parameter identification methods used. Static and dynamic stability derivatives were obtained for an angle-of-attack range from –20 deg to 53 deg. The results of the study indicated that the variations of the estimates with angle of attack were consistent for most of the static derivatives, and the effects of configuration modifications to the model (such as nose strakes) were apparent in the static derivative estimates. The dynamic derivatives exhibited greater uncertainty levels than the static derivatives, possibly due to nonlinear aerodynamics, model response characteristics, or additional derivatives. *NASA Langley Research Center, Hampton, Virginia. 1135. Barber, M. R.; and *Tymczyszyn, J. J.: Wake Vortex Attenuation Flight Tests: A Status Report. NASA Langley Research Center, CP-2178, The 1980 Aircraft Safety and Operating Problems, Part 2, March 1981, (see N81-19056 10-03), pp. 387–408, 81N19057, #. (See also 1201.) Flight tests were conducted to evaluate the magnitude of aerodynamic attenuation of the wake vortices of large transport aircraft that can be achieved through the use of static spoiler deflection and lateral control oscillation. These methods of attenuation were tested on Boeing B-747 and Lockheed L-1011 commercial transport aircraft. Evaluations were made using probe aircraft, photographic and visual observations, and ground based measurements of the vortex velocity profiles. The magnitude of attenuation resulting from static spoiler deflection was evaluated both in and out of ground effect. A remotely piloted QF-86 drone aircraft was used to probe the attenuated vortices in flight in and out of ground effect, and to make landings behind an attenuated B-747 airplane at reduced separation distances. *FAA, Los Angeles, California. ECN-7848 L-1011 Airplane With Smokers 1136. Maine, Richard E.; and Iliff, Kenneth W.: Use of Cramer-Rao Bounds on Flight Data With Colored Residuals. Journal of Guidance and Control, Vol. 4, No. 2, pp. 207–213, March–April 1981. This paper discusses the use of the Cramer-Rao bound as a means of assessing the accuracy of maximum likelihood parameter estimates obtained from dynamic flight data. Emphasizing practical considerations such as modeling error, the Cramer-Rao bound is evaluated with real and simulated data. Improved computations of the bound to correct large discrepancies caused by colored noise and modeling error are presented. This corrected Cramer-Rao bound is the best available analytical predictor of accuracy. 1137. Borek, R. W.: Practical Aspects of Instrumentation Installation in Support of Subsystem Testing. AGARD CP-299, Paper 25. AGARD Subsystem Testing and Flight Test Instr., (ISBN-92-835-0290-6), April 1981, (see N81-29065 20-01), AD-A101016, 81N29090, #. Some of the problems associated with using military specification MIL-W-5088H as a guideline for wire gage selection are discussed. Examples of proper use of this specification as a criterion for interfacing wire bundles and connectors are provided. The quantitative results of 22 projects that have used the technique known as sneak analysis are reviewed and examples are given. 197 1138. Redin, P. C.: Application of a Performance Modeling Technique to an Airplane With Variable Sweep Wings. NASA TP-1855, H-1131, May 1981, 81N24048, #. A performance modeling concept previously applied to an F-104F G and a YF-12C airplane was applied to an F-111A airplane. This application extended the concept to an airplane with variable sweep wings. The performance model adequately matched flight test data for maneuvers flown at different wing sweep angles at maximum afterburning and intermediate power settings. For maneuvers flown at less than intermediate power, including dynamic maneuvers, the performance model was not validated because the method used to correlate model and in-flight power setting was not adequate. Individual dynamic maneuvers were matched successfully by using adjustments unique to each maneuver. 1139. Jenkins, J. M.: A Comparison of Laboratory Measured Temperatures With Predictions for a Spar/Skin Type Aircraft Structure. NASA TM-81359, May 1981, 81N23067, # A typical spar/skin aircraft structure was heated nonuniformly in a laboratory and the resulting temperatures were measured. The heat transfer NASTRAN computer program was used to provide predictions. Calculated temperatures based on a thermal model with conduction, radiation, and convection features compared closely to measured spar temperatures. Results were obtained without the thermal conductivity, specific heat, or emissivity with temperature. All modes of heat transfer (conduction, radiation, and convection) show to affect the magnitude and distribution of structural temperatures. 1140. Carter, A. L.; and Sims, R. L.: Comparison of Theoretical Predictions of Orbiter Airloads With Wind Tunnel and Flight Test Results for a Mach Number of 0.52. NASA TM-81358, May 1981, 81N23066, #. The measurement and prediction of wing airloads for space shuttle orbiter 101 during approach and landing tests is discussed. Strain gage instrumentation, calibration, and flight data processing are covered along with wind tunnel and simulator results. The generation of theoretical predictions using the FLEXSTAB computer program is described, and the results are compared to experimental measurements. 1141. *Chambers, J. R.; and Iliff, K. W.: Estimation of Dynamic Stability Parameters From Drop Model Flight Tests. AGARD LS-114, AGARD Dyn. Stability Parameters, May 1981, (see N81-31105 22-01), 81N31116, #. (See also 1134.) The overall remotely piloted drop model operation, descriptions, instrumentation, launch and recovery operations, piloting concept, and parameter identification methods are discussed. Static and dynamic stability 198 derivatives were obtained for an angle attack range from –20 deg to 53 deg. It is indicated that the variations of the estimates with angle of attack are consistent for most of the static derivatives, and the effects of configuration modifications to the model were apparent in the static derivative estimates. *NASA Langley Research Center, Hampton, Virginia. 1142. Ehernberger, L. J.: Aspects of Clear Air Turbulence Severity Forecasting and Detection. Proceedings, 1st International Conference on Aviation Weather Systems, Montreal, Canada, May 4–6, 1981. pp. 146–152, 82A45823. Factors influencing the accuracy of the forecasts of incidences of clear air turbulence (CAT) are discussed, along with techniques for improved verification. Descriptive ranking terms for the intensity of CAT events, ranging from light to extreme, are developed, and meteorological parameters used for predictions are reviewed, including jetstream core location, vertical and horizontal wind shears, stable layers, tropopause height, trough speed, 500-mb vorticity, surface fronts, pressure centers and cyclogenesis, and wind speeds near mountain ridges. Methods of remote detection of CAT, particularly by using radiometry sensitive to the IR water vapor band, are noted to have had some success in detecting actual CAT events and decreasing false alarms. Statistical aspects of CAT encounter severity are discussed, including the establishment of confidence intervals for thresholds of detection of CATs of varying intensities. 1143. Borek, Robert W.: Some Practical Aspects of Minimizing the Weight and Volume of Airborne Instrumentation Systems. Fourth AGARD/DUT/CIT Special Course on Flight Test Instrumentation, Delft, the Netherlands, May 11–22, 1981. 1144. Maine, R. E.: Programmer’s Manual for MMLE3, a General FORTRAN Program for Maximum Likelihood Parameter Estimation. NASA TP-1690, H-1105, June 1981, 81N27813, #. The MMLE3 is a maximum likelihood parameter estimation program capable of handling general bilinear dynamic equations of arbitrary order with measurement noise and/or state noise (process noise). The basic MMLE3 program is quite general and, therefore, applicable to a wide variety of problems. The basic program can interact with a set of user written problem specific routines to simplify the use of the program on specific systems. A set of user routines for the aircraft stability and control derivative estimation problem is provided with the program. The implementation of the program on specific computer systems is discussed. The structure of the program is diagrammed, and the function and operation of individual routines is described. Complete listings and reference maps of the routines are included on microfiche as a supplement. Four test cases are discussed; listings of the input cards and program output for the test cases are included on microfiche as a supplement. 1145. *Mancuso, R. L.; *Endlich, R. M.; and Ehernberger, L. J.: An Objective Isobaric/Isentropic Technique for Upper Air Analysis. Monthly Weather Review, Vol. 109, June 1981, pp. 1326–1334, 81A43360. An objective meteorological analysis technique is presented whereby both horizontal and vertical upper air analyses are performed. The process used to interpolate grid-point values from the upper-air station data is the same as for grid points on both an isobaric surface and a vertical cross-sectional plane. The nearby data surrounding each grid point are used in the interpolation by means of an anisotropic weighting scheme, which is described. The interpolation for a grid-point potential temperature is performed isobarically; whereas wind, mixing-ratio, and pressure height values are interpolated from data that lie on the isentropic surface that passes through the grid point. Two versions (A and B) of the technique are evaluated by qualitatively comparing computer analyses with subjective handdrawn analyses. The objective products of version A generally have fair correspondence with the subjective analyses and with the station data, and depicted the structure of the upper fronts, tropopauses, and jet streams fairly well. The version B objective products correspond more closely to the subjective analyses, and show the same strong gradients across the upper front with only minor smoothing. *SRI International Atmospheric Science Center, Menlo Park, California. 1146. Sim, Alex G.: Althaus on Airfoils. Astronautics & Aeronautics, Vol. 19, No. 6, June 1981. Book review of Profilpolaren fur Den ModelflugWindKanalmessungen an Profilen im Kristischen Reynoldszahlbereich (Polars for Airfoils for Model Airplanes-Wind Tunnel Measurements on Airfoils at Critical Reynolds Numbers) by Dieter Althaus. 1147. Maine, R. E.; and Iliff, K. W.: The Theory and Practice of Estimating the Accuracy of Dynamic FlightDetermined Coefficients. NASA-RP-1077, H-1128, July 1981, 81N27865, #. Means of assessing the accuracy of maximum likelihood parameter estimates obtained from dynamic flight data are discussed. The most commonly used analytical predictors of accuracy are derived and compared from both statistical and simplified geometrics standpoints. The accuracy predictions are evaluated with real and simulated data, with an emphasis on practical considerations, such as modeling error. Improved computations of the Cramer-Rao bound to correct 199 large discrepancies due to colored noise and modeling error are presented. The corrected Cramer-Rao bound is shown to be the best available analytical predictor of accuracy, and several practical examples of the use of the Cramer-Rao bound are given. Engineering judgment, aided by such analytical tools, is the final arbiter of accuracy estimation. 1148. Weil, J.; and Powers, Bruce G.: Correlation of Predicted and Flight Derived Stability and Control Derivatives With Particular Application to Tailless Delta Wing Configurations. NASA TM-81361, July 1981, 81N26154, #. Flight derived longitudinal and lateral-directional stability and control derivatives were compared to wind-tunnel derived values. As a result of these comparisons, boundaries representing the uncertainties that could be expected from wind-tunnel predictions were established. These boundaries provide a useful guide for control system sensitivity studies prior to flight. The primary application for this data was the space shuttle, and as a result the configurations included in the study were those most applicable to the space shuttle. The configurations included conventional delta wing aircraft as well as the X-15 and lifting body vehicles. 1149. Hughes, D. L.: Comparison of Three Thrust Calculation Methods Using In-Flight Thrust Data. NASA TM-81360, H-1141, July 1981, 81N30111, #. The gross thrust of an experimental airplane was determined by each method using the same flight maneuvers and generally the same data parameters. Coefficients determined from thrust stand calibrations for each of the three methods were then extrapolated to cruise flight conditions. The values of total aircraft gross thrust calculated by the three methods for cruise flight conditions agreed within ±3 percent. The disagreement in the values of thrust calculated by the different techniques manifested itself as a bias in the data. There was little scatter (0.5 percent) for the thrust levels examined in flight. 1150. *Barrett, W. J.; *Rembold, J. P.; Burcham, F. W.; and Myers, L.: Flight Test of a Full Authority Digital Electronic Engine Control System in an F-15 Aircraft. AIAA Paper 81-1501. Presented at the 17th AIAA, SAE, and ASME Joint Propulsion Conference, Colorado Springs, Colorado, July 27–29, 1981, 81A40912, #. (See also 1242.) The Digital Electronic Engine Control (DEEC) system considered is a relatively low cost digital full authority control system containing selectively redundant components and fault detection logic with capability for accommodating faults to various levels of operational capability. The DEEC digital control system is built around a 16-bit, 1.2 microsecond cycle time, CMOS microprocessor, microcomputer system with approximately 14 K of available memory. Attention is given to the control mode, component bench testing, closed loop bench testing, a failure mode and effects analysis, sea-level engine testing, simulated altitude engine testing, flight testing, the data system, cockpit, and real time display. *United Technologies Corp., Government Products Div., West Palm Beach, Florida. 1151. Anon.: Preliminary Analysis of STS-1 Entry Flight Data. NASA TM-81363, August 1981, 81N29153, #. A preliminary analysis of data acquired during the first shuttle orbiter reentry is presented. Heating levels were higher than predicted. Variations in measured versus predicted lift to drag ratio and trim are discussed, as are plots showing time histories of control surface and jet activity. The confidence felt in the stability and control derivatives is only fair. Confidence in the derivatives extracted for Mach numbers below 3.5 is especially weak, because these derivatives were affected by sideslip data contaminated by wind and turbulence, nonindependent rudder motions, and buffet. The sources of the data used are described. Recommendations are presented for changes to the Aerodynamic Data Book, and for planning future flights. 1152. Ayers, T. G.; and Hallissy, J. B.: Historical Background and Design Evolution of the Transonic Aircraft Technology Supercritical Wing. NASA TM-81356, H-1148, August 1981, 81N32116, #. Two dimensional wind tunnel test results obtained for supercritical airfoils indicated that substantial improvements in aircraft performance at high subsonic speeds could be achieved by shaping the airfoil to improve the supercritical flow above the upper surface. Significant increases in the drag divergence Mach number, the maximum lift coefficient for buffer onset, and the Mach number for buffet onset at a given lift coefficient were demonstrated for the supercritical airfoil, as compared with a NACA 6 series airfoil of comparable thickness. These trends were corroborated by results from three dimensional wind tunnel and flight tests. Because these indicated extensions of the buffet boundaries could provide significant improvements in the maneuverability of a fighter airplane, an exploratory wind tunnel investigation was initiated which demonstrated that significant aerodynamic improvements could be achieved from the direct substitution of a supercritical airfoil on a variable wing sweep multimission airplane model. 1153. Sims, R. L.: User’s Manual for FSLIP-3, FLEXSTAB Loads Integration Program. NASA TM-81364, H-1158, August 1981, 81N30815, #. The FSLIP program documentation and user’s manual is presented. As a follow on program to the FLEXSTAB computer analysis system, the primary function of this FORTRAN IV program is to integrate panel pressure coefficients computed by FLEXSTAB to obtain total shear, 200 bending, and torque airloads on various surfaces, summed relative to user specified axes. The program essentially replaces the ALOADS module in FLEXSTAB with expanded capabilities and flexibility. As such, FSLIP is generalized to work on any FLEXSTAB model or other pressure data if in a compatible format. 1154. Berry, D. T.: Flying Qualities Criteria and Flight Control Design. AIAA Paper 81-1823. Guidance and Control Conference, Albuquerque, New Mexico, August 19–21, 1981, pp. 411–415, (see A81-44076 21-12), 81A44127, #. Despite the application of sophisticated design methodology, newly introduced aircraft continue to suffer from basic flying qualities deficiencies. Two recent meetings, the DOD/NASA Workshop on Highly Augmented Aircraft Criteria and the NASA Dryden Flight Research Center/Air Force Flight Test Center/AIAA Pilot Induced Oscillation Workshop, addressed this problem. An overview of these meetings is provided from the point of view of the relationship between flying qualities criteria and flight control system design. Among the items discussed are flying qualities criteria development, the role of simulation, and communication between flying qualities specialists and control system designers. 1155. Walker, H. J.: Analytic Study of Orbiter Landing Profiles. NASA TM-81365, H-1160, September 1981, 82N10001, #. A broad survey of possible orbiter landing configurations was made with specific goals of defining boundaries for the landing task. The results suggest that the center of the corridors between marginal and routine represents a more or less optimal preflare condition for regular operations. Various constraints used to define the boundaries are based largely on qualitative judgments from earlier flight experience with the X-15 and lifting body research aircraft. The results should serve as useful background for expanding and validating landing simulation programs. The analytic approach offers a particular advantage in identifying trends due to the systematic variation of factors such as vehicle weight, load factor, approach speed, and aim point. Limitations such as a constant load factor during the flare and using a fixed gear deployment time interval, can be removed by increasing the flexibility of the computer program. This analytic definition of landing profiles of the orbiter may suggest additional studies, including more configurations or more comparisons of landing profiles within and beyond the corridor boundaries. 1156. Borek, R. W.; *Pool, A.; and Sanderson, K. C.: Practical Aspects of Instrumentation System Installation, Volume 13. NASA TM-84067, AGARD-AG-160, Vol 13, September 1981, 82N13140, #. A review of factors influencing installation of aircraft flight test instrumentation is presented. Requirements, including such factors as environment, reliability, maintainability, and system safety are discussed. The assessment of the mission profile is followed by an overview of electrical and mechanical installation factors with emphasis on shock/vibration isolation systems and standardization of the electric wiring installation, two factors often overlooked by instrumentation engineers. A discussion of installation hardware reviews the performance capabilities of wiring, connectors, fuses and circuit breakers, and a guide to proper selections is provided. The discussion of the installation is primarily concerned with the electrical wire routing, shield terminations and grounding. Also included are some examples of installation mistakes that could affect system accuracy. System verification procedures and special considerations such as sneak circuits, pyrotechnics, aircraft antenna patterns, and lightning strikes are discussed. *National Aerospace Lab., Amsterdam, the Netherlands. 1157. Enevoldson, E. K.; Horton, and V. W.: “Light Bar” Attitude Indicator. Proceedings of 5th Advanced Aircrew Display Symposium, Patuxent River, Maryland, September 15–16, 1981, (see A83-16126 04-06), pp. 251–261, 83A16136, #. The development and evaluation of a light bar attitude indicator to help maintain proper aircraft attitude during high altitude night flying is described. A standard four-inch ADI was modified to project an artificial horizon across the instrument panel for pitch and roll information. A light bulb was put in the center of the ADI and a thin slit cut on the horizon, resulting in a thin horizontal sheet of light projecting from the instrument. The intensity of the projected beam is such that it can only be seen in a darkened room or at night. The beam on the instrument panel of the T-37 jet trainer is shown, depicting various attitudes. The favorable comments of about 50 pilots who evaluated the instrument are summarized, including recommendations for improving the instrument. Possible uses for the instrument to ease the pilot task are listed. Two potential problems in using the device are the development of pilot complacency and an uprightinverted ambiguity in the instrument. 1158. Thompson, M. O.; and Horton, V. W.: Exploratory Flight Test of Advanced Piloted Spacecraft—Circa 1963. Technical Review, presented at the 25th Society of Experimental Test Pilots Symposium, Beverly Hills, California, September 23–26, 1981, Vol. 16, No. 2, 1981, pp. 229–248, 82A14941, #. The NASA early experimental program for parawing and lifting body spacecraft recovery concepts is discussed. Simple hand drawings, in-house construction, and crude drop tests were used in lieu of a thorough stress analysis. The parawing (Parasev) was controlled by manually shifting the center of gravity with respect to the center of pressure; the craft would take off while being towed at 40 KIAS. The M-2 lifting body was originally constructed with a 3/32 in. 201 mahogany skin by a glider manufacturer and employed general aviation aircraft nose and main wheel assemblies. A minimum altitude of 200 ft was found acceptable for release of the parasev, allowing the pilot time to adjust for transients incurred at the tow release. A small landing assist rocket was furnished for the M-2/F-2 and X-24A lifting bodies to enhance stability, and landings at a maximum lift/drag ratio of 2.8 were successfully completed. The data gained were eventually applied in the development of the Shuttle. Parasev Vehicle E-8009 1159. Myers, A. F.; Earls, M. R.; and *Callizo, L. A.: HiMAT Onboard Flight Computer System Architecture and Qualification. AIAA Paper 81-2107. Technical Papers, presented at AIAA 3rd Computers in Aerospace Conference, San Diego, California, October 26–28, 1981, pp. 41–54, 82A10082, #. (See also 1268.) Two highly maneuverable aircraft technology (HiMAT) remotely piloted research vehicles (RPRV’s) are being flight tested at NASA Dryden Flight Research Center, Edwards, California, to demonstrate and evaluate a number of technological advances applicable to future fighter aircraft. Closed-loop primary flight control is performed from a ground-based cockpit utilizing a digital computer and up/down telemetry links. A backup flight control system for emergency operation resides in one of two onboard computers. Other functions of the onboard computer system are uplink processing, downlink processing, engine control, failure detection, and redundancy management. This paper describes the architecture, functions, and flight qualification of the HiMAT onboard flight computer systems. *Rockwell International Corp., Los Angeles, California. 1160. Anon.: Report on Research and Technology-FY 1981. NASA TM-81367, November 1981, 82N14047, #. More than 65 technical reports, papers, and articles published by personnel and contractors at the Dryden Flight Research Center are listed. Activities performed for the Offices of Aeronautics and Space Technology, Space and Terrestrial Applications, Space Transportation Systems, and Space Tracking and Data Systems are summarized. Preliminary stability and control derivatives were determined for the shuttle orbiter at hypersonic speeds from the data obtained at reentry. The shuttle tile tests, spin research vehicle nose shapes flight investigations, envelope expansion flights for the Ames tilt rotor research aircraft, and the AD-1 oblique wing programs were completed as well as the KC-135 winglet program. 1161. Sims, R. L.; and Carter, A. L.: Comparison of Wind Tunnel and Theoretical Aeroelastic Predictions With Flight Measured Airloads for the B-1 Aircraft. AIAA Paper 81-2387. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A14393, #. An aeroelastic analysis of the B-1 aircraft was generated using the FLEXSTAB computer program. Relatively simple aerodynamic and structural models were employed. Theoretical wing and horizontal stabilizer airloads were compared to wind tunnel predictions and flight data measured during quasi-steady pitch maneuvers at Mach numbers of 0.85 and 1.2 with the wing in the 67.5 degree full aft sweep position. The basic objective was to evaluate the usefulness of the FLEXSTAB program for pre-flight airloads analysis of large flexible aircraft. Significant aeroelastic increments were noted between rigid and flexible vehicle results. FLEXSTAB predicted airloads for the outer wing panel were in good agreement with measured data for both rigid airloads and elastic increments. FLEXSTAB results for the horizontal stabilizer were useful for defining general aeroelastic trends, but absolute load levels were not well predicted due to theoretical limitations and difficulties encountered in modeling the complex B-1 configuration. Overall, the FLEXSTAB program is viewed as a useful integrated tool for static aeroelastic analysis in support of flight programs. 1162. McMurtry, T. C.; Sim, A. G.; and Andrews, W. H.: AD-1 Oblique Wing Aircraft Program. AIAA Paper 81-2354. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A14390, #. A NASA program for evaluation of the handling and flying characteristics of the AD-1 oblique wing aircraft is discussed. The vehicle was flown to compare wind tunnel predictions with aerodynamic data, explore the control system requirements, and obtain a preliminary assessment of the aeroelastic effects. The fiberglass sandwich skin aircraft is designed for 8 g positive and 4 g negative loading at 202 175 knots, while the wing pivot can withstand 25 g loading. Flight monitoring was accomplished with a 41 channel pulse code modulation system for telemetry and by averaging of pilot ratings. Maneuvering tests are outlined, noting that pilot ratings indicated acceptable handling at up to 50 deg sweep. It is concluded that acceptable flying qualities can be achieved with a 60 deg sweep, and that aeroelastic tailoring can be used to satisfy cruise design technology. 1163. Bohn-Meyer, M.; and *Jiran, F.: Techniques for Modifying Airfoils and Fairings on Aircraft Using Foam and Fiberglass. AIAA Paper 81-2445. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A14383, #. The concept of using foam and fiberglass reinforced plastic to modify airfoils and fairings was applied successfully to highspeed aircraft at NASA Dryden Flight Research Center. An on-aircraft installation method was used to modify an F-15 wing glove and wing leading edge and an F-104 flap trailing edge in support of the Shuttle tile airload tests. A combination of methods, both an on-aircraft installation and an off-aircraft fabrication for installation on the aircraft, was used to modify a section of an F-111 supercritical wing with a natural laminar flow airfoil. Techniques, methods, problem areas, and recommendations are presented which indicate that using foam and fiberglass to modify airfoils and fairings on highspeed aircraft is a viable means of quickly developing airfoils and fairings with desired aerodynamic characteristics with little risk to the parent or carrier aircraft. *Fred Jiran Glider Repairs, Mojave, California. 1164. Saltzman, E. J.; and Ayers, T. G.: A Review of Flight-to-Wind Tunnel Drag Correlation. AIAA Paper 81-2475. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A14382, #. Comparisons are made of wind-tunnel-model and flight drag data for various configurations representing aircraft from the mid-1940s to the 1970s. Discrepancies between model and flight data such as Reynolds number effects, wall interference, and aeroelastic problems are discussed. Sting support effects and the inability of models to simulate surface deflections for longitudinal trim are also studied. A wind tunnel-to-flight correlation of turbulent friction drag confirms the incompressible Karman-Schoenherr variation of turbulent skin friction with Reynolds number and the T’ method for accounting compressibility effects. NASA tested 10 deg cone research indicates that model tests which are affected by tunnel noise may require the lower disturbance level environment available in flight, and it is concluded that new cryogenic facilities will improve the fidelity of model simulations of full-scale flight flow phenomena. 1165. DeAngelis, V. M.: In-Flight Deflection Measurement of the HiMAT Aeroelastically Tailored Wing. AIAA Paper 81-2450. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A14381, #. (See also 1229.) An electro-optical flight deflection measurement system was developed for NASA for use on the highly maneuverable aircraft technology (HiMAT) remotely piloted research vehicle (RPRV) to provide a means of evaluating the performance of the HiMAT’s aeroelastically tailored composite wing and canard. A description of the flight deflection measurement system is presented from a user’s viewpoint and includes the general method of operation, system capabilities and limitations, method of installation on the HiMAT vehicle, and calibration of targets. Also included is a general description of the HiMAT RPRV and its design goals. Preliminary flight deflection and bending moment data were obtained at Mach 0.8 and were extrapolated to the Mach 0.9 maneuver design condition for comparison to NASTRAN predictions and ground loads test results. The preliminary flight test results tended to agree with the results obtained from the static ground loads tests, that is, that the NASTRAN model overpredicted the streamwise twist of the composite outer wing panel. 1166. Swann, M. R.; Duke, E. L.; Enevoldson, E. K.; and Wolf, T. D.: Experience With Flight Test Trajectory Guidance. AIAA Paper 81-2504. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A14379, #. (See also 1276.) A system that provides the test pilot with flight test trajectory guidance is presently evolving at the NASA Dryden Flight Research Facility. In use, this system has resulted in discernible improvements in the ease and accuracy with which pilots have approached and maintained the desired flight test conditions or trajectories. This paper describes the use of the guidance system in several past flight programs at Dryden, including the F-111 TACT program, the F-15 airframe/propulsion system interaction program, the F-15 cone transition and boundary layer experiments, and the Space Shuttle tiles flight test program. 1167. Bever, G. A.: The Development and Use of a Computer-Interactive Data Acquisition and Display System in a Flight Environment. AIAA Paper 81-2371. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13946, #. The flight test data requirements at the NASA Dryden Flight Research Center increased in complexity, and more advanced instrumentation became necessary to accomplish mission goals. This paper describes the way in which an airborne computer was used to perform real-time calculations on 203 critical flight test parameters during a flight test on a wingletequipped KC-135A aircraft. With the computer, an airborne flight test engineer can select any sensor for airborne display in several formats, including engineering units. The computer is able to not only calculate values derived from the sensor outputs but also to interact with the data acquisition system. It can change the data cycle format and data rate, and even insert the derived values into the pulse code modulation (PCM) bit stream for recording. 1168. Meyer, R. R., Jr.; Jarvis, C. R.; and *Barneburg, J.: In-Flight Aerodynamic Load Testing of the Shuttle Thermal Protection System. AIAA Paper 81-2468. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13932, #. To contribute to the certification of the structural integrity of the Space Shuttle orbiter’s thermal protection system (TPS) before the first Shuttle flight, in-flight aerodynamic load tests of six simulated areas of the orbiter were conducted. The tests were performed on an F-104 and F-15 aircraft. This paper describes the test approach, techniques used, and results. Two areas of the orbiter TPS were redesigned and retested as a result of these tests. No TPS failures due to air-loads occurred in the areas that were evaluated in the flight tests during the Shuttle’s first flight. *NASA Johnson Space Center, Houston, Texas. 1169. Baer-Riedhart, J. L.: The Development and Flight Test Evaluation of an Integrated Propulsion Control System for the HiMAT Research Airplane. AIAA Paper 81-2467. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13931, #. The Highly Maneuverable Aircraft Technology airplane is a.44-scale version of an advanced fighter design. It is remotely piloted from a ground cockpit and is powered by a J85-GE-21 turbojet engine. The engine is electronically controlled by a digital computer onboard the airplane to operate at selected engine operation modes. The HiMAT design and development philosophy emphasized high-risk, low-cost and minimum testing, and also required that no single failure would cause loss of the vehicle. This philosophy generated unique requirements for design, computer simulation methods, specialized test techniques, and support systems which are discussed in this paper. 1170. Gera, J.; Wilson, R. J.; Enevoldson, E. K.; and *Nguyen, L. T.: Flight Test Experience With High-Alpha Control System Techniques on the F-14 Airplane. AIAA Paper 81-2505. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13906, #. Improved handling qualities of fighter aircraft at high angles of attack can be provided by various stability and control augmentation techniques. NASA and the U.S. Navy are conducting a joint flight demonstration of these techniques on an F-14 airplane. This paper reports on the flight test experience with a newly designed lateral-directional control system which suppresses such high angle of attack handling qualities problems as roll reversal, wing rock, and directional divergence while simultaneously improving departure/spin resistance. The technique of integrating a piloted simulation into the flight program was used extensively in this program. This technique had not been applied previously to high angle of attack testing and required the development of a valid model to simulate the test airplane at extremely high angles of attack. *NASA Langley Research Center, Hampton, Virginia. 1171. Ko, W. L.; Quinn, R. D.; Gong, L.; Schuster, L. S.; and Gonzales, D.: Preflight Reentry Heat Transfer Analysis of Space Shuttle. AIAA Paper 81-2382. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13882, #. Preflight predictions of the structural temperature distributions during entry are compared with data from the initial Shuttle flight. Finite element thermal analysis programming was used to model the heat flow on Shuttle structures and actual gas properties of air were employed in the analyses of aerodynamic heating. Laminar, separated, and turbulent heat fluxes were calculated for varying locations on the craft using velocity-attitude and angle-of-attack projections taken from the nominal STS-1 trajectory. Temperature time histories of the first flight are compared with laminar and turbulent flow assumptions and an unpredicted rapid cooling 1800 sec into entry is credited to inaccurate assumptions of structural heat dissipative properties or flow conditions in that time phase of the flight; additional discrepancies in descriptions of heating of the upper fuselage are attributed to a lack of knowledge of the complex flow patterns existing over that area of the Shuttle body. 1172. Iliff, K. W.; Maine, R. E.; and *Cooke, D. R.: Selected Stability and Control Derivatives From the First Space Shuttle Entry. AIAA Paper 81-2451. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13880, #. (See also 1267.) Primary stability and control derivative estimates garnered from the first Shuttle entry are reported. The craft was the first vehicle to maneuver over a wide range of hypersonic velocities, yielding data on flight characteristics from 204 previously unexplored regimes. The flight envelope was confined to entry and safe landing, with no additional maneuvers to gain control data. Data for a Mach number range of 25–1.5 and altitudes of 515,000–50,000 ft are provided, and functional ranges of the Shuttle control surfaces and attitude jets are outlined. On-board systems gathered data on aerodynamic coefficient identification, flight condition and Euler angles, and jet chamber pressures. A maximum likelihood estimation program, which contained unknown stability and control derivatives, was used for control; a control input determined the value of the unknown derivatives, and the input and spacecraft response were measured. Longitudinal and lateral directional maneuvers and their derivative estimates are described, noting wind contamination of the sideslip measurements below Mach 3. Further maneuvering and stability tests are projected for subsequent flights. *NASA Johnson Space Center, Houston, Texas. 1173. Burcham, F. W., Jr.; Myers, L. P.; Nugent, J.; Lasagna, P. L.; and Webb, L. D.: Recent Propulsion System Flight Tests at the NASA Dryden Flight Research Center. AIAA Paper 81-2438. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13874, #. The article presents a summary of the propulsion system tests conducted on a number of aircraft at the NASA Dryden Flight Research Center. The tests included digital engine control systems, engine-inlet compatibility, inlet-airframe interactions, nozzle-boattail drag and advanced turboprop acoustics. Among the aircraft evaluated were the F-15, HiMAT, F-14, and the JetStar. 1174. Matheny, N. W.; and *Panageas, G. N.: HiMAT Aerodynamic Design and Flight Test Experience. AIAA Paper 81-2433. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13871, #. Consideration is given to the design phase of the highly maneuverable aircraft technology program. Design objectives are examined, noting full-scale design and the remotely piloted research vehicle. Attention is given to subsonic, transonic, and supersonic design. Design results are discussed with reference to aerodynamic efficiency, aeroelastic tailoring, and the flight test program. *Rockwell International Corp., North American Aircraft Div., Los Angeles, California. 1175. Petersen, K. L.: Flight Experience With a Remotely Augmented Vehicle Flight Test Technique. AIAA Paper 81-2417. AIAA, SETP, SFTE, SAE, ITEA, and IEEE, 1st Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981, 82A13857, #. A flight technique which uses the remotely augmented vehicle (RAV) concept is developed to flight test advanced control law concepts. The design, development and flight test validation of a RAV system mechanized on a digital fly-bywire aircraft are described, and future applications are discussed. Flight experiments investigate complete inner loop, low sample rate, and adaptive control system mechanisms. The technique, which utilizes a ground-based FORTRAN programmable digital computer and up and down telemetry links is found to provide the flexibility necessary to effectively investigate alternate control law mechanisms in flight. 1176. Hedgley, D. R.: Solution to the Hidden-Line Problem. Astronautics and Aeronautics, Vol. 19, November 1981, 82A12803, #. It is pointed out that realistic three-dimensional renderings of solid objects or surfaces by computers have long been needed. The NASA Dryden Flight Research Center will soon publish a report and the computer program on an algorithm that solves for hidden lines. The computer program is written in FORTRAN IV and its size is approximately 35N + 9500 words, where N is the number of elements. A number of pictures are presented which were drawn by a computer using the algorithm. 1177. *Williams, D. A.; Pool, A.; and Sanderson, K. C.: AGARD Flight Test Instrumentation Series. Volume 14: The Analysis of Random Data. AGARD-AG-160-VOL-14, November 1981, 82N21099, #. *Cranfield Institute of Technology, Bedford, U. K. 1178. Smith, J. W.: Analysis of a Longitudinal PilotInduced Oscillation Experienced on the Approach and Landing Test of the Space Shuttle. NASA TM-81366, December 1981, 82N13149, #. During the final free flight (FF-5) of the shuttle’s approach and landing tests, the vehicle experienced pilot-induced oscillations near touchdown. The light test data showed that pilot inputs to the hand controller reached peak-to-peak amplitudes of 20 deg at a frequency between 3 and 3.5 radians per second. The controller inputs were sufficient to exceed the priority rate limit set in the pitch axis. A nonlinear analytical study was conducted to investigate the combined effects of pilot input, rate limiting, and time delays. The frequency response of the total system is presented parametrically as a function of the three variables. In general, with no dead time, for controller inputs of 5 deg or less, the total system behaves in a linear fashion. For 10 deg of 205 controller input, independent of the delay time, the elevon loop will be rate saturated above a frequency of 4 radians per second. 1179. Sefic, W. J.: NASA Dryden Flight Loads Research Facility. NASA TM-81368, December 1981, 82N15079, #. The Dryden Flight Loads Research Facility (NASA) and the associated equipment for simulating the loading and heating of aircraft or their components are described. Particular emphasis is placed on various fail-safe devices which are built into the equipment to minimize the possibility of damage to flight vehicles. The equipment described includes the ground vibration and moment of inertia equipment, the data acquisition system, and the instrumentation available in the facility for measuring load, position, strain, temperature, and acceleration. 1180. Kelley, W. W.; and Enevoldson, E. K.: Limited Evaluation of an F-14A Airplane Utilizing an AileronRudder Interconnect Control System in the Landing Configuration. NASA TM-81972, December 1981, 82N13148, #. A flight test was conducted for preliminary evaluation of an aileron-rudder interconnect (ARI) control system for the F-14A airplane in the landing configuration. Two ARI configurations were tested in addition to the standard F-14 flight control system. Results of the flight test showed marked improvement in handling qualities when the ARI systems were used. Sideslip due to adverse yaw was considerably reduced, and airplane turn rate was more responsive to pilot lateral control inputs. Pilot comments substantiated the flight data and indicated that the ARI systems were superior to the standard control system in terms of pilot capability to make lateral offset corrections and heading changes on final approach. 1181. *Kalev, I.: Cyclic Plasticity Models and Application in Fatigue Analysis. Presented at 3rd Nonlinear Finite Element Analysis and ADINA Conference, Cambridge, Massachusetts, June 10–12, 1981. Computers and Structures, Vol. 13, October–December 1981, pp. 709–716, 81A38340. An analytical procedure for prediction of the cyclic plasticity effects on both the structural fatigue life to crack initiation and the rate of crack growth is presented. The crack initiation criterion is based on the Coffin-Manson formulae extended for multiaxial stress state and for inclusion of the mean stress effect. This criterion is also applied for the accumulated damage ahead of the existing crack tip which is assumed to be related to the crack growth rate. Three cyclic plasticity models, based on the concept of combination of several yield surfaces, are employed for computing the crack growth rate of a crack plane stress panel under several cyclic loading conditions. *National Research Council, Washington, D. C. 1182. Lasagna, P.; and Mackall, K.: Acoustic Flight Testing of Advanced Design Propellers on a JetStar Aircraft. NASA CP-2208, NASA Langley Research Center Advan. Aerodyn.: Selected NASA Res., December 1981, (see N84-27660 18-01), pp. 1–10, 84N27661, #. Advanced turboprop-powered aircraft have the potential to reduce fuel consumption by 15 to 30 percent as compared with an equivalent technology turbofan-powered aircraft. An important obstacle to the use of advanced design propellers is the cabin noise generated at Mach numbers up to .8 and at altitudes up to 35,000 feet. As part of the NASA Aircraft Energy Efficiency Program, the near-field acoustic characteristics on a series of advanced design propellers are investigated. Currently, Dryden Flight Research Center is flight testing a series of propellers on a JetStar airplane. The propellers used in the flight test were previously tested in wind tunnels at the Lewis Research Center. Data are presented showing the narrow band spectra, acoustic wave form, and acoustic contours on the fuselage surface. Additional flights with the SR-3 propeller and other advanced propellers are planned in the future. Improvements in cruise efficiency on the order of 15 to 40% are obtained by increasing the extent of laminar flow over lifting surfaces. Two methods of achieving laminar flow are being considered, natural laminar flow and laminar flow control. Natural laminar flow (NLF) relies primarily on airfoil shape while laminar flow control involves boundary layer suction or blowing with mechanical devices. The extent of natural laminar flow that could be achieved with consistency in a real flight environment at chord Reynolds numbers in the range of 30 × 10(6) power was evaluated. Nineteen flights were conducted on the F-111 TACT airplane having a NLF airfoil glove section. The section consists of a supercritical airfoil providing favorable pressure gradients over extensive portions of the upper and lower surfaces of the wing. Boundary layer measurements were obtained over a range of wing leading edge sweep angles at Mach numbers from 0.80 to 0.85. Data were obtained for natural transition and for a range of forced transition locations over the test airfoil. 1184. Maine, Richard E.; and Iliff, Kenneth W.: Formulation and Implementation of a Practical Algorithm for Parameter Estimation With Process and Measurement Noise. SIAM Journal on Applied Mathematics, Vol. 41, No. 3, December 1981, (See also 1109, 1209.) A new formulation is proposed for the problem of parameter estimation of dynamic systems with both process and measurement noise. The formulation gives estimates that are maximum likelihood asymptotically in time. The means used to overcome the difficulties encountered by previous formulations are discussed. It is then shown how the proposed formulation can be efficiently implemented in a computer program. A computer program using the proposed formulation is available in a form suitable for routine application. Examples with simulated and real data are given to illustrate that the program works well. 1982 Technical Publications 1185. *Kelly, G. L.; *Berthold, G.; and Abbott, L.: Implementation of the DAST ARW II Control Laws Using an 8086 Microprocessor and an 8087 FloatingPoint Coprocessor. Mini and Microcomputers in Control and Measurement, Proceedings of the International Symposium, San Francisco, California, Acta Press, Anaheim, California and Calgary, Alberta, Canada, 1982, pp. 58–60, 83A11910. A 5 MHZ single-board microprocessor system which incorporates an 8086 CPU and an 8087 Numeric Data Processor is used to implement the control laws for the NASA Drones for Aerodynamic and Structural Testing, Aeroelastic Research Wing II. The control laws program was 206 ECN-15662 C-140 JetStar Propeller Testbed 1183. Montoya, L. C.; Steers, L. L.; and Trujillo, B.: F-111 TACT Natural Laminar Flow Glove Flight Results. NASA CP-2208, NASA Langley Research Center Advan. Aerodyn.: Selected NASA Res., December 1981, (see N84-27660 18-01), pp. 11–20, 84N27662, #. executed in 7.02 msec, with initialization consuming 2.65 msec and the control law loop 4.38 msec. The software emulator execution times for these two tasks were 36.67 and 61.18, respectively, for a total of 97.68 msec. The space, weight and cost reductions achieved in the present, aircraft control application of this combination of a 16-bit microprocessor with an 80-bit floating point coprocessor may be obtainable in other real time control applications. *University of Kansas, Lawrence, Kansas. 980025 KC-135 Airplane, Three-View Drawing 1188. Montoya, L. C.; *Jacobs, P.; *Flechner, S.; and Sims, R.: KC-135 Wing and Winglet Flight Pressure Distributions, Loads, and Wing Deflection Results With Some Wind Tunnel Comparisons. NASA CP-2211, KC-135 Winglet Program Rev., (see N84-27686 18-02), January 1982, pp. 47–102, 84N27688, #. A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/–4 deg, 15/–2 deg, and 0/–4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/–2 deg incidence winglet configuration. *NASA Langley Research Center, Hampton, Virginia. 207 EC80-14090 Firebee DAST RPV 1186. Anon.: KC-135 Winglet Program Review. NASA CP-2211, H-1165, NAS 1.55:2211. Symposium held at Edwards, California, September 16, 1981, January 1982, 84N27686, #. 1187. Barber, M. R.; and *Selegan, D.: KC-135 Winglet Program Overview. NASA CP-2211, KC-135 Winglet Program Rev., (see N84-27686 18-02), January 1982, pp. 1–46, 84N27687, #. A joint NASA/USAF program was conducted to accomplish the following objectives: (1) evaluate the benefits that could be achieved from the application of winglets to KC-135 aircraft; and (2) determine the ability of wind tunnel tests and analytical analysis to predict winglet characteristics. The program included wind-tunnel development of a test winglet configuration; analytical predictions of the changes to the aircraft resulting from the application of the test winglet; and finally, flight tests of the developed configuration. Pressure distribution, loads, stability and control, buffet, fuel mileage, and flutter data were obtained to fulfill the objectives of the program. *AFWAL, Wright-Patterson AFB, Ohio. 1189. Lux, D. P.: In-Flight Lift and Drag Measurements on a First Generation Jet Transport Equipped With Winglets. NASA CP-2211, KC-135 Winglet Program Rev., (see N84-27686 18-02), January 1982, pp. 103–116, 84N27689, #. A KC-135A aircraft equipped with wing tip winglets was flight tested to demonstrate and validate the potential performance gain of the winglet concept as predicted from analytical and wind tunnel data. Flight data were obtained at cruise conditions for Mach numbers of 0.70, 0.75, and 0.80 at a nominal altitude of 36,000 ft. and winglet configurations of 15 deg cant/–4 deg incidence, 0 deg cant/–4 deg incidence, and baseline. For the Mach numbers tested the data show that the addition of winglets did not affect the lifting characteristics of the wing. However, both winglet configurations showed a drag reduction over the baseline configuration, with the best winglet configuration being the 15 deg cant/–4 deg incidence configuration. This drag reduction due to winglets also increased with increasing lift coefficient. It was also shown that a small difference exists between the 15 deg cant/–4 deg incidence flight and wind tunnel predicted data. This difference was attributed to the pillowing of the winglet skins in flight which would decrease the winglet performance. 1190. Kehoe, M. W.: KC-135A Winglet Flight Flutter Program. NASA CP-2211, H-1169, KC-135 Winglet Program Rev., (see N84-27686 18-02), January 1982, pp. 171–188, 84N27692, #. The evaluation techniques, results and conclusions for the flight flutter testing conducted on a KC-135A airplane configured with and without winglets are discussed. Test results are presented for the critical symmetric and antisymmetric modes for a fuel distribution that consisted of 10,000 pounds in each wing main tank and empty reserve tanks. The results indicated that a lightly damped oscillation was experienced for a winglet configuration of a 0 deg cant and –4 deg incidence. The effects of cant and incidence angle variation on the critical modes are also discussed. Lightly damped oscillations were not encountered for any other winglet cant and incidence angles tested. 1191. *Dougherty, N. S., Jr.; and Fisher, D. F.: BoundaryLayer Transition Correlation on a Slender Cone in Wind Tunnels and Flight for Indications of Flow Quality. NASA TM-84732, NAS 1.15:84732, AD-A111328, AEDC TR-81-26, Arnold AFB, Tennessee, February 1982, 82N25228, #. Boundary layer transition location measurements were made on a 10 deg sharp cone in 23 wind tunnels in the United States and Europe and in flight. The data were acquired at subsonic, transonic, and supersonic Mach numbers over a range of unit Reynolds numbers in an effort to obtain an improved understanding of the effect of wind tunnel flow quality on transition location. The data indicate that the transition 208 mechanism in both wind tunnels in flight is associated with the formation of Tollmien-Schlichting waves in the laminar boundary layer. However, the location of the end of transition was found to be primarily a function of the noise under the laminar boundary of the cone surface and, within + or –20 percent, independent of Mach number and unit Reynolds number. *ARO, Inc., Tullahoma, Tennessee. 1192. Hedgley, D. R., Jr.: User’s Guide for SKETCH. NASA TM-81369, H-1169, February 1982, 82N17878, #. A user’s guide for the computer program SKETCH is presented. The removal of hidden lines from images of solid objects is a problem in computer graphics which is solved by SKETCH. 1193. Hedgley, D. R., Jr.: A General Solution to the Hidden-Line Problem. NASA RP-1085, H-1162, NAS 1.61:1085, March 1982, 82N21907, #. The requirements for computer-generated perspective projections of three dimensional objects has escalated. A general solution was developed. The theoretical solution to this problem is presented. The method is very efficient as it minimizes the selection of points and comparison of line segments and hence avoids the devastation of square-law growth. 1194. *Elfstrom, G. M.; *Kostopoulos, C.; **Peake, D. J.; and Fisher, D. F.: The Obstacle Block as a Device to Measure Turbulent Skin Friction in Compressible Flow. AIAA Paper 82-0589. Presented at 12th Aerodynamic Testing Conference, Williamsburg, Virginia, (see A82-24651 10-09), March 22–24, 1982, pp. 131–138, 82A24664, #. The obstacle block, developed as an alternative to the Preston tube for indirectly measuring skin friction on smooth surfaces in incompressible flows, is examined as a device for compressible flows as well. The block, which is congruent with a surface static pressure orifice, has a geometry which is easily specified and thus has a universal calibration. Data from two independent studies are used to establish such a calibration using “wall” variables, valid for Mach numbers up to about 3. Various aspects concerning practical application of the device are examined, such as sensitivity to yaw and the minimum permissible axial spacing between blocks. Several examples showing the utility of the device are given. *National Aeronautical Establishment, High Speed Aero Laboratory, Ottawa, Canada. **NASA Ames Research Center, Moffett Field, California. 1195. Smith, Harriet J.; and Enevoldson, Einar: The Application of a Six-Degree-of-Freedom Piloted Simulation in Support of the F-14 Flight Research Program. Presented at the SES-SFTF Symposium on Simulation Aircraft Test and Evaluation, Naval Air Test Center, Maryland, March 16–17, 1982. 1196. Ko, W. L.: Structural Properties of Superplastically Formed/Diffusion-Bonded Orthogonally Corrugated Core Sandwich Plates. AIAA Journal, Vol. 20, No. 4, April 1982, (see AIAA 80-0304), pp. 536–543, 80A18305, #. (See also 1080.) A new orthogonally corrugated sandwich structure that can be fabricated by using the superplastic forming-diffusion bonding (SPF-DB) process is described. Formulas and the associated plots for evaluating the effective elastic constants and the bending stiffness for the core of this new sandwich structure are presented. The structural properties of this sandwich structure are compared with the conventional honeycomb core sandwich structure under the conditions of equal sandwich density. The SPF-DB orthogonally corrugated sandwich core has higher traverse shear stiffness than the conventional honeycomb sandwich core, but has lower stiffness in the sandwich core-thickness direction. 1197. Berry, D. T.; Powers, B. G.; Szalai, K. J.; and Wilson, R. J.: In-Flight Evaluation of Control System Pure Time Delays. J. of Aircraft, Vol. 19, No. 4, April 1982 (see AIAA 80-1626), pp. 318–323. (See also 1108.) 1198. Glover, R. D.: Aircraft Interrogation and Display System: A Ground Support Equipment for Digital Flight Systems. NASA TM-81370, NAS 1.15:81370, April 1982, 82N21175, #. A microprocessor-based general purpose ground support equipment for electronic systems was developed. The hardware and software are designed to permit diverse applications in support of aircraft flight systems and simulation facilities. The implementation of the hardware, the structure of the software, describes the application of the system to an ongoing research aircraft project are described. 1199. Berry, D. T.: Flying Qualities—A Costly Lapse in Flight-Control Design. Astronautics and Aeronautics, Vol. 20, April 1982, pp. 54–57, 82A28280, #. Generic problems in advanced aircraft with advanced control systems which suffer from control sensitivity, sluggish response, and pilot-induced oscillation tendencies are examined, with a view to improving techniques for eliminating the problems in the design phase. Results of two NASA and NASA/AIAA workshops reached a consensus that flying qualities criteria do not match control system development, control system designers are not relying on past experience in their field, ground-based simulation is relied on too heavily, and communications between flying qualities and control systems engineers need improvement. A summation is offered in that hardware and software have 209 outstripped the pilot’s capacity to use the capabilities which new aircraft offer. The flying qualities data base is stressed to be dynamic, and continually redefining the man/machine relationships. 1200. Dittmar, J. H.; and Lasagna, P. L.: A Preliminary Comparison Between the SR-3 Propeller Noise in Flight and in a Wind Tunnel. NASA TM-82805, NAS 1.15:82805. Presented at 103rd Meeting of the Acoust. Soc. of Am., Chicago, Illinois, April 27–30, 1982, 82N21998, #. The noise generated by supersonic-tip-speed propellers is addressed. Models of such propellers were tested for acoustics in the Lewis 8-by-6-foot wind tunnel. One of these propeller models, SR-3, was tested in flight on the JetStar airplane and noise data were obtained. Preliminary comparisons of the maximum blade passing tone variation with helical tip Mach number taken in flight with those taken in the tunnel showed good agreement when corrected to the same test conditions. This indicated that the wind tunnel is a viable location for measuring the noise of these propeller models. Comparisons of the directivities at 0.6 and 0.7 axial Mach number showed reasonable agreement. At 0.75 and 0.8 axial Mach number the tunnel directivity data fell off more towards the front than did the airplane data. A possible explanation for this is boundary layer refraction which could be different in the wind tunnel from that in flight. This may imply that some corrections should be applied to both the airplane and wind tunnel data at the forward angles. At and aft of the peak noise angle the boundary layer refraction does not appear to be significant and no correction appears necessary. 1201. Barber, M. R.; and *Tymczyszyn, J. J.: Wake Vortex Attenuation Flight Tests—A Status Report. Presented at the 13th Annual Symposium of the Society of Experimental Test Pilots, Rottach-Egern, West Germany. Published in Cockpit, Vol. 17, April–June 1982, pp. 6–26, 83A11806. (See also 1135.) *FAA, Los Angeles, California. 1202. *Shideler, J. L.; **Swegle, A. R.; and Fields, R. A.: Development of René 41 Honeycomb Structure as an Integral Cryogenic Tankage/Fuselage Concept for Future Space Transportation Systems. AIAA Paper 82-0653. Presented at AIAA, ASME, ASCE, and AHS 23rd Structures, Structural Dynamics and Materials Conference, New Orleans, Louisiana, May 10–12, 1982, pp. 66–75, 82A30084, #. (See also 1203, 1342.) The status of the structural development of an integral cryogenic-tankage/hot-fuselage concept for future space transportation systems is reviewed. The concept comprises a honeycomb sandwich structure that serves the combined functions of containing the cryogenic fuel, supporting the vehicle loads, and protecting the spacecraft from entry heating. The inner face sheet is exposed to cryogenic temperature of –423 F during boost; the outer face sheet, which is slotted to reduce thermal stress, is exposed to a maximum temperature of 1400 F during a high-altitude gliding entry. Attention is given to the development of a fabrication process for a Rene 41 honeycomb sandwich panel with a core density of less than 1 percent that is consistent with desirable heat treatment processes for high strength. *NASA Langley Research Center, Hampton, Virginia. **Boeing Aerospace Co., Seattle, Washington. 1203. *Shideler, J. L.; *Swegle, A. R.; and Fields, R. A.: Development of René 41 Honeycomb Structure as an Integral Cryogenic Tankage/Fuselage Concept for Future Space Transportation Systems. NASA TM-83306, NAS 1.15:83306, Presented at the 23rd AIAA, ASME, ASCE, and AHS Structural, Structural Dynamics and Materials Conference, New Orleans, Louisiana, May 10–12, 1982, June 1982, 82N30328, #. (See also 1202, 1342.) The status of the structural development of an integral cryogenic-tankage/hot-fuselage concept for future space transportation systems (STS) is discussed. The concept consists of a honeycomb sandwich structure which serves the combined functions of containment of cryogenic fuel, support of vehicle loads, and thermal protection from an entry heating environment. The inner face sheet is exposed to a cryogenic (LH2) temperature of –423 F during boost; and the outer face sheet, which is slotted to reduce thermal stress, is exposed to a maximum temperature of 1400 F during a high altitude, gliding entry. A fabrication process for a Rene’ 41 honeycomb sandwich panel with a core density less than 1 percent was developed which is consistent with desirable heat treatment processes for high strength. *NASA Langley Research Center, Hampton, Virginia. **Boeing Aerospace Co., Seattle, Washington. 1204. Fisher, D. F.; and *Dougherty, N. S., Jr.: In-Flight Transition Measurement on a 10 Deg Cone at Mach Numbers From 0.5 to 2.0. NASA TP-1971, H-1117, NAS 1.60:1971, June 1982, 82N26227, #. Boundary layer transition measurements were made in flight on a 10 deg transition cone tested previously in 23 wind tunnels. The cone was mounted on the nose of an F-15 aircraft and flown at Mach numbers room 0.5 to 2.0 and altitudes from 1500 meters (5000 feet) to 15,000 meters (50,000 feet), overlapping the Mach number/Reynolds number envelope of the wind tunnel tests. Transition was detected using a traversing pitot probe in contact with the surface. Data were obtained near zero cone incidence and adiabatic wall temperature. Transition Reynolds number was found to be a function of Mach number and of the ratio of wall temperature to adiabatic all temperature. Microphones mounted flush with the cone surface measured free-stream disturbances imposed on the laminar boundary layer and 210 identified Tollmien-Schlichting waves as the probable cause of transition. Transition Reynolds number also correlated with the disturbance levels as measured by the cone surface microphones under a laminar boundary layer as well as the free-stream impact. *ARO, Inc., Tullahoma, Tennessee. 1205. Baer-Riedhart, J. L.: Evaluation of a Simplified Gross Thrust Calculation Method for a J85-21 Afterburning Turbojet Engine in an Altitude Facility. AIAA Paper 82-1044. Presented at 18th AIAA, SAE and ASME Joint Propulsion Conference, Cleveland, Ohio, June 21–23, 1982, 82A34978, #. A simplified gross thrust calculation method was evaluated on its ability to predict the gross thrust of a modified J85-21 engine. The method used tailpipe pressure data and ambient pressure data to predict the gross thrust. The method’s algorithm is based on a one-dimensional analysis of the flow in the afterburner and nozzle. The test results showed that the method was notably accurate over the engine operating envelope using the altitude facility measured thrust for comparison. A summary of these results, the simplified gross thrust method and requirements, and the test techniques used are discussed in this paper. 1206. Myers, L. P.; Mackall, K. G.; Burcham, F. W., Jr.; and *Walter, W. A.: Flight Evaluation of a Digital Electronic Engine Control System in an F-15 Airplane. AIAA Paper 82-1080. Presented at 18th AIAA, SAE and ASME Joint Propulsion Conference, Cleveland, Ohio, June 21–23, 1982, 82A37683, #. Benefits provided by a full-authority digital engine control are related to improvements in engine efficiency, performance, and operations. An additional benefit is the capability of detecting and accommodating failures in real time and providing engine-health diagnostics. The digital electronic engine control (DEEC), is a full-authority digital engine control developed for the F100-PW-100 turbofan engine. The DEEC has been flight tested on an F-15 aircraft. The flight tests had the objective to evaluate the DEEC hardware and software over the F-15 flight envelope. A description is presented of the results of the flight tests, which consisted of nonaugmented and augmented throttle transients, airstarts, and backup control operations. The aircraft, engine, DEEC system, and data acquisition and reduction system are discussed. *United Technologies Corp., Pratt and Whitney Aircraft Group, West Palm Beach, Florida. 1207. Mackall, K. G.; Lasagna, P. L.; *Dittmar, J. H.; and Walsh, K.: In-Flight Acoustic Results From an AdvancedDesign Propeller at Mach Numbers to 0.8. AIAA Paper 82-1120. Presented at 18th AIAA, SAE and ASME Joint Propulsion Conference, Cleveland, Ohio, June 21–23, 1982, 82A35017, #. Acoustic data for the advanced-design SR-3 propeller at Mach numbers to 0.8 and helical tip Mach numbers to 1.14 are presented. Several advanced-design propellers, previously tested in wind tunnels at the Lewis Research Center, are being tested in flight at the Dryden Flight Research Facility. The flight-test propellers are mounted on a pylon on the top of the fuselage of a JetStar airplane. Instrumentation provides near-field acoustic data for the SR-3. Acoustic data for the SR-3 propeller at Mach numbers up to 0.8, for propeller helical tip Mach numbers up to 1.14, and comparison of wind tunnel to flight data are included. Flowfield profiles measured in the area adjacent to the propeller are also included. *NASA Lewis Research Center, Cleveland, Ohio. 1208. Webb, L. D.; and Nugent, J.: Selected Results of the F-15 Propulsion Interactions Program. AIAA Paper 82-1041. Presented at 18th AIAA, SAE and ASME Joint Propulsion Conference, Cleveland, Ohio, June 21–23, 1982, 82A34976, #. A better understanding of propulsion system/airframe flow interactions could aid in the reduction of aircraft drag. For this purpose, NASA and the United States Air Force have conducted a series of wind-tunnel and flight tests on the F-15 airplane. This paper presents a correlation of flight test data from tests conducted at the NASA Dryden Flight Research Facility of the Ames Research Center, with data obtained from wind-tunnel tests. Flights were made at stabilized Mach numbers around 0.6, 0.9, 1.2, and 1.5 with accelerations up to near Mach number 2. Wind-tunnel tests used a 7.5 percentscale F-15 inlet/airframe model. Flight and wind-tunnel pressure coefficients showed good agreement in most cases. Correlation of interaction effects caused by changes in cowl angle, angle-of-attack, and Mach number are presented. For the afterbody region, the pressure coefficients on the nozzle surfaces were influenced by boattail angles and Mach number. Boundary-layer thickness decreased as angle of attack increased above 4 deg. 1209. Maine, R. E.; and Iliff, K. W.: Formulation of a Practical Algorithm for Parameter Estimation With Process and Measurement Noise. Identification and System Parameter Estimation 1982. Proceedings of the Sixth IFAC Symposium, Washington, DC, Vol. 2, June 7–11, 1982, pp. 1139–1144, 84A18611. (See also 1109, 1184.) A new formulation is proposed for the problem of parameter estimation in dynamic systems with both process and measurement noise. The formulation applies to continuoustime state space system models with discrete-time measurements. Previous formulations of this problem encountered several theoretical and practical difficulties which are overcome by the new formulation. The most 211 important element of the new formulation is a reparameterization of the unknown noise covariances. A computer program that implements the new formulation is available. 1210. Meyer, R. R., Jr.: A Unique Flight Test Facility — Description and Results. Proceedings, 13th International Council of the Aeronautical Sciences Congress, and AIAA Aircraft Systems and Technology Conference, Seattle, Washington, Vol. 1, August 22–27, 1982, pp. 433–448, 82A40925, #. (See also 1223.) The Dryden Flight Research Facility has developed a unique research facility for conducting aerodynamic and fluid mechanics experiments in flight. A low aspect ratio fin, referred to as the flight test fixture (FTF), is mounted on the underside of the fuselage of an F-104G aircraft. The F-104/FTF facility is described, and the capabilities are discussed. The capabilities include (1) a large Mach number envelope (0.4 to 2.0), including the region through Mach 1.0; (2) the potential ability to test articles larger than those that can be tested in wind tunnels; (3) the large chord Reynolds number envelope (greater than 40 million); and (4) the ability to define small increments in friction drag between two test surfaces. Data are presented from experiments that demonstrate some of the capabilities of the FTF, including the shuttle thermal protection system airload tests, instrument development, and base drag studies. Proposed skin friction experiments and instrument evaluation studies are also discussed. 1211. Kempel, R. W.: Flight Experience With a Backup Flight-Control System for the HiMAT Research Vehicle. AIAA Paper 82-1541. Presented at AIAA Guidance and Control Conference, San Diego, California, August 9–11, 1982, 82A40429, #. The NASA Dryden Flight Research Facility is conducting flight tests of two remotely piloted, subscale, advanced fighter configurations; the tests are part of the Highly Maneuverable Aircraft Technology (HiMAT) project. Closed-loop primary flight control is performed from a ground-based cockpit and digital computer in conjunction with an up/down telemetry link. A significant feature of these vehicles is an on-board, digitally active, backup control system designed to recover the vehicle in the event of a transfer from primary control. Automatic transfers occur upon certain critical ground or airborne system malfunctions. Control modes are provided that enable a ground or airborne controller to guide the vehicle to a safe landing. This paper describes the features, operational development, and flight evaluation of the HiMAT backup flight control system. 1212. Shafer, M. F.: Flight-Determined Correction Terms for Angle of Attack and Sideslip. AIAA Paper 82-1374. Presented in 9th AIAA Atmospheric Flight Mechanics Conference, San Diego, California, August 9–11, 1982, 82A40290, #. The effects of local flow, upwash, and sidewash on angle of attack and sideslip (measured with boom-mounted vanes) were determined for subsonic, transonic, and supersonic flight using a maximum likelihood estimator. The correction terms accounting for these effects were determined using a series of maneuvers flown at a large number of flight conditions in both augmented and unaugmented control modes. The correction terms provide improved angle-ofattack and sideslip values for use in the estimation of stability and control derivatives. In addition to detailing the procedure used to determine these correction terms, this paper discusses various effects, such as those related to Mach number, on the correction terms. The use of maneuvers flown in augmented and unaugmented control modes is also discussed. 1213. *Gupta, N. K.; and Iliff, K. W.: Identification of Aerodynamic Indicial Functions Using Flight Data. AIAA Paper 82-1375. Presented in 9th AIAA Atmospheric Flight Mechanics Conference, San Diego, California, August 9–11, 1982, 82A39136, #. It is pointed out that the use of indicial function representation provides a model superior to the aerodynamic derivative model. Specific derivatives can be approximated from the indicial models. The model can also be used to compute equivalent stability and control parameters not usually available from flight data. It is shown that derivatives regarding the angle-of-attack and the side slip angle can be derived directly from the indicial functions without any identifiability problem. Attention is given to the pitch moment coefficient, linear indicial function representation, the identification problem for the pitch moment equation, the identifiability of linear systems, parametric representations of the indicial functions, an identification technique, angle-ofattack and pitch rate dynamics in the pitch plane, multivariate linear models, nonlinear aerodynamic indicial functions, measurement system accuracy, and poststall and spin-entry data from a scaled research vehicle. *Integrated Systems, Inc., Palo Alto, California. 1214. Iliff, K. W.; and Maine, R. E.: NASA Dryden’s Experience in Parameter Estimation and Its Uses in Flight Test. AIAA Paper 82-1373. Presented in 9th AIAA Atmospheric Flight Mechanics Conference, San Diego, California, August 9–11, 1982, 82A39135, #. (See also 1294.) An explanation of the parameter estimation method used at the Dryden Flight Research Facility is presented, and an overview is provided of experience related to the employment of this method, taking into account the utilization of this experience in flight tests. According to a definition of the aircraft parameter estimation problem, the system investigated is assumed to be modeled by a set of dynamic equations containing unknown parameters. To 212 determine the values of the unknown parameters, the system is excited by a suitable input, and the input and actual system response are measured. The values of the unknown parameters are then inferred, based on the requirement that the model response to the given input match the actual system response. Examples of parameter estimation in flight test are discussed, giving attention to the F-14 fighter, the HiMAT (high maneuverable aircraft technology) vehicle, and the Space Shuttle. 1215. Curry, R. E.; and Sim, A. G.: Unique Flight Characteristics of the AD-1 Oblique-Wing Research Airplane. AIAA Paper 82-1329. Presented at 9th AIAA Atmospheric Flight Mechanics Conference, San Diego, California, August 9–11, 1982, 82A39106, #. (See also 1257.) Flight characteristics associated with an oblique-wing airplane have been studied with limited scope and complexity using the AD-1 research vehicle. The AD-1 is a low-speed, low-cost, manned airplane with an aeroelastically tailored wing that can be pivoted 0 to 60 deg asymmetrically. Results of the flight tests include aerodynamic parameter extraction, verification of the aeroelastic wing design criteria, trim requirements, stall characteristics, and an evaluation of the handling qualities and basic control system requirements. Some of the unique characteristics of these results that pertain to the oblique-wing design are presented. 1216. Maine, R. E.; and Iliff, K. W.: Selected Stability and Control Derivatives From the First Three Space Shuttle Entries. AIAA Paper 82-1318. Presented at the 9th AIAA Atmospheric Flight Mechanics Conference, San Diego, California, August 9–11, 1982, 82A39096, #. Stability and control derivative estimates obtained from the first three Space Shuttle entries are presented. The derivative estimates were obtained using the established modified maximum likelihood estimation program (MMLE3). The method of analysis used by the MMLE3 program is reviewed, the Shuttle configuration and data system are described, and the various types of maneuvers analyzed from the three entries are illustrated. Finally, the paper presents selected derivative results and compares them with predictions. Most of the flight-derived estimates agreed fairly well with predictions, considering the lack of experience in these new flight regimes. The most notable exception was the aerodynamic interference caused by firing the reaction control jets in the atmosphere. The flight results showed this interference to be considerably smaller than predicted. 1217. *Bailey, R. E.; *Smith, R. E.; and Shafer, M. F.: An In-Flight Investigation of Pilot-Induced Oscillation Suppression Filters During the Fighter Approach and Landing Task. Flight Testing Technology: A State-of-theArt Review; Proceedings of the Thirteenth Annual SETP Symposium, New York, New York, September 19–22, 1982, pp. 185–191, A84-44451. An investigation of pilot-induced oscillation suppression (PIOS) filters was performed using the USAF/Flight Dynamics Laboratory variable stability NT-33 aircraft, modified and operated by Calspan. This program examined the effects of PIOS filtering on the longitudinal flying qualities of fighter aircraft during the visual approach and landing task. Forty evaluations were flown to test the effects of different PIOS filters. Although detailed analyses were not undertaken, the results indicate that PIOS filtering can improve the flying qualities of an otherwise unacceptable aircraft configuration (Level 3 flying qualities). However, the ability of the filters to suppress pilot-induced oscillations appears to be dependent upon the aircraft configuration characteristics. Further, the data show that the filters can adversely affect landing flying qualities if improperly designed. The data provide an excellent foundation from which detail analyses can be performed. *Calspan Advanced Technology Center, Buffalo, New York. 1218. Duke, E. L.: Automated Flight Test Maneuvers— The Development of a New Technique. Flight Testing Technology: A State-of-the-Art Review; Proceedings of the Thirteenth Annual SETP Symposium, New York, New York, September 19–22, 1982, pp. 101–119, 84A44464, #. A new flight test technique using a maneuver autopilot is being applied at the Dryden Flight Research Facility of the NASA Ames Research Center. The flight test maneuver autopilot (FTMAP) is designed to provide precise, repeatable control of an aircraft during certain prescribed maneuvers so that a large quantity of high quality test data can be obtained with a minimum of flight time. This paper discusses the control algorithms, the flight test application, and the preliminary flight demonstration results of the FTMAP. 1219. Harney, P. F.: Real-Time Data Display for AFTI/F-16 Flight Testing. ITC/USA/’82; Proceedings of the International Telemetering Conference, San Diego, California, September 28–30, 1982, pp. 13–33, (see A84-32401 14-32), 84A32403. (See also 1225.) Advanced fighter technologies to improve air to air and air to surface weapon delivery and survivability is demonstrated. Real time monitoring of aircraft operation during flight testing is necessary not only for safety considerations but also for preliminary evaluation of flight test results. The complexity of the AFTI/F-16 aircraft requires an extensive capability to accomplish real time data goals; that capability and the resultant product are described. Previously announced in STAR as N83-13095. AFTI F-16 Airplane EC92-10061-10 1220. Shafer, M. F.: Low Order Equivalent Models of Highly Augmented Aircraft Determined From Flight Data Using Maximum Likelihood Estimation. Journal of Guidance, Control and Dynamics, Vol. 5, No. 5, September–October 1982, (see AIAA Paper 80-1627), pp. 504–511. (See also 1107.) This paper presents the results of a study of the feasibility of using low order equivalent mathematical models of a highly augmented aircraft, the F-8 digital fly-by-wire (DFBW), for flying qualities research. Increasingly complex models were formulated and evaluated using flight data and maximum likelihood estimation techniques. The airframe was first modeled alone. Next, equivalent derivatives were used to model the longitudinal unaugmented F-8 DFBW aircraft dynamics. The most complex model of the unaugmented aircraft incorporated a pure time shift of the pilot input, a first order lag, and the basic longitudinal airframe model. This same model was then implemented for the F-8 DFBW aircraft in a highly augmented mode. Excellent matching of the dynamics resulted for this model, indicating that low order equivalent models that are good representations of the highly augmented F-8 DFBW aircraft can be formulated with these methods. 1221. Murray, J. E.: User’s Manual for THPLOT, A FORTRAN 77 Computer Program for Time History Plotting. NASA TM-81374, NAS 1.15:81374, October 1982, 83N12881, #. A general purpose FORTRAN 77 computer program (THPLOT) for plotting time histories using Calcomp pen plotters is described. The program is designed to read a time history data file and to generate time history plots for selected time intervals and/or selected data channels. The capabilities of the program are described. The card input required to 213 define the plotting operation is described and examples of card input and the resulting plotted output are given. The examples are followed by a description of the printed output, including both normal output and error messages. Lastly, implementation of the program is described. A complete listing of the program with reference maps produced by the CDC FTN 5.0 compiler is included. 1222. *Peake, D. J.; Fisher, D. F.; and **McRae, D. S.: Flight, Wind Tunnel and Numerical Experiments With a Slender Cone at Incidence. AIAA Journal, Vol. 20, No.10, October 1982. The three-dimensional leeward separation about a 5 deg semi-angle cone at an 11 deg incidence was investigated in flight, in the wind tunnel, and by numerical computations. The test conditions were Mach numbers of 0.6, 1.5, and 1.8 at Reynolds numbers between 7 and 10 million based on freestream conditions and a 76.2 cm (30-in.) length of surface. The surface conditions measured included those of fluctuating pressures and mean static, as well as recovery pressures generated by obstacle blocks to provide skin friction and separation line positions. The mean static pressures from flight and wind tunnel were in reasonably good agreement. The computed results gave the same distributions, but were slightly more positive in magnitude. The experimentally measured primary and secondary separation line locations compared closely with computed results. There were substantial differences in level between the surface root-mean-square pressure fluctuations obtained in flight and in the wind tunnel, due, it is thought, to a relatively high acoustic disturbance level in the tunnel compared with the quiescent conditions in flight. *NASA Ames Research Center, Moffett Field, California. **North Carolina State University, Raleigh, North Carolina. 1223. Meyer, R. R., Jr.: A Unique Flight Test Facility: Description and Results. NASA TM-84900, NAS 1:15:84900. ICAS Paper 82-5.3.3. Presented at 13th Congr. of the ICAS/AIAA Aircraft Systems and Technology Conference, Seattle, Washington, August 22–27, 1982, November 1982, 83N13124, #. (See also 1210.) The Dryden Flight Research Facility has developed a unique research facility for conducting aerodynamic and fluid mechanics experiments in flight. A low aspect ratio fin, referred to as the flight test fixture (FTF), is mounted on the underside of the fuselage of an F-104G aircraft. The F-104G/FTF facility is described, and the capabilities are discussed. The capabilities include (1) a large Mach number envelope (0.4 to 2.0), including the region through Mach 1.0; (2) the potential ability to test articles larger than those that can be tested in wind tunnels; (3) the large chord Reynolds number envelope (greater than 40 million); and (4) the ability to define small increments in friction drag between two test surfaces. Data are presented from experiments that 214 demonstrate some of the capabilities of the FTF, including the shuttle thermal protection system airload tests, instrument development, and base drag studies. Proposed skin friction experiments and instrument evaluation studies are also discussed. 1224. Roncoli, R. B.: A Flight Test Maneuver Autopilot for a Highly Maneuverable Aircraft. NASA TM-81372, H-1176, NAS 1.15:81372. Presented at the AIAA Region 6, 32nd Annual Student Conference, Irvine, California, April 28–May 1, 1982, November 1982, 83N13115, #. A flight test maneuver autopilot (FTMAP) is currently being flown to increase the quality and quantity of the data obtained in the flight testing of the highly maneuverable aircraft technology (HiMAT) remotely piloted research vehicle (RPRV). The FTMAP resides in a ground-based digital computer and was designed to perform certain prescribed maneuvers precisely, while maintaining critical flight parameters within close tolerances. The FTMAP operates as a non-flight-critical outer loop controller and augments the vehicle primary flight control system. The inputs to the FTMAP consist of telemetry-downlinked aircraft sensor data. During FTMAP operation, the FTMAP computer replaces normal pilot inputs to the aircraft stick and throttle positions. The FTMAP maneuvers include straight-and-level flight, level accelerations and decelerations, pushover pullups, and windup turns. The pushover pullups can be executed holding throttle or Mach number fixed. The windup turns can be commanded by either normal acceleration or angle of attack. The operational procedures, control mode configuration, and initial simulation results are discussed. 1225. Harney, P. F.: Real-Time Data Display for AFTI/F-16 Flight Testing. NASA TM-84899, NAS 1.15:84899, November 1982, 83N13095, #. (See also 1219.) Advanced fighter technologies to improve air to air and air to surface weapon delivery and survivability is demonstrated. Real time monitoring of aircraft operation during flight testing is necessary not only for safety considerations but also for preliminary evaluation of flight test results. The complexity of the AFTI/F-16 aircraft requires an extensive capability to accomplish real time data goals; that capability and the resultant product are described. 1226. Carter, A. L.: Strain Gage Load Measurement on the Shuttle Orbiter. NASA TM-84898, NAS 1.15:84898. Presented at SESA 1982 Fall Meeting, Hartford, Connecticut, November 7–12, 1982, 83N12135, #. This paper describes the application of the calibrated strain gage load measurement method to the shuttle orbiter. Descriptions of instrumentation and calibration are included, along with comparisons of measured results with wind tunnel and FLEXSTAB analytical predictions. 1227. Fisher, D. F.; and *Dougherty, N. S., Jr.: Flight and Wind-Tunnel Correlation of Boundary-Layer Transition on the AEDC Transition Cone. NASA TM-84902, NAS 1.15:84902. Presented at the AGARD Flight Mechanics Panel Symposium, Cesme, Turkey, October 11–14, 1982, November 1982, 83N14433, #. (See also 1243.) Transition and fluctuating surface pressure data were acquired on a 10 deg included angle cone, using the same instrumentation and technique over a wide range of Mach and Reynolds numbers in 23 wind tunnels and in flight. Transition was detected with a traversing pitot-pressure probe in contact with the surface. The surface pressure fluctuations were measured with microphones set flush in the cone surface. Good correlation of end of transition Reynolds number RE(T) was obtained between data from the lower disturbance wind tunnels and flight up to a boundary layer edge Mach number, M(e) = 1.2. Above M(e) = 1.2, however, this correlation deteriorates, with the flight Re(T) being 25 to 30% higher than the wind tunnel Re(T) at M(e) = 1.6. The end of transition Reynolds number correlated within ± 20% with the surface pressure fluctuations, according to the equation used. Broad peaks in the power spectral density distributions indicated that Tollmien-Schlichting waves were the probable cause of transition in flight and in some of the wind tunnels. *Rockwell International Corp., Huntsville, Alabama. 1228. Saltzman, Edwin J.: A Summary of NASA Dryden’s Truck Aerodynamic Research. SAE Paper 821284. Presented at SAE Truck & Bus Meeting and Exposition, Indianapolis, Indiana, November 8–11, 1982. A combination of subscale wind tunnel model tests and fullscale coast down and highway fuel consumption tests have been conducted on baseline and low-drag tractor-trailer configurations. Fuel savings calculated for the low-drag configuration, based on the model drag data or the full-scale drag data, correlate quite well with fuel savings obtained from the over-the-roads tests at highway speeds. Subscale drag test for flow-vane and boattail devices provided drag reductions of about 48 percent and 15 percent, respectively, for bus- or motor-home-type vehicles. Full-scale boattail drag data are also presented. 1229. DeAngelis, V. M.: In-Flight Deflection Measurement of the HiMAT Aeroelastically Tailored Wing. Presented at 1st AIAA, SETP, SFTE, SAE, ITEA, and IEEE Flight Testing Conference, Las Vegas, Nevada, November 11–13, 1981. Journal of Aircraft, Vol. 19, December 1982, pp. 1088–1094, 83A13167, #. (See also 1165.) 1230. Gong, L.; Quinn, R. D.; and Ko, W. L.: Reentry Heating Analysis of Space Shuttle With Comparison of 215 Flight Data. NASA CP-2216. NASA Langley Research Center Computational Aspects of Heat Transfer in Struct., 1982, (see N82-23473 14-34), pp. 271–294, 82N23490, #. Surface heating rates and surface temperatures for a space shuttle reentry profile were calculated for two wing cross sections and one fuselage cross section. Heating rates and temperatures at 12 locations on the wing and 6 locations on the fuselage are presented. The heating on the lower wing was most severe, with peak temperatures reaching values of 1240 C for turbulent flow and 900 C for laminar flow. For the fuselage, the most severe heating occurred on the lower glove surface where peak temperatures of 910 C and 700 C were calculated for turbulent flow and laminar flow, respectively. Aluminum structural temperatures were calculated using a finite difference thermal analyzer computer program, and the predicted temperatures are compared to measured flight data. Skin temperatures measured on the lower surface of the wing and bay 1 of the upper surface of the wing agreed best with temperatures calculated assuming laminar flow. The measured temperatures at bays two and four on the upper surface of the wing were in quite good agreement with the temperatures calculated assuming separated flow. The measured temperatures on the lower forward spar cap of bay four were in good agreement with values predicted assuming laminar flow. 1231. Ko, W. L.; Quinn, R. D.; and Gong, L.: Reentry Heat Transfer Analysis of the Space Shuttle Orbiter. NASA CP-2216, NASA Langley Research Center Computational Aspects of Heat Transfer in Struct., 1982, (see N82-23473 14-34), pp. 295–325, 82N23491, #. A structural performance and resizing finite element thermal analysis computer program was used in the reentry heat transfer analysis of the space shuttle. Two typical wing cross sections and a midfuselage cross section were selected for the analysis. The surface heat inputs to the thermal models were obtained from aerody