RTI
Report
No.
TRR-33
Final
Report,
RTI
Program
RU-277
SURVEYOR
LANDING
RADAR
TEST
PROGRAM
REVIEW
CONTRACT
NO.
951603
I I
I
Prepared
for
Jet California
Propulsion Institute
Laboratory of Technology California
Pasadena,
January
24,
1967
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27709
RTI
Report
No.
TRR-33
SURVEYOR
LANDING
RADAR
TEST
PROGRAM
REVIEW
CONTRACT RTI PROJECT
NO. NO.
951603 RU-277
Submitted
to:
Jet
Propulsion
Laboratory Institute of Technology
California Pasadena,
California
D.
F.
Palmer
Approved
by:
P. G. V. Borgiotti [_ #t_,'-6,
Gene
Smith, Systems
Director Laboratory
Radiation
FOREWARD This report was preparedby the Radiation SystemsLaboratory of the Research Triangle Institute, ResearchTriangle Park, North Carolina, under California Institute of Technology(JPL) Contract 951603,a subcontract of NAS7-100. The work wasadministered by Section 273 of the Jet Propulsion Laboratory, Pasadena, California. Mr. S. A. Cohenwas the JPL coordinator for the contract. The programstudies beganon June I, 1966andwere completedJanuary 31, 1967. Participating RTI Staff Members ere: w P. G. Smith, Director of Radiation SystemsLaboratory D. F. Palmer, Project Leader G. V. Borgiotti, Member f Technical Staff o
TABLE OFCONTENTS
Page
CONCLUSIONS ANDRECOMMENDATIONS A. GENERAL RESULTS B. SPECIFIC RECOMMENDATIONS II. INTRODUCTION III. BACKGROUND INFORMATION A. INTRODUCTION B. ENVIRONMENTAL ANDSYSTEM-PERFORMANCE DEFINITION C. SUMMARY OFRADVSHARACTERISTICS C D. OUTLINE FANTICIPATED O RADVS PERATIONAL O PROBLEMS E. OUTLINE FRADVSUNCTIONAL O F DETAILS IV. "DESIRABLE" PROGRAM TEST DESCRIPTION A. PHILOSOPHY "DESIRABLE" PROGRAM OFA TEST B. "DESIRABLE" FLIGHT-READINESS PROGRAM TEST C. "DESIRABLE" SPECIAL EST T PROGRAM V. PRESENT PROGRAM TEST DESCRIPTION A. OVERALL PROGRAM OUTLINE B. TEST EQUIPMENT C. DEVELOPMENTAL ANDTYPE ACCEPTANCE TESTS D. VERIFICATION ACCEPTANCE AND TESTS VI. EVALUATION OFPRESENT PROGRAM TEST A. INTRODUCTION B. COMPARISON OFTEST SPECIFICATIONS MISSION WITH REQUIREMENTS C. COMPARISON OFPRESENT AND"DESIRABLE" PROGRAMS D. DOCUMENTATION ADEQUACY E. TESTING CONSISTENCY VII. SUGGESTED MODIFICATIONS TEST A. FLIGHT-READINESS PROGRAM B. SPECIAL-TEST PROGRAM APPENDICES A. BIBLIOGRAPHY OFDOCUMENTATION ANDLIST OF REFERENCES B. RADVSRANSMITTER T RECEIVER LEAKAGE C. SIDELOBE EFFECTS ASRELATED TORADVSEST T PROGRAM REVIEW
I.
I-I i-i i-i 2-i 3-i 3-1 3-1 3-5 3-12 3-15 4-1 4-I 4-5 4-16 5-1 5-i 5-1 5-2 5-10 6-1 6-i
6-1 6-10 6-18 6-18 7-1 7-i 7-11
A-I B-I
C-I
ii
Table
of
Contents
(continued_
D. E. F. G.
AVAILABLE DETAILS BUYER OF FAT
DETAILS THE
OF
VENDOR SYSTEM
UNIT
TESTS TEST
D-I E-I F-I
VENDOR
ACCEPTANCE
REQUIREMENTS OF THE STEA
LISTING SIGNAL SIMULATION
DISCUSSION TECHNIQUE
G-I
iii
LIST Figure 3-1 No. Antenna points RADVS RF RA 3-4 5-I and is and beam
OF
ILLUSTRATIONS
Page
configuration, into plane of RADVS. paper°) (Z-axis 3-6 3-7 of the DVS. (Section for 3-8 simplified signal shown as is shown. misalignment Fig. 6-2 for angle vs 6-5 assumed to computed curves of 6-6 block simulation typical of diagram portion all DVS of and 5-3 3-11 downward block preamp similar.) tracker, of The the RADVS
diagram° sections
Frequency Diagram STEA. RA
channel except and
channels
6-1
Spectral velocity assumed.)
width
velocity (See
magnitude.
relationships
6-2
Relationships Fig. 6-1. and
7-1
Stability circuit. Simplified audio
deviation-linearity
measurement 7-6
7-2
version and
of
discriminator
circuit
using 7-8
detection to STEA Blocks dotted
amplication. implementing in from double phase are spread-sprectrum lines splitter in are to existing switches
7-3
Addition testing. system; show position.
for shown lines
existing
mode;
switches
spread-spectrum 7-10
7-4
Block
diagram
for
an
on-board
spectrum
analyzer.
7-13
B-I B-2
lllustration lllustration shown is
of of
transmitter-receiver experimental of the system
leakage Antenna in Fig. B-I.
B-2
configuration. illustrated
a part
C-I C-2
Acquisition Response Fig. C-I.
circuit, curves of
block lowpass
diagram and bandpass filters in C-8
C-3
Detector lobe
outputs with
(Fig.
C-I) as
versus
cross-coupled
sideC-8
power,
PM/PccsL
a parameter.
iv
List
of
Illustration
(continued#
Figure C-4
No. Sidelobe Fig. 1500 tion C-I Hz suppression for Bb = (W.B.), and capability 8 kHz, when = O. Bt = Hz, and PCCSL _ of 300 = circuit Hz and (N.B.), Acquisiare PM 1/2. in
Page
Bd = 4 apply PCCSL
thresholds
interchanged Attenuator
C-12 C-II
characteristic.
lllustration of a large is separated effect the spectrum, spectrum. slightly spectrum occur G-I
of signal signals to
limiting plus (i.e.,
action a small
on
the
resultant For well the spectrum of
signal.
non-overlapping), to
change (before
a non-symmetrical limiting) center of signals, shifts of the of
resultant
a syn_netrical the the the large-signal effect is resultantsignal do C-15
centered at the For fluctuating different; center from limiting. relative axis spacecraft small that
larger
after
Sideband threshold DVS beam vertical.
levels vs of a
to slant at
the
signal for to
acquisition the the upper lunar G4
roll
range 5 ° angle
G-2
Signal component levels VSo range for sinusoidal frequency modulation of I kHz deviation at modulating frequencies up to about 1.5 kHz. G-5
circuit has been suggested(see AppendixC). Finally, the possibility of antennamodification could be more extensively investigated. These solutions should be analyzed and tested thoroughly in order to determine their adequacy and to detect any adverse conditions imposedby their use. In addition, tests of the simultaneous presenceof two signals in one receiver channel should be madeto determinewhether any natural suppression of the weakersignal exists. Without the modifications and associated testing, the CCSL problem is considered to be sufficiently serious to decreaseappreciably the the probability of mission success; although restriction of roll angle appearscapable of reducing the dangerof CCSL effects to acceptable levels for certain lunar approachangles [52], it is not an adequatesolution to the problem for all anticipated missions. The testing recommended aboveis estimated to require about six man monthsengineering time over a four monthperiod° (2) Morecarefully evaluate the transmitter-receiver situation. leakageproblem
Further tests are recormnended obtain additional information about to the characteristics of the transmitter-receiver leakagesignal under actual lunar descent conditions. The following two tasks are desirable: a) review previous vibration test and comparelevels with those measuredon the Surveyor 1 spacecraft to determine adequacy and possible need for retest; and b) perform onboard measurement the transmitter-receiver leakagespectrum. (These tests are of described in Section VII.) Suchinvestigations are important becausethe actual nature of the transmitterreceiver leakagesignal during lunar descent is still unknown. It is very desirable to learn these characteristics to determine their effect on the remainder of the Surveyorprogramand future programsinvolving similar radar-controlled landing systems. Performance the item (a) recommended of above is primarily a matter of data gathering andanalysis. It is quite possible that no further vibration tests will be necessaryif results obtained previously can be interpolated or extrapolated to Surveyor i conditions. This analytical work is estimated to _equire about four man-monthsf engineering effort. o The implementationof item (b) is estimated to require approximately three man-months ngineering for design, e construction, and testing of a breadboardunit. An additional period of about 1-2
I.
CONCLUSIONSRECOMMENDATIONS AND
A. GENERAL RESULTS Themajor weakness the present RADVS of test programappearsto be in the area of design verification (as opposedto flight acceptancetesting). In particular, deficiencies are believed to exist in investigations of sidelobe signal pickup, transmitter-receiver leakage effects, and retro tankage echodiscrimination. Of lesser importance is the apparent lack of design margin determination in environmental tests. Finally, the adequacy the ionization layer environmental test of remains in question becauseof unavailability of documentation. Thepresent flight acceptancetest programseemsto be basically complete except for absenceof full simulation of retro engine induced stresses. The adequacy
of certain portions, however, Tests accuracy, of is of concern because of lack tracker circuitry of realism and in the operaare insignal tion, volved. Documentation cedures ment listings. to be with the of the present program appears unit Test levels, tests are to be level adequate testing, except where are to be for test proequipconsimulation. range mark acquisition cross-coupled sensitivity, sidelobe converter
and
performance
This less
particularly
affects assembled.
tends
permanently of unit
requirements which not seem
generally variable.
sistent,
exception for
acquisition and system
Environmental
levels
completely
consistent.
B.
SPECIFIC The
RECOMMENDATIONS of this section Most of is to sunm_arize specific reconTnendations are A is given discussed rough here, resulting
purpose present Section
from detail and
the in
study. VII,
these Test each
recormmendations Modifications°" modification
in greater of on time past
"Suggested to fulfill
estimate based
manpower
required
experiences. The suring (I) recor_nendations mission Perform problems. As logic tions. is a result currently Also, an of previous being alternate studies [52],* to the cross-coupled-sidelobe potential problem suppression (CCSL) situasuccess: thorough analyses and experimentation of cross-coupled sidelobe are listed below in order of decreasing importance in as-
modified solution
[66] of
eliminate a
adding
small-signal
Bracketted
numbers
refer
to
references
listed
in Appendix
A.
i-I
three monthswould be required for complete incorporation of the circuit into a flight spacecraft. (3) Provide additional test equipmentand procedures to incorporate measurement klystron frequency coherenceand sweeplinearity of into the flight acceptanceprogram.
Problemsinvolving frequency incoherenceand sweep nonlinearity cannot be detected with use of the present test equipment. Yet, they can causeloss of sensitivity and false locks, as discussed in Appendix G. Loss of range accuracy is also a conlnon effect of sweepnonlinearity. The equipmentneeded,which is described in Section VII, is estimated to require about six engineering man-months nd eight technician man-months a for completion of six units. An additional two man-months ould be required w for installation at test facilities andmodifications of test requirements. (4) Provide additional test equipmentand proceduresto allow testing with realistic signal spectra in the flight acceptanceprogram.
The tracker, analog converter, range mark, and cross-coupled sidelobe circuitry are not completely checkedusing the present line spectruminputs, as noted in AppendixG. In addition, closed-loop descent testing lacks the realism necessary to fully check subsysteminteraction. The required circuitry, which is described in Section VII, would necessitate about two man-months f engineering and two man-months f o o technician time to complete a prototype. Construction and installation of all units wouldprobably consume additional six man-months f technician an o time. Thepossibility exists, however, that Ryanalready has someof the circuitry designed. (5) Thoroughlyexaminethe sufficiency of systemdesign and test requirements in view of retro-tankage effects.
Further analytical and experimental work should be performed to determine the range of effects the retro-tankage can cause. Theanalysis would consist of determining the possible profiles of retro-tankage separation from the spacecraft, and the use of these profiles for estimating the retro-tankage signal level and velocity combinations. Signals having these characteristics should then be applied to the RADVS from a signal simulator such as STEA evaluate the rejection capability and responseof the SDC. to
1-3
Theanalytical work described above is estimated to require about three engineering man-months. The requirementsfor performing the experimental work depends upon the range of signal levels andvelocities obtained from the analytical study. If the present STEA can supply these required signals, the test will be relatively simple; otherwise, special tests will have to be planned. (6) Modify present flight-acceptance test programto fill gaps.
Table 7-1 of Section VII indicates portions of the existing flightacceptanceprogramwhich are not considered to be adequate. With the exception of the unit acceleration tests, which are discussed separately below (8), these changesare mainly small items to increase systemconfidence. About two man-months f engineering time is expected to be required o to institute the changesin Table 7-1 which do not appearelsewhere in this enumeration. Full conformity to Table 7-1 also requires performance of items 3, 4, and 8 of the present summary f recommendations. o (7) Renew type-acceptancetesting to determinemargins of operation within the expected environmental conditions and to ana|yze fatigue effects of flight-acceptance testing.
Previous type-acceptancetesting appearedto lack the thoroughness neededto makeit valuable for RADVS, described in Sections IV.C.4 and as VII.B.3. Completionof the programwould probably require about 18 manmonthsof combinedengineering-technician time. (8) Adda constant acceleration test in the flight simulate retro engine deceleration. acceptanceprogramto
The argumentfor the needof this test is given in Section VII.A.I. Basically, the reason is that such an environmentcould easily imposethe most severe mechanicalstress on the system, and, therefore, eachunit should be tested for ability to withstand it. It is estimated that about 12 man-months ould be neededto place this w test in the program.
1-4
(9)
Add the
a
sinusoidal
vibration
test
with (if
the the
RADVS
operating is insure
to match realistic). thorough-
retro-descent unit level
specification test procedure
specification to
(10)
Provide ness and rigid
documentation
uniformity. acquisition systems. to JPL all Ryan engineering which and change might proposals otherwise (ECP) be to help sensitivity levels to assure rejection of sub-
(ii)
Set
standard (12) Circulate make only known to
system
peculiarities in design
evident
those
engaged
construction.
1-5
II.
INTRODUCTION
The purposeof the study programreported herein was to review the present Surveyor landing radar test programand to recommend desirable and realistic modifications. This effort wasdefined as Phase1 of an overall programfor achieving a higher confidence level in the ability of the Surveyor Radar Altimeter and Doppler Velocity Sensor(RADVS) systemto perform its function of enabling soft lunar landings. The first task of the study was to become familiar with the radar systemand certain parts of the test program. During this early period, the basic tenets of a test-program philosophy were also developed. Subsequently,detailed studies of a "desirable" test programand of the current test programwere conducted; to reduce biases of the former programby knowledgeof the latter, these tasks were undertakenas independently as possible. This approachis clearly indicated by the report outline: Sections III, IV, and V contain backgroundinformation, a "desirable" test programdescription, and the present programdescription, respectively. Following sections contain an evaluation of the present test program (mainly by comparisonwith the "desirable" program)and a set of suggested test modifications. Section I contains a surmnary f conclusions and recommendao tions. Several important conditions influenced the conductand conclusions of the program. First, the time schedule of the Surveyor Programis determinedby important factors outside the purview of the test-program review and is not likely to be caused to changematerially unless serious problemsare encountered. Second, from a time-duration viewpoint the Surveyor Programis entering its latter stages. Consequently, the current practicality of implementingsuggestedmodifications to the programis an uppermostconsideration. Thesetwo factors dictate that the test programbe reviewed from an adequacy standpoint rather than from a standpoint of improvement. A third condition which enters very strongly into the programis that completely realistic earth testing is out-of-the-question. Compromises between realistic testing under lunar conditions and reasonable testing costs and delays are clearly in order. Although careful consideration has beengiven to the desirability and usefulness of suggestedtest programmodifications, no attempt has beenmadeto place numerical values on the confidence levels (for successful RADVSerformance)to be achieved p by the various recormnendations.The RTI teambelieves that such numerical assignmentswould have little basis and therefore little value. Indications are given in the Recommendations Section of the relative importanceattached to the recommendations. 2-1
ALONG Frequency 5-40 40-1500 100-1500 (Hz) 2.5 2.0 2.0 Level
THRUST
AXIS Duration
g peak g peak g rms,
sinusoidal sinusoidal white
Throughout Throughout
powered powered
flight flight
gaussian 100-1500 4.5 g rms
random white random ALONG LATERAL AXIS
Throughout powered flight except lift-off and/or Mach Liftoff or Mach 1
1
gaussian
Frequency 1-2.5 2.5-40 40-1500 100-1500
(Hz)
Level
Duration double amplitude Power Power Power flight flight flight except lift-off
4.0 1.25 2.0 2.0
inches g peak g peak g rms,
sinusoidal sinusoidal white
gaussian 100-1500 4.5
random white
Power flight or Mach 1 Lift-off or
g rms,
Mach
1
gaussian (c) During fairing flat is Acoustic the Centaur to 20 Environment firing, be Hz no to
random
the greater
overall than
sound 145 db
pressure over 2
level • 10 -4
inside dynes/cm
the 2
Centaur a
estimated from Pressure
(with
spectrum (d) The
i0 kHz).
pressure (e) Pitch
changes rate rate
from
atmospheric
to
10 -4
torr
within
three
minutes.
The periods. 2.
maximum
pitch
will
not
exceed
5 deg/sec
during
thrust,
coast,
or
turn
Transit (a)
Phase and Vibration
Shock
Not
appreciable. (b) Pressure is anticipated to be less than 10 -12 torr.
The
pressure (c)
Temperature radiant flux is 4.40 BTU/Hr - Ft 2", reradiation is into a background
The at
incident
-460°F.
3-2
III.
BACKGROUND INFORMATION
A. INTRODUCTION Thoroughinvestigation of RADVS testing demands etailed knowledgeof the d three fundamentalelementsof the problem: (I) the environmentalconditions to be imposedupon the system, (2) the systemperformancerequired within the environments, and (3) the characteristics of the system. (Fewof the parameterscan be known with completecertainty, of course.) Presentation in this report of all information gathered would be of little value to those familiar with the Surveyorprogram. Certain details mustbe listed to support the analyses and conclusions, however. Thepurposeof the section being introduced, therefore, is to provide manyof the necessarydetails in a concise manner o A by-product of gathering the backgroundinformation was the uncovering of areas in which RADVSperational problemsmight be anticipated. An outline of these o ideas is presented as a logical extension of the details listed; more complete analyses are contained in AppendicesB and Co B. ENVIRONMENTAL ANDSYSTEM-PERFORMANCE DEFINITION Environmentalconditions are ascribable to the four main mission phases: boost, transit, midcourse, and descent. RADVS must operate only in the last phase noted (but must survive the others, of course). Details are outlined below and consolidated in Table 3-1: (The main source of environmental information is HAC document224800,Detail Specification, EnvironmentalConditions, SurveyorSpacecraft [5]. Also see [1,2].) i. Boost Phase 411 minutes)
(a) During The Static the Boost will cutoff. transverse Acceleration Phase, have At a static acceleration to a maximum cutoff, the the maximum of of 2.8 5.9 g will g's at will will be the be experienced. instant of
acceleration engine In (b) The the
increased Centaur direction,
booster 5 g's.
acceleration acceleration
approximately be 0. i g.
Vibration vibration levels are experienced at the S/C--Centaur separation
following
plane:
3-1
3.
The 4.
Midcourse expected Descent (a) environments Phase are less severe than in other conditons.
Shock retro-rocket of 5 g and ignition: a duration Terminal of peak sawtooth acceleration pulse
Shock with a
from magnitude (b)
250-350
msec.
Static
Acceleration burning, of engine axis.) the static acceleration (No significant reaches static 10.8 g along the
Due thrust appears
to axis
retro-rocket at the the end
burn-out.
acceleration
along (c)
lateral
Vibration due and
(retroburning) is a combination excitation of 2g (peak) sinusoidal applied at along
Vibration 100-1500 any axis Hz for
to retroburning g rms white of
0.2
gaussian 50 seconds.
independently
a maximum
time
Table STATIC PHASE Boost
3-1.
Summary
of
the
main
missiom
environments
expected
ACCELERATION Max. 5.9 g axis)
VIBRATION Max 4.5 g rms white Gaussian (both and at the on thrust axis lateral S/C
TEMPERATURE 50 ° (Data Surveyor Flight) to lO0°F from I In
OTHER Pressure: three minutes from atmospheric to 10 -4 torr Acoustic:
(thrust
Centaur Separation plane
White from
Spectrum 20 Hz to db torr
i0 kHz, 145 over 2.10 -4
Transit
Not
Appreciable
KPSM: 50°F SDC: 75°F Preamps: to 75°F (Data S/C
0 ° to i0-12 25 ° to 0° At of Van
Pressure torr
:
Radiation the center the outer belt Allen
:
from i flight)
Max. 6 1 x i0 protons/cm 2 sec. (> 40 MeV) and I x 108 electrons/cm 2 sec.
Descent
Max 10.8 g along the thrust axis (retrorocket)
Along any axes: Combined 100-1500 Hz, 2 g peak sinusoidal and 0.2 g rms white Gaussian for of a maximum 50 sec. time
Same
as
in Phase
Shock: Sawtooth Acceleration pulse of 5 g magnitude and a duration of _ 300 milliseconds
Transit
3-3
(d) Description of DescentProfile Details of the terminal descentprofile are outlined below and consolidated in Table 3-2: The relative speedof approachto the moon the slant range of about 60 at miles is about 9000 fps. At the 60 miles slant range, the altitude marking radar (AMR)generates a trigger signal. The following sequence events then occurs: of (i) After a delay cormnanded Flight Control Programmerstorage, the vernier into engines are ignited; (2) one second(nominally) later the main retro rocket engine is ignited; (3) about one half secondlater, power is supplied to RADVS. During the retrophase, the S/C attitude remains fixed and the S/C is in the inertial mode. TheRADVSltitude, velocity, and reflectivity data are telemetered a back to earth. Control of attitude is fulfilled using the vernier engines. Roll control is obtained by swivelling one of the vernier enginesabout a radial line perpendicular to the roll axis. The retro-rocket thrust slowly increases until a certain point after which it rapidly decreases° When acceleration reachesa the nominalvalue of 3.5 g, an inertia switch provides a signal to the Flight Control Programmer (FCP)to initiate the retro-rocket separation sequence. The thrust level of the verniers is increased to the maximum programmed level. After a fixed time delay (to allow the retrorocket thrust to be reducedto a negligible value), theretro-separation units are blown apart. After another delay to permit the retrorocket engine to clear the S/C, the FCP provides an arming signal which enables transfer of yawand pitch control to the doppler reference if the RODVS signal is present. Otherwise, the S/Cwill remain in the inertial modeuntilthe signalappears. In the time before RODVS present and in any case before reaching the optimum is (fuel-wise) descent curve, the vernier engine thrust is servoed to maintain a constant thrust-to-mass ratio equivalent to 0.9 lunar g. Theburnout condition must be within the operational ranges of the doppler sensors. The doppler radars are required to operate within the desired accuracy only for velocity smaller than 850 fps. . When optimumdescent trajectory is reached, the thrust is controlled to the bring the vehicle downthe desired range-velocity curve. At i000 feet, a signal from the radar altimeter will changethe Doppler Systemscale factor. At a speedof i0 fps the thrust control is switched to the doppler velocity reference. A constant velocity of nominally 5 fps is commanded, the pitch and and yawcontrol is switched to the inertial hold mode. A signal from the radar altimeter shuts off the vernier engines at an altitude of 14 feet. TheRADVS turned off after landing. is
3-4
Table 3-2.
Chronological sequenceof events during the descent phase
EVENTS ANDCONDITIONS i. AMR on 2. Vernier EngineIgnition 3. Main Retro Ignition (Vehicle attitude relative to the lunar verticle not to exceed45°. Attitude at acquisition not to exceed25° for engineering missions, 45° for scientific missions. Max. slant range for acquisition, 50 kft. Static acceleration not to exceed 380 ft/sec 2. Velocity magnitudeis +3000 to I00 fps.) 4. MainRetro Motor Burnout (BO) 5. Main Retro Casing Separation (12 sec after BO) (Vehicle Static Accelerations along the vehicle roll axis not to exceed12 ft/sec 2. Maxvelocity is 850 ft/sec.) 6. Inertial Modeat i0 fps velocity mark. 7. Verniers off at 14 ft mark. 8. Landing
RADVSEQ'T R Inact ive
turn-on retro quire 0.55 ignition when sec after (acpossible)
RADVS after RADVS
control 3 sec. Descent
enable
Control
Generate RADVS off
14
ft
mark
C.
SUMMARY The
OF
RADVS and above
CHARACTERISTICS beam the configuration spacecraft. Because DVS each channels DVS of Fig. of the 3-2 the will RADVS shows is an shown overall, of the here. split retain into two quadrature sense of but in the 90 db rein Fig. 3-I, looking block
antenna from of
downward diagram trackers,
simplified four frequency
this only
sub-system. one of the 3-3, PI
similarity be described is to passed Fig. to
Referring channels, ceived balanced the entire PI
to Fig. / O° and The
receive Beam i,
channel in order then in used taken
/90-o for doppler
doppler
signals.
two
signals which PI i00 is
are shown is
through The
separate signal
preamplifiers, doppler (i.e., band
one (i00
of Hz the
3-3. control from
contained of or
kHz)
the 40
gain-state db, 65 db,
preamplifier gate). range These (maximum
whether
signals keep 33 the db
are
the
gain-state signal of
switches approximately portion of
output above
signals the
within
a limited threshold). 3-3.
dynamic Major
acquisition in Table
characteristics
this
RADVS
are
summarized
3-5
YawAxis
/ t 1
\ Beam 2 I I
/ /
Ant. (Beams 2 &
2 3)
(Beams
Ia4)
Pitch
-X
Axis
+Z
Axis ! | \
/ Beam 3 / \ I I
Downward (Roll Axis)
-y
Fig.
3-i.
Antenna (Z-axis paper.)
and
beam
configuration, downward into
RADVS. plane of
points
3-6
4¸
3-7
004
Q;
i
>. P.
%
COO .,.4 O_ n._._
1
-F
I
g
rn
A v
|0_
3-8
Table
3-3.
Major RF and
characteristics preamplifier
of
RADVS
Beam
Configuration
--
See
Fig.
3-1
for
antenna-spacecraft
relationship DVS RA DVS Beams Beam ---25 ° off along 2.0 13.3 RA Klystron -250 12.9 RF Filters -To + watts GHz milliwatts, GHz reject spurious and filter shares with mixer mixers local components other in on-board Beam from RADVS + Z axis
Z axis (per beam)
Klystron
transmitters Additional cause Isolators -One it used
equipment. be-
i receiver altimeter. to
antenna each of
with four
mixers
help
maintain Mixers -Balanced to FM Preamplifiers -reject sweep;
balance. used in altimeter AM in order by in DVS.
oscillator
caused used
single-ended -i00
mixers kHz
Upper
cut-off
Low-frequency Velocity
roll-off channels -3 kHz corner in 40 frequency, and 65 db
6 db/octave gain at in Altimeter 30 Gain-State-Switches states; 1.2 90 kHz db
roll-off a second gives
corner
frequency roll-off
12 db/octave
gain
state. -same corner _ _ as above but with
channels kHz -and
5 kHz
frequencies. 0.2 sec
Time
constant
Hysteresis
i db
3-9
Oneof the DVSfrequency trackers is illustrated in Fig. 3-4. TheSSBM consists of a pair of balancedmodulators phasedin such a way that the lower sidebandsof outputs 1 and 2 reinforce for positive-doppler inputs and their upper sidebandscancel; negative-doppler inputs producethe opposite effect. This permits rejection of negative-doppler signals during searchby use of a limited range of frequency search, as explained below. The IF amplifier provides a I0 kHz "window"about the VCO frequency; the IF output is used to provide reflectivity data, as well as for frequency tracking. The two quadrature channelsbetweenthe IF amplifier and the discriminator provide sensing of frequency errors betweenthe input signal spectra and f c from the crystal oscillator. During the track mode, the discriminator output is applied to an integrator which controls the VCO frequency to drive the tracking error to zero. The search modeis initiated by application of a 0. i second"flyback" pulse to the integrator circuit. Dischargeof the integrator capacitor sweeps the VCO downward frequency until the sweep-limit switch is activated at f e + 800Hz. in
(The which of the lower limit the for VCO the RA is f
C
+
2kHz.) the upper
Another sweep
flyback limit.
pulse Tbe
is
then
generated parameters
returns sweep Start
frequency are
to
important
operation Sweep
(approximately):
Frequencies: burnout: burnout: 85 kHz kHz RA, above below ikft Ikft range: range: 91.5 22.5 kHz kHz
DVS,
before after
26.5
Search 60
Rates: kHz/sec for wide sweeps, 15 kHz/sec for narrow sweeps
Search has
ceases
whenever
the
signal the
passing threshold
through circuit is
the
tracker If the have
low-pass track analog this
filter concondelayed
sufficient for by at
strength least
to exceed the
level. to not
tinues verters gate
0. i sec., Gate"
tracker (The
output RA
applied does
"Doppler
circuits.
tracker
feature). The data conversion either section tracker if This the contains it appears circuitry, belief additional to be which that circuitry locked is onto termed for the the beams same 2 and echo as 3 the
which other lobe
unlocks (through logic,"wms
a sidelobe). based on
"cross-coupled would cause
original to
sidelobe
coupling
mainbeam
signals
in one
channel
exceed
corresponding
cross-coupled
sidelobe
3-10
%
3-11
signals have beams other true, by
in
the
other the in the
channel same
by
at
least It no
30 was
db,
and
that
the
two
signals only
would
essentially 2 and beams. and the 3
frequency. that
originally would be have mission possible.
employed experienced shown and
between the
belief
trouble and
between not the
Subsequent present of roll
measurements approach is if this
analyses each to be
this avoid
to be difficulty are
to analyze proves
selection
angle,
Other
solutions
under
consideration. velocity estimates are provided by the following relationships
Analog
VI-V 2 v ----; x 2A
V2-V 3 J ----; y 2A z
VI + 2B
V3
where
_. = _ I 2
fdi
% = _
(fvcoi
fc )
A = B = These
sin cos
45 ° 25 °
sin
25 ° = = are
0.30 0.91 in a straightforward is the subtracted resultant manner from to give calibration V y velocity I and 2 2f c by using the the DVS VCO
computations Beam frequency
performed
outputs. output A
3 VCO is
output subtracted coupled
frequency from with an are
; then
Beam
2 VCO of V z .
a digital constant, sense
measure then
frequency V z .
counter, Beams
appropriate used to obtain
gives
analog by using
2 and
3 VCO's channels,
; velocity being VCO's
is obtained directly to obtain
dual
quadrature
with Similarly,
analog Beams
obtained are used
from analog
the V x
sign-sensing . range f
C
circuit.
Slant against by a
is
obtained is always
from
Beam
4 VCO. and analog range. in Table a
The
frequency
of
this
VCO
is
beat
(sense
positive), An
frequency-analog measure of V z is
conversion subtracted
is made from this
frequency-counting to obtain RADVS
circuit. of slant
output
a measure outputs are
Other D. OUTLINE The CW nature
summarized RADVS
3-4.
OF ANTICIPATED anticipated the radar and
OPERATIONAL of RADVS
PROBLEMS operation and testing result arising from during the
major of
problems the
unusual
environmental can arise CW because class of
conditions of the
lunar leakage this
descent. problem,
Operational which is is
problems inherent fact to that
transmitter-receiver A no major serious conditons aspect of
the
radars. cause
leakage if
problem it were
the
it would by
probably the
operational existing
difficulty
not
greatly
aggravated
environmental
3-12
Table 3-4. Other RADVSutputs o
Range Marks -- i000 foot mark and 14 foot mark generatedby comparinganalog slant range and zener references. Altitude scale is changed i000 feet by changein FMdeviation (4 Mc at to 40 Mc) andby 2:1 changein analog circuits. CRODVS (conditional reliable operate doppler-velocity sensor)-generatedby "or" circuit with Beams 2, and 3 lock-on sigi, nals. Usedwith RODVS "or" gate to give RODVS into output. OnceRODVS signal has beengenerateddue to all beamslocking, the CRODVS signal is gated out (after one seconddelay). RODVS (reliable operate doppler-velocity sensor) -- generated by "and" circuit with lock signals from all three velocity beams,feeding "or" circuit with CRODVS signal. Usedto switch systemto RADVSontrol, once the initial cycle of operate under c CRODVS occurred. has RORA (reliable operate radar altimeter) -- generatedby "and" circuit with lock signals from Beams 3, and 4. i,
3-13
at the time of lunar descent. The instabilities induced on the transmitters and on the leakagepaths by retro and vernier engine vibration and by rocket plumes are the major contributors to the leakageproblem. As can easily be imagined, these unusual environmental conditions makeit difficult to test RADVSnder u realistic conditions. The operational problem causedby leakage is one of falsesignal lock-on; the false signals arise from modulation on the compositeleakage signal entering the pre-amplifier. Themost difficult modulation to correct is that on RFleakagepaths; however,other sources can introduce serious problems (e.g., vibration effects on the RFmixer which maymodulate the leakagesignal at frequencies up to several kHz). It is expected that most such spurious signals will fall in the doppler bandbelow I0 kHz. Other forms of false-signal lock can also occur. Onecause could be passage of the ejected retro tankage through one of the mainbeamso Although reflections from this source will havenegative doppler, its radar cross section is so large that the negative-doppler rejection capability of the receiver maynot be adequate; note that this capability is critically dependentupon the matchbetweenthe preamplifiers of a given channel. A secondeffect causedby passageof the retro tankage through a mainbeam would be to reduce the gain-state of the corresponding preamplifiers, in effect blinding the particular channel to weakerground-reflected signals. False lock can also be causedby cross-coupled sidelobe signals. These signals result from transmission on onemainbeam reception on a sidelobe of and an alternate beam. This problemcan become very severe for large lunar approach angles. Another type of problemwhich can occur is referred to as the "coherence-loss" problem. This problembecomes increasingly serious at the higher altitudes. Frequency modulation of the klystron transmitters will causea frequency beam betweentime-delayed echoesand the klystron reference signal to appear on preamplifier signals; This beam will causespectral lines to appear in the doppler band. In addition, serious spectral spreading of the preamplifier signal can result, with subsequentloss in acquisition sensitivity and in frequency tracking ability. Causesof the FMare microphonic vibrations in the klystron resonant structure and ripple on the klystron powersupply. Both AMand FMon the klystron output can pose serious problems. The effects
of AM, for the however, AM can be removed serious effectively spectral by the use of of balanced mixers. In order the to produce spreading ground-reflected signals,
3-14
depth of modulation must be several per cent; such severe cases would seldombe encountered,and if they were the accompanying FMwould usually causea muchmore serious effect than the AM. Another class of problemswhich should be considered in evaluating this test program is referred to as adaptive control errors. This is concernedwith the fact that certain RADVSarametersare programmeds a function of the position p a in a series of events which makeup the landing sequence. For example,at the generation of the i000 foot mark the RAklystron deviation is changed a factor by of i0. Simultaneously, the analog scale factor is changed. Similarly, the 14 foot mark is used to shut off the vernier engines to permit free fall for the remainder of the flight. Obviously, failure to perform these adaptive measuresat the proper time could result in mission failure. E. OUTLINE FRADVSUNCTIONAL O F DETAILS Proper operation under various environmental and dynamicconditions requires successful serial/parallel functioning of the manymoduleswithin the RADVSnits. u Consideration of all of the required processes is necessaryin any thorough testing program. For completenessof the present study, therefore, moduleshave been separated into functional groups which are the fundamentalelements of operational sequences; these are listed in Table 3-5 along with information necessary to help define tests. Table 3-5 will be used and analyzed in later report sections, but certain features should be noted here. First, the choice of grouping is not meantto imply that each group functions (or will need to be tested) individually. Instead, the intent is to group important characteristics which must not be overlooked in defining tests. For example, thoroughexamination of the klystrons' outputs also gives adequateinformation about powersupply and modulator operation; however, the definition of "thorough" mustbe basedupon the characteristics listed for the power supply andmodulator sub-units. It is also important to understandthat the numerical values given in Table 3-5 are not necessarily performancerequirements. In fact, most of themare adjusted as the systemis better understoodand refined. The values given in the table are mainly for reference; the only real criterion of successful performance must be based on systemfunctional requirements. Regular unit connectors are listed as test accesspoints in Table 3-5 whenever possible. Otherwise, moduletest points (TP) are given. Only unit connector points will be available, however,in most tests. 3-15
Abbreviations used in the table are listed below: BAL balanced BP band pass BW bandwidth CKT circuit DET detector DISC discriminator DTC dual time constant DVS doppler velocity sensor HV high voltage KPSMklystron power supply and modulator LP low pass LVPS low voltage powersupply NOM nominal QUADquadrature RA radar altimeter RCVDreceived R/T receive/transmit SDC Signal Data Converter TKR tracker VCO voltage controlled oscillator XMT transmit
3-16
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3-20
IV.
"DESIRABLE"
TEST
PROGRAM
DESCRIPTION
A.
PHILOSOPHY i.
OF
A "DESIRABLE"
TEST
PROGRAM
Introduction The following have been to discussion generated base and the the its summarizes at the outset philosophy a testing of the on philosophy, Surveyor only typical In of the ,,f one fact, RADVS
which RTI
might has
Program.
attempted requirements of was period and nature should,
testing performance test
a knowledge while In had
mission by the
characteristics, now in effect. personnel had
remaining to best the with
unbiased do this,
a knowledge write-up
actual
program after time
order
prepared at the of of JPL,
immediately at which profile. program, come from as
RTI team
finished familiar briefly a detailed the now
initial the to
orientation RADVS the of system general it. is basic RTI It
the
become had been
mission the test
Although they no great of the quite those had
they not
exposed
acquired that
knowledge described because between are is
course, different of a test
surprise test
philosophy in effect
below the the
not
greatly
that
program
principles suggestions to become
program tests as
are and
fundamental. actually being
Differences performed of the
for more
desirable apparent
expected given. 2. Any tions grams tests weighed failure. research bility must of be the of
a more
detailed
description
former
General test
Considerations which viewed for does with not duplicate exactly In fact, to the all be the from actual military encountered acceptable partial or operating and space condiproearly must be
program be
will allow under
always some actual
suspicion.
margin
unanticipated conditions.
problems Obviously, may any result
during margin
operating against
carefully It is an
penalties fact completely detect because realistic above
which that
complete for in space the capa-
unfortunate cannot be to
pre-flight Some operational part actual
test degree
program of risk
vehicles a test
realistic. prevent
program not just of
and of
failures failures operational
of but
a vehicle also because
accepted, impracticality the
unpredictable of
simulation may be so
conditions. it leads to
Although what is
point to be lack
made some of
obvious
to appear
trivial,
believed (i) some space (2)
important realism must in a be of
conclusions: pre-operational test program for
vehicles
accepted; the confidence of testing levels which may are be
numerical assigned
estimates to certain
portions
programs
rather
4-1
meaningless; the relative merits of manyaspects of different test programs,or in modifications to a given program,are basedon scientific and engineering judgements which are open to debate. The overall test programphilosophy described below in broad terms is intended to represent a goodcompromise betweencompleteand realistic testing and costs (in dollars and schedule). For further discussion, it is useful to consider test phasesas corresponding to the major divisions of the RADVSelivery program. Suchtest classifications d would be as follows: (i) Special tests (design verification) (2) Unit tests (unit construction verification) (3) Vendorsystemtests (systemassemblyverification) (4) Buyer systemtests on S/C (installation verification) (5) Prelaunch systemtests on S/C (launch configuration verification) The first of these phaseswould consist of special tests to determine whether problemswere inherent in the basic systemdesign coupledwith the environmental conditions and all of the anticipated descent profiles. The results of such special tests could of course range from re-design, through the imposition of individual test requirementson eachRADVS tests reveal marginal conditions), (if to the conclusion that no comparabletesting of eachRADVS necessary (if tests is reveal that no problemsare likely to be incurred). The remaining phaseswould be fundamentalto the preparation of every flight system; they might be termed "flight-readiness" test phases. 3. Special Tests
The operation testing is made As radiation The the nature basic to under evident indicated task permit as of special tests is to yield of enough information This of about implies realism system functional desired much simpler testing as flight systems. The the
realistic by noting
conditions the problems the
practical. for
degree intended
expected
application. conditions CW radar dependent and and
previously,
interaction
between
environmental strong is of poses very the for
performance of the
characteristics
is particularly leakage need signal
systems. upon range
transmitter-receiver conditions. with and The vernier possibly why
much
environmental
for operation operation difficult tests and to be
velocity most
sensors RADVS
simultaneously testing problem, several
engine the most
the
difficult problem.
operational of the
There tional conducted
are
reasons both which
realistic
(exclusive radiating flown:
actual
operabe
flight) on
involving spacecraft
environmental are intended
conditions
cannot
4-2
any low altitude operation of the S/C mountedin its upright position would be seriously hampered,and quite possibly invalidated, by the presenceof strong ground reflections; (2) firing the vernier engines during such tests is quite impractical becauseof contamination of S/C surfaces and components; (3) mounting the assembled S/C in an inverted position, in order to avoid groundeffects, is undesirable becauseof handling problems (with the possibility of damaging the system), and becauseof difficulties in operating the vernier engines in this position; and (4) the vacuum conditions existing on the moon difficult to are simulate in the earth's environmentunder conditions also permitting firing the vernier engines. The conclusion to be drawn from these considerations is that realistic testing of the environmental interaction with radiating performanceis impractical for an assembledflight spacecraft. This interaction maybe very important, however, and it is very desirable that any significant degradation of systemperformancewhich it causesbe evaluated and corrected, if necessary. It maybe possible to do this with a special "one-time" test performed on a mock-upS/C containing a partial RADVSystemand one or more vernier engines. Experimental evidence that no serious s problemexists becauseof vernier engine effects on transmitter-receiver leakage would obviously be extremely valuable in establishing a high level of confidence in the capability of the RADVS play its role in soft landing, without the to need for evaluating these effects oneachS/C.Ontheoth_ hand, experimental evidence of the existence of a serious problem, or of a marginal situation, would indicate the need for corrective action; after such action the experimental set up could be used for evaluating its effectiveness. 4. Unit Tests
These under ture, tests are defined as and etc.). units, The the those under Except testing nature of which can be performed on the units of RADVS laboratory vacuum, conditions vibration, radiating problems. of simulated for large environmental mechanical conditions units and (temperaactive prethe
(i)
electromagnetic sent no serious
under RADVS
simulated would
conditions that
should testing
indicate
electrical would
properties the major
antennas in
under the
realistic unit tests.
environmental It may, of
conditions course, be
present
difficulty
4-3
desirable to forego parts of such tests entirely, checking certain antenna characteristics in conjunction with other units during systemtests. For example, for a RADVSntennawhich has beenproven to be of soundmechanicaldesign, it is a believed to be unnecessaryto check the antenna pattern characteristics under varying temperatureand vacuum conditions. However,it would be desirable to test the antennamatchand transmitter-receiver leakageduring vibration. The major purposeof the unit tests should be to establish that eachunit fulfills its design requirementsand to yield confidence of successful future operation as a system. 5. Vendor System Tests (Ryan)
If system at the the tests vendor at the be unit is tests have been proper at the performed mating buyer of for of of very thoroughly, Such because with the tests it the only are requirement made of to assure than units. preferably to
(Ryan) former. desirable,
(Hughes) these added tests
is easier system
accomplish on the
fixes S/C
Duplication however, possibility in
installed
would
assurance. thorough and all unit tests is for unlikely each
Unfortunately, because unit. of
the
sufficiently all to signals be: of
difficulties
simulating are expected
environments
Particular (i)
problems of
simulation vibration;
structural
resonances
the
S/C
frame
in
(2)
simulation connecting
of all of of
electromagnetic components the heat thermal transfer to
interference the S/C; and which the
effects
without
(3)
simulation only S/C form is
environment between
exists
when and
the the
surroundings
radiative. environment-simulation system on a S/C, much of problems the can be completely usefulness solved of system withtests
Since out at
none
of
these the
installing Ryan is lost.
potential
Consideration tem tests should
of be
these primarily
different concerned
aspects with
leads
to
concluding system
that
vendor
sys-
verifying
performance
under
ambient 6.
environmental Buyer The System purpose and
conditions. Tests of the these other S/C to (Hughes) tests parts is of to the in check S/C. its in out Full the proper inter-marriages testing under attention the should severe should be
between be made
RADVS to
environmental state, Special
insure
proper
operation be
assembled space.
environmental
conditions
encountered
4-4
paid to testing those units susceptible to interference from other S/C systems (e.g., electrical noise pick-up on the klystron supply voltages). Although it is very desirable to radiate and receive signals from the RADVS antennas, whenthe complexities of locating the spacecraft so that these antennas "look" through essentially free-space toward remote targets are considered, it appears that a compromise aybe required, or at least maybe desirable from cost m and schedule standpoints. A first compromise would be to couple the RADVSntennas a through feed adapters andwaveguideto other antennaswhich could radiate toward and receive echosfrom special targets, such as signal repeaters which imposea doppler shift and bandwidthspreading on the re-radiated signals. In this manner, real delay is imposedupon the signals; this _s quite important to testing the range measurement klystron coherencelosses. Signal bandwidth spreading is also and important from the standpoint of differences in the responseof the frequency trackers and the analog output circuits to actual "noiselike" signals rather than to sinusoidal signals. Of almost equal value would be tests for which delay is producedby a long length of transmission line or a delay line (suitably operated at an intermediate frequency). Bandwidthspreading could be imposedby an active circuit inserted at any convenient point in the signal path. 7. Pre-launch System Tests (Cape Kennedy)
The of the the S/C purpose system be of of a of after these tests is and to other rather system check on the survival tests. and It proper operation that shipment pre-flight than is desirable to detect any and tracking be with
tests
functional
nature, An of overall time be
environmental, to check
degradation is very
components. Because They tests
test
sensitivity these previous
desirable. simple. new
and
facility
limitations, a back-up of
should tests,
relatively no B. basically
should
essentially
being
performed. TEST PROGRAM leads the to other the of it design flight can be of two complementary tests. more
"DESIRABLE" Consideration
FLIGHT-READINESS of one is and the of stated special for
philosophy tests first and study
testing The
programs, program
readiness approached system
latter
chosen it
because
systemmatically, and testing A in the
promises details. program
to yield
greater
insight
into
operation
requirement testing steps: the of
complete following (i)
can
be
generated
from
the
foregoing
information
_nspect a list
RADVS
Functional
Details which are
Table, minimally
Table
3-5,
to
determine at the
characteristics
sufficient
4-5
unit level to assure successful operation.* (The practicality of all tests listed need not be considered at this point.); (2) do the samefor the systemlevel; (3) determinewhich of the characteristics listed in (i) and (2) are likely to be affected by the environmental conditions described in Section III; and (4) combinethe results of the first three steps with considerations of test practicality and desired redundancy(for improvedreliability) to obtain a practical, thorough test program. (Further modifications would be likely during actual implementationof the program.) Assumptions about the extent of test signal realism are required before the steps listed can be undertaken. The mostbasic is that all units will be exercised with signals resulting from the full range of possible doppler and range signals. Other assumptionsare listed below so that they can be referred to numerically as needed: (i) range rates, doppler rates, and spectral shapeswill be realistically simulated; (2) a complete test with negative doppler and range will be performed; (3) the range signal will increase a decadein frequency during sweep return; (4) delay times corresponding to propagation delay from high altitudes will no___ttprovided during normal signal simulation. be The results of the first three steps, under the aboveassumptions,are shown in Table 4-1. The first four environmentslisted are onesduring which RADVS is to operate. Most of the characteristics checkedin these columnsare expected to be influenced by the environment; others are listed to check the system's or unit's "state of health." The last columnrefers to the nonoperatingenvironmentexpected at launch and during transit. Requirements checkedthere are mainly to ascertain general "state of health." Characteristics from Table 3-5 which are not included in Table 4-1 are listed in Table 4-2 along with an indication of why they were omitted from the former table. The last step in generation of the test programrequires a statement of criteria for determining the desirable sequence. Thesecriteria, which are mainly For the purposesof this program, testing below the unit level is undesirable becauseof the difficulty of simulating the manyinterconnection effects. 4-6
derived from the stated programphilosophy, are listed below: (i) Thoroughunit level testing is desirable becausethe analysis and correction of faults is generally less time consuming there than at the systemlevel. (2) Environmental tests should be repeated with the system installed on the spacecraft becausesimulation of the mission environmentis not likely to be very accurate during tests of individual units. (3) EMI tests are not likely to be meaningful at the unit level becausemost problemsare due to interconnections and grounding of units. (4) Constantacceleration testing of the entire spacecraft is probably not practical. (5) There is no basic need for testing the completesystem while not installed on the spacecraft except, perhaps, as a final reference test before leaving the vendor; sucha test need not be extensive. (6) Stability tests are easily handled by performing pertinent tests in every phaseand comparingresults. (7) A brief prelaunch test sequenceis desirable to check for damage during transit to the launch site. (8) Nonoperatingenvironmentsare anticipated to be imposed upon the entire spacecraft in the course of testing other systems; no unit level checks are required except for increased insurance of passing later tests. The resulting "desirable" preflight test programis given in Table 4-3. (The overall systemcharacteristics of "warm-uptime" and "powerconsumption"were addedat this point.) Details of performing the required tests are discussed in Sections VI and VII, where the present programand the "desired" programare compared and modifications are recommended.
4-7
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Table 4-2.
Listing of Characteristics in Table 3-2 for which Tests are JudgedUnnecessary becauseof Items in Table 4-1Noted.(See Text for Meaningof Special AssumptionNumerals;)
Characteristics Not Requiring SeparateExamination
Unit a. KPSM I. Ripple Voltage & Stability Supplies of & System Level
ResponsibleItems in Table 4-t
Special Assumptions Required
1,2,3,4
2.
Time
Delay
for
HV
Turn-on
3
3.
Sweep
Voltage
Timing
4
b.
R/T 4.
UNIT Separate Balances and Gate Gain & Phase Stages 13
for Preamp Matrices
c.
SDC 5. LVPS Regulation & Ripple 17,18,20,23,24
6.
Carrier & Extraneous band Elimination in
SideSSBM
23,24
1,2
7.
Spurious
Outputs
23,24
I
8.
IF Passband
Shape
23
1
9.
SSBM
& IF Amplifier
Gain
18
Stability
i0.
RA
IF
Gate
Performance
23
3
ii.
Proper ing Loop
Operation Gain, Bandwidth
of VCO of
TrackTime StaLinear 19,23,24
Components:
Constants, bility, Operation
4-12
Table 4-2.
Continued
Responsible in Table Items 4-1 Special Assumptions
Characteristics Not Requiring SeparateExamination
Unit c. SDC 12. & System Level
Required
(Cont'd.) Proper BP (Search Tracker Mode) SLP and 20 1,2
Filter
Operation
13.
Threshold
Detector all Preamp 20
Accuracy in Gain States
14.
Relative Signals Doppler
Phase through Gates
Shifts the
of 23
15.
Reference Stability
F_eq.
Generator 23
System a. KPSM i. Amplitude
Level
Only
Modulation
23
2.
Noise Outputs
& Other
Spurious
23
3.
Blanking and Timing
Signal
Amplitude 23
b.
R/T 4.
UNITS Transmitter-Receiver 23 Leakage
5.
Insertion
Loss/VSWR
20
6.
Noise
Figure
20
7.
Balance
of
Gains
& Phases
23
8.
Preamp
Gain
Stability
23
4-13
Table 4-2.
Continued
Responsible in Table Items 4-] Special Assumptions
Characteristics Not Requiring SeparateExamination
System b. R/T 9. UNITS Preamp Level Only
Required
(Cont'd.) Passband Shape 23
i0.
Preamp Gain Selection Accuracy & Hysteresis
23
1
ii.
Spurious
Outputs
18,23
c.
SDC 12. Tracker and Rates Search Ranges 20,23
d.
WAVEGUIDE 13.
ASSEMBLY 20,23
Performance
4-14
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4-15
C. "DESIRABLE" SPECIALEST T PROGRAM An anticipation of operational problemsand an awareness testing limitations of forms the basis for specification of tile special test program. As noted in the foregoing philosophy, checking of certain interactions basic to design and operation is not expected to be feasible or desirable on flight spacecraft. Thesetest areas are better knownnow, having delineated the flight-readiness program. The purposeof the present section is to itemize the extra tests required to yield high confidence of successful operation. i. Transmitter-Receiver Leakage Tests
A detailed description for of the leakage the of of power be the problem is given problem given here. the product of the in Appendix are B. Several and possible evaluated. In leakage to nate must is the order factor total if experiments Only to measuring summary RADVS leakage is described
a brief avoid and the
results
degradation leakage must
tracker in of the the
sensitivity, tracker-filter order of -160
bandwidth db. Stated
(normalized in of This an the altertracker
leakage this in
power)
way, be the
requirement order to with on
is not avoid CW
met
the on
acquisition the leakage The small
sensitivity component. combined
reduced major and
locking radar
problem of
difficulty modulation the in
systems. are met so
allowable that
effects estimation
leakage as a
this
leakage can be
a reasonable only by
to whether given of In system RADVS an
requirement the
is made it must by far
possible operate. biggest
experience unusual
with
environment lunar find descent
in which poses
The
environ-
ment
during to
the
testing
problem. of the The simulattransto to to the an
attempt
a reasonable possible
method
for were of
realistic suggested
measurement and
leakage first ing
problem, experiment above
several to be the leakage
experiments consisted the
evaluated. (or a
considered earth, firing
hanging engines in
a spacecraft and observing ground
system)
vernier
the
mitter-receiver an acceptable but the
signal. out is
The this
difficulty The order
reducing
reflections is similar
level
rules
method. in
second to reduce
experiment ground in
first,
spacecraft However,
inverted the
reflections the a
acceptable engines similar
level. upside method down is high and
difficulties the application
encountered of this
firing
vernier third and
discourage to to tether reduce the
method. above
Still the level, the three
a balloon-supported ground reflections
spacecraft to an
earth, again methods,
sufficiently observing the latter
acceptable Of
analyzing most
transmitter-receiver
leakage.
offers
promise.
4-16
All three of the test methodsmentionedabove have common shortcomingsof a serious nature. First, the acoustical air-coupling which exists in the tests, but is not present in the lunar environment, tends to maskthe desired results. In theory, this coupling can be reducedto an acceptable degree by various acoustical shielding techniques. An evenmoreserious difficulty is the fact that plume-coupling effects would not be realistically tested by any of the tests because the plumecharacteristics would be grossly different in the lunar environment than in the test environmentbecauseof the atmosphere. This limitation is believed to be sufficiently serious to discourage use of any of the three tests for studying the effects of vernier plume on the transmitter-receiver leakage. Only brief consideration was given to conducting tests in a vacuum chamber. Overall RADVSystemtests with vernier engine operation are impractical. Perhaps s a combined analytical and experimental study where relevant plume characteristics are measured and subsequentlyused to analyze the leakageproblemwould be very helpful. However,such a programwould be lengthy and costly and is believed to be impractical at this point in the Surveyor program. A completely analytical approachto the leakageproblemcan be conducted; one such JPL study was performed [67]. However,the uncertainties of plumecharacteristics and of antennacharacteristics (in particular the near-field levels outside the center of major field concentration), require that the computed results be viewedwith caution. It is believed, with the present state of knowledge concerning these uncertainties, that a completely analytical approachwould have very limited usefulness. An earth test is described which is believed to be very useful in evaluating vibration effects, but which will not test for plumeeffects. This consists of a two-step process: measurements the driving-force vibration characteristics of of the retro rocket and the vernier engines; and application of these measured vibration levels to an inverted spacecraft containing RADVS. (A modification would be to use Surveyor I vibration data which were obtained during retro fire and vernier engine operation of the lunar descent, rather than the data obtained as described in the first step.) During the secondstep all preamplifier output signals would be recorded and/or analyzed in order to obtain spectral plots of these signals. Becausethese vibration tests do not include plumeeffects, their value may be questioned. It maybe useful to point out that there are several mitigating factors to the plumeeffects; consequently, those tests described abovewhich do 4-17
not include these effects are still
follows: (i) Plume doppler circuits coupling. (2) It is expected that coupling band; will will be
quite valuable.
predominantly in
The
mitigating
factors
are
as
the
negative rejection of this
consequently provide
negative-doppler rejection
significant
plume character the
coupling with a
will
have
a
random, bandin
thermal-noise-like width; the consequently,
fairly
wide threshold
noise-developed will provide so
acquisition of
circuits
a significant that false-lock non-thermal is undeis
degree not noise as
receiver to
desensitizing occur as for such to not
likely
narrow-band
components; it
although
desensitization lock.
sirable, Unfortunately, plet_y spectral receiver correct only for The ignoring
is preferable arguments lack are of
false
these plume
sufficiently about
conclusive the For lock, degree example, may is than also
to of
justify coupling much
comand its
effects; is of
knowledge
characteristics desensitization, lock-on the final to desired portion studied are this VCO
considerable it may (This doppler on-board during
concern. avoid false
too
although signals. of is the an
prevent to occur
desensitization band, test say, where less
expected i0 kHz.)
lower test signals
spectral lunar is
characteristics Two promising analyzer
of
preamplifier methods employing The other of a
obtained data are steps
an
actual One
descent.
obtaining stepped
described. a narrow-band many of contiguous of
a simple through bands
spectrum the
which
filter spectral
doppler the
band.
simultaneously band last test) of two interest tests
observes by means described to be
covering
doppler The on-board be
banks (the
doppler
filters. vibration and test are and the to
above very test
earth-bound and
are
believed of
useful program.
practical
considered
valuable 2. It
parts
a desirable
Flight is of
Tests to in conduct order It of to and is to a series verify of its from flight tests on an early experimental certain completely In opera-
desirable the radar
model
operational the foregoing
capability discussion cannot be
under
realistic realistic fact, tion there during
conditions. simulation appears retro
apparent
that achieved. RADVS for
lunar be no
environmental practical firing. its design way
conditions to simulate RADVS
realistically can be tested when
vernier
However, for
high signals
altitude
operation
to verify
proper
operation
realistic
4-18
are present. The major attraction of such tests is that they test the system's capability for acquiring and tracking low-level signals which have realistic fluctuations and spectral characteristics. Klystron frequency instabilities will showup during such tests as a "coherenceloss" or, stated another way, as a spectral spreading loss; such instabilities will produceno observable effect during ground tests in which only small delays are imposedon the received test signals. Any anomalies of acquisition, tracking, and signal processing of realistic signals will be discovered during such tests and corrections can be made. Although preamplifier noise signals resulting from transmitter-receiver leakage will not be a good indication of those existing during lunar descent, the reduction of such components acceptable levels will certainly enhancethe RADVS' to capability for operating under lunar descent conditions. Fromthe standpoint of such noise characteristics, then, the high altitude tests must be viewed as essentially qualitative in that they highlight trouble spots which require corrective action. If during the flight tests certain problemareas are discovered which are sensitive functions of environmental conditions, correction of these problemsfor the flight tests alone maynot be sufficient. For example, if during these tests marginal corrections are madefor the transmitter-receiver leakageproblem, special attention should be given to additional tests which ensure that lunar descent conditions will not seriously aggravate the problem. The flight tests should be conductedunder conditions which are as realistic as possible. Operating altitudes should preferably be as high as 40,000 feet and the antennashould be tiltable from 0° to 70° relative to vertical (i.e., the limits encounteredfor RADVSescents). Thealtitude requirement cannot be d met by the helicopter; becausethis is otherwise a good choice it maybe desirable to compromise the altitude requirement. A subsonic, fixed-wing aircraft on cannot provide the hover testing of a helicopter, but generally offers a superior "flying laboratory" becauseof the greater available space(as for exampleoffered by the KC-135). Altitude limitations of someaircraft can be partially compensated y inclusion of flight tests conditions which present low b signal level; flights over smoothseas or flat sandyterrain offer oneway of satisfying this condition. Themajor deficiency of flight tests, as described here, is that lunar
descent vernier viewed descent. for radar conditions firing as are not are realistically not present. of RADVS' simulated; Therefore, capability tests are in the for particular, flight tests retro and be effects cannot lunar a necessity
complete In spite
verification of this
controlling as
limitation,
such
considered
design
verification.
4-19
3.
There outputs. nals, correct One The so be
InterferinR will, It is of
Signal cours_ be
Tests undesirable determine can signals whether cause appearing false-lock in can the preamplifier on such sigthe
important their
to
occur on
or whether signal. such
presence
deleterious
effects
tracking
undesirable and velocity
signal
arises
from of tank
reflections this signal
from should
the be
retro-rocket quite
tank.
amplitude that
distribution of of retro
predictable, beams negativeit. there However, is a should
realistic The
simulation velocity capability short it to range cause scatter the a
passage will will be
through negative; discriminate
the
antenna the
possible.
this of the
target circuit
thus,
doppler because good
rejection of the for It
against the signal, as
and
relatively-high Possible transmitted below the
strength effects and/or tracking
of
chance (i)
difficulty. enough echo
are
follows: signal this is
might
received threshold;
to drop considered (2) It might signal expected (3) It might
lunar normal
effect. enough the preamp energy gain to suppress this the is lunar also an
back-scatter by switching result. back-scatter image is large image
level;
enough enough might
energy to even
so
that
its
positive tracker some cases.) amplitude at
doppler behavior. (4) It might to pass
cause be
erratic in
(Its have through
tracked and or high
low a
enough tracker's
frequency lowpass
enough
bandpass
filters
a significantly (5) Its presence
high in mixers
level. with true signals might cause
trackable Because performed Another which cases have these of these with
intermodulation very
components. thorough design verification tests should be
possibilities, such signals. of interfering to present are small in a
source been
signals a serious to
is
through
cross-coupled (Appendix mainbeam The main C).
sidelobes, In all practical is
shown
problem the
signals
relative given
correct
signal, concern Although
which is
simultaneously to the incorrect
present signal,
receiver have
channel. disastrous
lock-on analysis
which
could
results.
4-20
showsquite clearly that the present RADVSill normally lock-on certain crossw coupled sidelobe (CCSL) ignals, the effect is important enoughthat it should be s thoroughly tested. For example, the test would showwhether there is somenatural weak-signal suppression in the receivers and trackers, which is not discovered by analyses assuminglinear-circuit operation. Changes presently being madein are RADVS include CCSL to logic for all beamcombinations; a very thorough analysis and testing of the resulting systemshould be made,at least one time, to discover any unanticipated interactions of such multiple-logic circuitry. Theabove discussion would indicate that interfering signal tests and complete tests of any CCSL should be run as a special, or one-of-a-kind, test. However, fix tests of negative-target rejection capability could easily be run on each RADVS system. Decision of the extent of testing in the flight-readiness programshould be based on operating margins found in special tests. Narrowmargins are dangerous becauseproper operation dependson critical circuit balancesto eliminate negative doppler signals in the trackers. 4. Environmental Overtests
The small dence number much be The of systems available for special to be thus with a a of large testing obtained; determined, is anticipated i.e., rather as to little "mean however, be too confito for could statistical assigned coupling to of significance any quantities
such analysis,
time can
failure." contribute The component to
overtesting without be
engineering number of
useful basis for
information testing Since
test
samples. starting be at the
should statistics unit
system component
reliability failures
analysis can
level.
generally can be then
obtained within to
sufficient
confidence, intervals. various of
and
system
failure
statistics tests
computed be used
useful reveal
confidence whether
Properly
instrumented were would for
would
component the program test
interactions mentioned requirements overtesting tests on the
correctly indicate the
anticipated. design modification program. predict needs.
Early Later,
results would phase of
they
determine of
flight-readiness be employed of
Another the effects the
environmental
should system. a few
to help the show
flight-readiness
A cycling times
system any degrada-
through tion
anticipated be of
flight-readiness from special testing. tests Basically, to the
program
would
that The
might
expected these
details
cannot
be
listed all
without
knowing
the
reliability would be
analysis varied from
results. a low level
though, point where
environmental became
conditions imminent.
failure
4-21
V. PRESENT PROGRAM TEST DESCRIPTION A. OVERALL PROGRAM OUTLINE The Surveyor test programhas four main facets: developmentaltests, type acceptance(or approval) tests (TAT), reliability tests, and flight acceptance (or approval) tests (FAT). The first two types of tests differ from the latter in that they do not generally involve flight spacecraft. Theyboth have the basic task of proving the design but differ by their positions in the program sequence. The third, reliability tests, can involve special sequences either on flight or test vehicles but is normally entwined in unit construction and regular FAT. The fourth set is used to determine flight readiness of systemswhich must actually perform the missions. An additional test group within the programmight be termed"quality assurance tests." This group is actually a part of construction which helps assure passage through other tests; it will not be considered separately in the study. Similarly, reliability tests will not be viewed as a separate group. B. TEST EQUIPMENT All of the formal type acceptanceand flight acceptancetesting by the buyer is performedwith use of SystemTest EquipmentAssemblies(STEA's). There are about three STEA'slocated at the E1Segundofacility, two at the Eastern Test Range,and one used at other installations as needed. In addition, the sametype of RADVS test equipmentassemblyis usedby the vendor for systemFAT. All of these assemblies can be considered to be identical for purposesof the present study. Details of STEA contents and operation are found in HAC publication 6594500, "STEA Operation and Maintenance Manual,"Vol . I and II. For completeness disof cussion, an abbreviated diagramof the portion of STEA which provides simulated signals to RADVS shownin Fig. 5-1. Other STEA is connectionswith RADVSre posa sible either through adapters placed at the normal moduleconnectors or by use of the spacecraft's telemetry system; the latter requires use of STEA'sRF test racks. Throughthese connections STEA permits examination of preampoutputs, tracker lock indicators, range marks, blanking signals, CRODVS indicator, RODVS indicator, RORA indicator, reflectivity outputs, analog outputs, and preamp gain state signals. An eight channel oscillograph (Brush, mark 200) and a digital voltmeter (Nonlinear Systems,484A)can be selected to monitor most of the signals. In addition, STEA provides indicator lampsshowingthe states of the RADVS bilevel-signal outputs. 5-i
Provisions for loading and filtering the RADVSnalog outputs are also cona tained in STEA. The purposeof the loading is to simulate the normal spacecraft (Flight Control) terminations wheneverthe actual connections do not exist. The reason for filters is to simulate the spacecraft responseso that effects on operation of analog output noise and ripple can be determined; the filter transfer functions are G(s) = 5 (2.6 s + i) (0.ii s + i) 2
for Vx and Vy, and G(s) = (0.08 s + 1)2
for Vz [35]. Monitoring can be performedeither with or without the filters. Another capability of STEA to simulate spacecraft dynamicsin closed loop is control tests. The simulated signal received by RADVS these tests is the same in as shownin Fig. 5-1 except that the input frequencies are determinedby voltage controlled oscillators (VCO's) instead of the sourcesshown. TheVCO's, in turn, are driven by signals obtained from computed spacecraft motion. Therefore, the only real difference to RADVS that its simulated return signals vary in freis quencyrather than remain essentially fixed. An evaluation of the use of STEA will be withheld until evaluation of the entire program. At present it will be pointed out that only the simulated return signal is essentially a single sinusoid which tracks the current transmitted signal with negligible time delay. (Also, see AppendixG.) C. DEVELOPMENTAL ANDTYPE ACCEPTANCE TESTS i. Vendor Tests *
Type 1964; approval report tested the tests at the was vendor, the Ryan, were essentially for serial conditions shock, (EMI). as the completed these number were tests. one in late The (S/N-l). Ryan 51765-IA was Ryan the TAT, controlling model document produced,
specimen During
first the
regular following
environmental
applied low are temper-
(separately): ature storage, in
vibration, and Table 51766-1
constant
acceleration, interference
thermal-vacuum, These phases and on
electromagnetic 5-1 with an (The
briefly in
described Ryan report
outline poor
of
results of
recorded captions
analyzed
[61]. complete vendor on
quality
the
reproduced
records
precludes of
re-analysis.) developmental type acceptance 5-2 tests was not is available for herein. the present study.
Documentation Available
information
tests
presented
I
5-3
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0 "0
_0
44 O 0
;:1 0
m
_D r_ 4.-I U 0 0 14 0
:> c_
"_ q_ 0 ('_ Z
0 r_ _ 0 0_ L_ u 3 N 1._ ¢1t q_ 0
_o
eq .--I I N
_
o_
e_
g
Z
o0O
,-4 0 o_ [-_ Z 0 ._
_u
0
_'_ _._
0 U
g'_ ..z
[--t rD
_
d
d
0 Z
d
0 0 _m
0
0 Z 0 z
0
0 Z 0 Z
d
0
[-.-I
5-5
Certain (I)
additional The unit was
features
are
noted
below: was applied by driving shaker. structural to placed to be but made have on a each No RADVS attempt
vibration separately made to the
environment with an the
electromagnetic spacecraft's intended
simulate mountings with called used
characteristics. flat response.
Instead, Monitoring Whenever and test the was
were
was tests KPSM
accelerometers for in the units
the
mountings. R/T unit units under
operating, only with the damped by
were
conjunction were SDC were to
vibrated; Signals acceleration
connections for the
flexible
waveguide. (2) Constant use the (3) Shock Sand the 2. Buyer of Tests the overall of
provided each unit
oscillators. with of
was Units
applied were
separately a portion
a centrifuge.
operated
during
test. testing Drop tests. was performed model 73. on each No unit with were use of a Barry during
Machine,
units
operated
A view is useful listing RADVS is
spacecraft the a more
special
test of
program
by RADVS
the
buyer,
Hughes, Such a
in understanding given the below; list.
relationships detailed
special of
tests.
description
the
portions
involving
follows DESIGN I.
ACTIVITIES: Reduced designated of 2. obtaining scale M-I and full scale M-13, design full scale spacecraft were models and for mockups, purposes
through
constructed
subsystem scale and
compatibility. models, MA-I and MA-2, were
One-forth used for MT-I,
antenna a full
tests. scale was spacecraft used for model evaluation with of thermally thermal simcon-
3.
The
ulated trol 4. A
components, provisions. with static
spaceframe given
components and vibration
simulated structural S-2A, was
by
point tests.
masses,
S-l,
was 5.
A more shock,
elaborate and static
spaceframe, loadings. system S-7.
tested
with
vibration,
(1963-65). tests were of performed these, was using which used space-
6.
Vernier frames
propulsion S-4 through
The S/C and
last
employed to establish
durmny masses vibration
to simulate for FAT
components, TAT. (Through
levels
mid
1965).
5-6
7. Spaceframe was used for flight control/propulsion interS-8 action tests at the Air Force Missile Development enter (AFMDC). C 8. The S-8 Spaceframe also wasused for RADVSibration tests in v the upsidedown position. (late 1964) 9. TheS-IO framewas used to determine thermal performanceof all subsystems and qualify the S/C thermal design. (early 1966) i0. The T-I test vehicle wasused for drop tests of the landing gear and for spacecraft/Centaur separation tests. RELIABILITY ANDSYSTEM-TEST ACTIVITIES i. The T-2H "vehicle" wasan installation of the QA-I RADVSnd test a equipmenton a helicopter to evaluate RADVSerformance. (design/ p development hase: mid 1963; veritication phase: mid 1964) p 2. TheT2N-I/-2 test vehicles were used for RADVS/flight control/ vernier propulsion subsystemtests during descents from a balloon. TheX-3 and X-4 RADVSasused. (Sept. 1965through w May1966) 3. The T-21 prototype vehicle, which is essentially identical to flight model spacecraft, wasused for the formal SystemType Approval Test Program;its purposewas to verify design and to check compatibility with ground equipmentat the Eastern Test Rangeand deep spacenetworks. It used the QA-I RADVS. 4_ Spacecraft SC-I and SC-2were used to check noise generation characteristics. (mid 1965and Jan. 1966, respectively) As noted, the formal TATmadeuse of the T-21 vehicle. The portions of this programwhich affected RADVS were the SystemFunctional Test (SFT), Vibration Test (VT), and Solar-Thermal-Vacuum (STV) Test Phases[13,14,15]. The SFTphase was for systemperformanceverification and calibration in normal laboratory surroundings. Theother two phasesare outlined in Table 5-2. All tests were performedusing a systemtest equipmentassembly(STEA) similar to that described in report Section V.B.
5-7
Table
5-2.
Listing TAT Using
of
the the
EnvironmenL_l T-21 Vehicle
Portion (with QA-I
of
Fo_m_l
System
RADVS)
ENVIRONMENT
TESTS VIBRATION: g peak, (2.0 g peak sinusoidal, axes, @ 2.0 (4.5 PLUS g rms in swept 40-100 Hz in i00XMTR power Tracker sensitivities
1,
lab
ambient
after 5-40
(launch) Hz @ 2.25 z axis swept on Hz for
io
sinusoidal, @ 1.20 lateral 1500 along lateral during 100-1500 100-1500 on three Hz
2.
g peak along directions); @ 2.0 g peak for 100-1500
three random
bandlimited z axis
i0 minutes
g rms
directions (descent) Hz @ 2.8 Hz random axes.
2 minutes). swept PLUS g rms for sinusoidal, 2 minutes i. 2. 3. Closed loop terminal descent test. XMTR power Tracker sensitivities
2.
VIBRATION: g peak, @ 0.2
bandlimited
3o
after Solar i00
IONIZATlON-layer-simulation Vacuum and % mm Chamber Hg; + and at pressures STV of after
in 130
the + 5
i. 2.
between
XMTR power Tracker sensitivities
watts/fE Z , -310 i x I0 -v torr
10°F
background,
and-
The (For
salient
features see
of
other 13,
special 14, 16,
tests 22, 33, to
involving 34, 35,
RADVS 38, an 42, A-21
are 43,
described and 48
below; in Appendix A.)
details, I.
references An S-8
VIBRATION: eject at was the
spaceframe in an
fitted
simulate by
vehicle system. to
after Shakers overall the
retro
supported three
inverted
position were with level were state on the on
a shock-cord with noise
attached force 80 from to
vernier i0 (An and
engine 56 pounds
points rms
driven flat of
obtain to
outputs 2000 engine Hz
between range.
spectra i0 pounds
bandlimited rms was
expected Preamp preamp were
mission outputs gain monitored of
established tape (i0 kHz
firing Analog and of one
tests.) outputs, marks spectral
recorded number a 2,
on magnetic tracker lock
bandwidth). RODVS, sequent filter results RORA, plots (and were: (a)
signals, Subusing a 50 Hz
range the second
galvanometric outputs were
recorder. made
content
preamp the
integration
time)
(looped)
tape
playback.
Significant
A
tracker
with to
a 3 db
acquisition but (The those
threshold with a
was 9 db
very level greatly
sushad in-
ceptible no
false
lockon,
significant tracker high gain
trouble.
higher to
threshold double of
creases in the
desensitization state over
sideband interest.)
signals,
frequencies
5-8
(b) An antennawithout shockmountingand with different surface coating showedappreciable return from foot pads and crushable blocks. The broadbandpowerwas sufficient to switch preampgain states. All DVStrackers were susceptible to false lock from this unit's output at force levels of 28 and 56 pounds. (c) Evidenceof leakagebetweenR/T units was noted (but not completely analyzed). (d) Isolators were found to be required. (e) The altimeter was stated to be so insensitive to vibration that no data for it waspresented or analyzed. 2. FLIGHT TESTS: The T-2Hphaseof the T-2 test programflight-tested RADVS with use of a helicopter. Themodel used in the tests contained all of the main features of flight models. Themaximum altitude flown wasabout 6,000 ft over the terrain. On-boardinstrumentation consisted of a magnetic tape recorder for preampoutputs, analog outputs, range marks, and reliable operate signals; a recording oscillograph for tracker lock signals in addition to those mentioned; and a camerato record the terrain being viewed by RADVS.This samesignal information was also telemetered. Data analysis included spectral analysis of preamp outputs and comparisonof analog outputs with optical tracking data from ground installations. Significant results of the 1964 tests were: (a) Analog output accuracywasgenerally within tolerance when the systemwas tracking normally. (b) The 14 ft range mark wasfrequently triggered by noise at ranges greater than 18 ft. (c) In flights over water, trackers 2 and 3 locked onto beamone through a sidelobe. Also, the CCSL logic betweentrackers 2 and 3 was found to operate properly over water, but no such situation could be imposedover land. (d) Noise on the analog outputs appearedto be higher than expected. (e) Checkof altimeter performanceover rough and mountainous terrain showedsatisfactory performance. (Accuracywas not checked.) (f) The DVS analog outputs wereperturbed whenpreamp gain state switching occurred.
5-9
3. DESCENT TESTS: Descenttests were performedwith the T2N-I and T2N-2 vehicles, which are special frames fitted with RADVS, flight control, and vernier engine propulsion subsystems. Mainmodifications madeto RADVS test purposes for included: (a) altering the waveguideruns to fit the frame; (b) locking the RA in the high deviation mode bypassing by the deviation control SCR; (c) disabling the signal-to-noise acquisition modeby disabling all preamp high gain threshold detectors (to mitigate vernier engine noise degradation); (d) bypassing the cross-coupled sidelobe logic circuitry; (e) restricting DVS operation to the narrow-bandmodeby applying a permanentburnout signal; and (f) restricting RAoperation to the narrow bandmodeby providing a permanentdeviation signal to the tracker filters. Teleme_red data included the 14 ft mark, reliability signals (except CRO),analog outputs, i0 fps detector, preampgain states, somepreampoutputs, and tracker lock signal. Tests were run from releases at about 1,450 ft to parachute recovery at about 600 ft and from releases at about 900 ft to landing. Significant results were: (a) The DI tracker locked onto leakage from the RAXMTfeed. Problemwasdiminished by tuming the RAklystron for reducedAMand by adding isolators to the DVS XMT waveguide. (b) Transients appearedin analog velocity outputs at preamp gain switching points. (c) Mechanicalisolation of the klystrons was found to be needed. (d) All other operation was considered satisfactory and within tolerances. D. VERIFICATION ACCEPTANCE AND TESTS i. Vendor Unit Tests
Unit procedure tests are which outlined are in SDC KPSM RA/VS DVS performed the Test Test as part of the vendor construction verification following Ryan documents:
51765-9 51765-10 51765-11 51765-12 51765-13 51765-14 51765-16
Requirements Requirements Test Test Requirements Requirements Test Tests Procedures Procedures
Antenna Antenna
Antenna Special KPSM
Manufacturing Temperature Test
Ranging
5-10
Unit tests which form part of the buyer's acceptancetest proceduresare outlined in Ryandocuments: 51765-2B,Part III, Unit AcceptanceTests 51765-2B,Part II, EnvironmentalTests (unit vibration only) Tests performedare outlined in Table 5-3. Other available details are contained in AppendixD. 2. Vendor System Tests
The simulated conditions. Tests and other vendor system tests consist of All operational tests is Ryan Table 5-4. E V.B. are checks performed number standard test during under a sequence laboratory Part of ambient I, [60]. (STC) used is operational The performed details the conditions.
controlling are are same
document in
report The The
51765-2B, test
outlined
conditions assembly
contained as described
in Appendix in Section Tests and
equipment
essentially 3. The document completeness are been outlined Buyer total
FlightAcceptance Hughes A, the test
sequence
requirements A-21) System of
are Test this
concisely
described [31].
in
HAC For RADVS have
3023926 of
Surveyor present below. inclusion Tests
Spacecraft report, In the
Specification document which to
contents test F. performed
affect RADVS
briefly for
addition,
requirements
relating
reproduced Flight
in Appendix by Hughes and are
Acceptance completely the
only units
on
vehicles used must
which have FAT,
have
been
(essentially) torily is tem passed
assembled lower
aligned. (vendor)
All
satisfactherefore, that of sys8
appropriate with
level the
FAT. of
The units
Hughes and
mainly
concerned
verifying are RADVS. the met.
compatibility This
checking a
functional of which first
requirements 6 concern is phase data
is accomplished
through
sequence
phases, The name and cised
phase this
termed yields for tests. Solar
Initial
System
Checkout of 4
(ISCO)
Test
Phase. of
As
the
implies, gives are
initial
verification The the next
compatibility phases in which
subsystems RADVS is exer-
reference environmental
future These Thermal Test
phases. are Vacuum Phase.
Mission
Sequence/Electromagnetic Vibration performance (AFETR) 5-5 includes Test for (VIB), and
Interference Vernier tests Engine are
(MS/EMI), Vibration
(STV)
Functional, set Range Table of
(VEV) the
Finally, Test
verification Phase. reference. ISCO to
performed from of and this any
during Appendix table, ambient
Airforce F has been
Eastern compiled
Information For phase note purposes tests is that
into
easier both the
the
"lab
ambient" tests
test for are
listing
readiness the VIB
other listed
phases. in in the
Another "vibr.
feature survival" all
only
tests tests
within in other of
phase usually
column, tests
although offer
phases
follow;
particular,
"prelaunch"
verification
survival.
5-11
Table 5-3.
Outline of VendorTests on Flight Units
UNIT
CHARACTERISTIC TESTED
UNITVERIUNIT FICATION ACCEPTANCE AMB. TEMP. A_m I VTR TVMp.
X X X X X X X X X X X X X X X X X X X X X X
KPSM Amplitude Modulation Other spurious outputs RAklystron rate Output powers Output frequencies Blanking signal amplitude, width, risetime Powerconsumption Warm-up time HVtime delay Klystron supply voltages, regulation, ripple Modulation inhibit circuit (for test use) R/T Antennapatterns Noise figure Preamp gain & phasebalance Preamp gain selection accuracy Preamp passbandshape& gain VSWRt XMT RCV a & flanges Preamp microphonics Insertion loss of special test horns Microwaveisolation betweenfeeds
Reflectivity-output Signal-tracking Signal-tracking Signal-tracking Response Analog Analog Analog Range Cross time output output output mark coupled accuracy, ripple, accuracy, sidelobe linearity noise (sine linearity logic (sine input) (doppler spectrum) input) calibration thresholds thresholds thresholds (sine (sine input) plus noise) spectrum)
X X X X X X X
X X X X X X
X X
SDC
X X X X X X X X X X X X X X X X X X X
(doppler
accuracies
Power consumption LVPS outputs & ripple
5-12
Table 5-4.
Outline of VendorSystemFlight AcceptanceTest
CHARACTERISTICS TESTED RAklystron sweeprate XMTR powers XMTR frequencies Preamp gain selection accuracy Reflectivity-output Signal tracking threshold (DVS) Signal tracking threshold (RA) Acquisition time (to RODVS) Acquisition time (to RORA) Analog output accuracy, linearity (velocities) Analog output accuracy, linearity (altitude) Analog output noise & ripple (at S/C filter output) i000 ft range mark accuracy 14 ft range mark accuracy Reliable operation indicating circuit operation Cross-coupledsidelobe logic operation (sine input) Logic signal amplitudes Analog transients due to preampgain switching Delay time from power-onto ROsignals Negative doppler rejection Warm time up Mechanical test & inspection Powerconsumption Thermalsensor integrity
STCNUMBERS --- (for reference) --- (for reference) --- (for reference) i, 4, 6, 8 2, 4, 6 i, 4, 6 2, 4, 6 i, 2, 4, 5, 6, 7, 8, 9, lO 2, 4, 5, 6, 7, 8, 9, i0 I0 7 i0 combination ii 7 6 5 3 i, 5
5-13
Table 5-5.
Listing of the Buyer Flight AcceptanceTests
Operating Conditions
,-4
CHARACTERISTICS
TESTED
•_ 0 -_
> • _ > m
HAC _= _ (see
TEST NUMBER
REQ.
AppendixF)
C
_
1,
Ranging Ranging
accuracy accuracy
(waveguide (freespace width
simulator) simulator) (high & low)
X X X X outputs X X X freq. X X X X X X X X X X X X X X X X range X X X X X X X X X X X X X X X X X X X X X X X X X
X X X X X X X X X X X X
RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA RA
135-i 136-1 116-1 107/108-1 105/106-1 133/134-1 111-1/122-1 111-2 122-I 122-2 122-3 109-1 124 112-1, -2 112-3/104-2 125/126-1 114/115-1,-2 114/115-3 102/103-1,-3 102/103-2 129-1 123-1 127-i 121-1 i01-I 104-1 i17_120-i,-2 130-1 132-1
2. 3. 4. 5. 6. 7. 8. 9. I0. ii. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29.
RA klystron XMTR output XMTR Preamp Preamp Preamp output
deviation power frequency
output noise level & spurious gain state logic & accuracy gain state signal false output at calibration accuracy in quiescent
Reflectivity-output Reflectivity-output Reflectivity-output
one
& repeatability state (sensitivity)
Signal-tracking thresholds Acquisition time Analog False Analog Range output lock accuracy susceptibility
and
analog
zero
accuracy
X X X
X X X X IX X X X X X X X X X X X
output noise mark accuracies false-lock circuit circuit
& ripple susceptibility logic & delay false outputs logic
Range mark Reliability Reliability Cross-coupled Waveguide Waveguide Warmup Power Unit
sidelobe
leakage integrity grounding time from primary over power input by TM by TM to TX voltage
consumption temperatures
Tracker-lock
indication
indicated
Negative velocity Range mark lockout
and range rate rejection until 3.7 sec. after BO
NOTES I. 2. 3. 4.
: About About 600-900 1700 feet feet. equivalent free space distance. level (both below deviation acquisition. modes)
Measured with Test conducted
the simulated return signal with the KPSM undeviated.
5-14
VI. EVALUATION OFPRESENT PROGRAM TEST A. INTRODUCTION The definitions of various portions of the programsare reiterated below to help avoid possible misinterpretations: 1. Unit Verification Tests are performed on all flight units by the vendor,
Ryan, 2. the prior Unit to the acceptance Acceptance cognizance II and III. Acceptance under I. Tests and (FAT) Cape are performed as on all in flight Hughes systems report by . Tests (FAT) of the are performed Hughes, on as all flight in tests detailed (FAT) the in Ryan are report 51765-2B on all as [60]. units Ryan by (Fli_ht) Ryan, under parts S_stem vendor, Tests of performed Hughes, flight in
vendor,
buyer,
detailed
report 3. systems Ryan
51765-2B, Vendor by the
(Flight) Ryan, part
cognizance
buyer,
detailed
report 4.
51765-2B,
Buyer-Flisht Hughes, [31]. at
Acceptance E1Segundo
the
buyer,
Kennedy
detailed
3023926A 5. tests 6. performed 7. formed
"Flight-Readiness performed Type on of flight
Tests" units (or systems is a
is or
a name systems; Tests
used tests (TAT_ for this
in
this listed and
report above
to are
encompass in this
all category. were
Acceptance units or Tests" of the
Approval) not name
Developmental (see to Section
Tests V.C). all
intended used in
flight report program.
"Special outside
encompass
tests
per-
flight-readiness
testing
B.
COMPARISON i. The
OF
TEST
SPECIFICATIONS
WITH
MISSION
REQUIREMENTS
Introduction purpose exist The with of this section RADVS _ is to determine adequately whether environmental The study or functional of various phases must compared
conditions two test of parts: phases
for first the no
which
not
tested.
consists the
compares actual
environmental to be to
conditions encountered functional to be
simulated during
during the
environment is the given
various RADVS are
the
mission; In test main the
consideration part,
requirements by RADVS
satisfy. with the The in
second
operations
performed
requirements. reference for test specifications in Appendix and Actual F. Environments in Section Ill.B.), phase. the various is the buyer FAT, which is outlined
Section 2. On
V.D.3.
and of of
detailed Simulated mission be
Comparison the of basis
the can
profile with
(sketched the
parts
the
mission
compared
appropriate
test
6-1
(a) Pre-Launch(PL) Phase In this phase, the rm_inenvironmental condition RADVS to withstand is the has EMIat the launch pad. The S/C in the MS/EMI est phase, sequencethree, goes t through a real-time simulated flight during which it is commanded through all modes of operation. Therefore, the survival of RADVS EMI in the PL and in the subto sequent launch phase is automatically checked. According to the HACtest specification, test levels are equal to or greater than those expected from all sourcesexcept the Centaur C-bandradar transponder. This is of no great consequence RADVS, to however, becausetests with the S/C telecommunicationstransmitter are at higher powerdensity and nearly the samefrequency. Furthermore, RADVSontains no pyrotechnic devices or other components c that might fail due to low-level RFheating. (b) Boost Phase Static acceleration and acoustic environmentsexpected during boost are not simulated in tests. The first of these is discussed in view of descent condition in a later section. The effect of the latter, acoustic pressure, during nonoperating conditions is expected to be less severe than vibration becauseof attenuation by the shroudand by the long propagation distance from the source. Also, the T-2N tmstindicates that nonoperatingsurvival of acoustic environmentsis no great problem. Boost vibration levels are expected to exceedthose of the VIB phaseof the buyer FAT. It appears, though, that the vendor unit acceptancetests are sufficient ; a direct comparisoncannot be madebecauseof the unknown effects of structural resonances. (c) Transit Phase During the transit phase, the most severeenvironmental conditions RADVS must withstand are related to the combination of solar radiation and vacuum. Comparison of actual and test environmentsis as follows:
Parame ter Actual -460°F 10 -12 flux: noted should the have 130 little of in RADVS the torr Expected - 300°F 5 x 130 10 -6 w/ft 2 torr (variable) S imula ted
Temperature Pressure: Incident The differences
of
background:
w/ft 2 effect is Van for on the
temperature checked. are not
reached
during
transit;
therefore,
survival expected be
sufficiently Allen this belt case
Radiation A special test
conditions (TAT)
imposed
in
testing. is
should
sufficient
because
susceptibility
6,2
very unlikely to vary among systemsof the samedesign. (Sucha test appears to have been conductedwith the T-21 vehicle, hut details are lacking in the available documents.) (d) DescentPhase The shockand constant acceleration causedby retro-rocket ignition and burning are not simulated in test. The shock environmentdoes not need to be considered separately becausethe rise time involved is slow compared the response times to of any RADVSomponents. Stati_ acceleration is important, however, becauseit c stresses every component nd connection to a high degree. It is also a factor dura ing boost, as mentioned, but the level during descent is about twice as high. Furthermore, RADVS required to operate during descent. is The expectedwidebandvibration level due to all vernier engines is I0 pounds rms, which is muchless than the total input of 60 poundsspecified in the buyer FATVEVphase. Relative to the vernier engine level alone, therefore, the VEV phaseovertests by a factor of 6. For a typical S/Cweight during VEVof 650 pounds, the correspondingacceleration level (roughly) is 60/650_ 0.I g-rms. Since this closely compares with the 0.2 g level expected during retro burning, the VEVphase probably yields a sufficient test of wideband vibration during descent. No tests are ever conductedon flight systemsin which RADVSperation during o sinusoidal vibration is checked. If the HAC_ environmental specification ([5], Section 3.2.3.4) is realistic, then such a test should be added. An easy place would be in the buyer FATVIB phase, where levels are near those expectedduring descent. Temperature and pressure are essentially the sameat the beginning of descent as during transit. After turn-on, RADVS temperaturesrise. This condition is realistically tested in the STV-TD phaseof buyer FAT. 3. Comparison of Test and Actual Functional Requirements
The JPL Surveyor [1,23] of System are Specification written spectral computation spectral in terms width, of width (No. of and 30240) functional power for is and HAC procurement and From in detail these, specifications requirements must and The be the requirements. signals used
frequency, The of of
simulations
determined. determination
frequencies is easily
usually
straight-forward, to usable accuracy. of test
approximated involves enforceable.
determination the
power,
although are
straight-forward, not strictly
the
estimation
unknowns; power
results, must be are
therefore, examined. based on power, the
Nevertheless,
levels
Computations (a)
following 31.8
factors: dbm
transmitter
DVS:
6-3
(b) transmitted power, RA: 24.0 dbm (c) antenna gain (one way), both: 28.0 db (d) minimum Muhlemaneflection coefficient: r -7.1 db Since spread spectra are not generally used in tests, the spectral spreading loss must also be computed. This is accomplishedby assumingthe spectra to have a Gaussianshapeand the filters to have rectangular passbands with widths: (a) for DVSbefore burnout: 3 kHz (b) for DVS after burnout: 600Hz (c) for RAbefore deviation signal: 4 kHz The 3 db width of the assumed signal spectrumis 2V Af = _- (_8) sinG, (6-i)
whereV is the velocity magnitude,X is the free space signal wavelength, _8 is the two-wayantennabeamwidth,and @is the angle betweenbeamcenterline and velocity vector [68]. Representative figures for the angle, e, can be obtained by assumingan angle of 45° betweenroll axis and lunar vertical, and 44° between the velocity vector and vertical at start of retro-fire. If the initial velocity is 8,800 fps and if the S/C retains its attitude relative to the lunar vertical throughout retrofire, then results for a beamat the worst roll angle are as shown in Fig. 6-1. (SeeFig. 6-2 for relationships assumed.) Initial misalignments of velocity and roll axis a few times greater than the i ° assumed Fig. 6-1 results for in little changefor velocities aboveabout 750 fps. Below 750 fps, the change would be noticeable but not great. a. DVS Beam Power Oneof the worst conditions of available poweroccurs whenthe return power is lowest and the spectrumis widest. Themaximum range for a beamoccurs when the vehicle is at the maximum operating slant range of 50 kft and its attitude with respect to the lunar vertical is 45° [i]. Theworst-case beamis then at an angle of 70 with the lunar vertical. ° The return powerfor this beamis computed as follows [50, and HACIDC 2253.3/359]: . &8 Equation 6-1 is a valid approximation for 8 _-_ and _8 less than about 15_
6-4
L_ m °
4.J °,.4 0
4_ o 0 I _J tf_ O4 > _J >
4-I _J
O0 c_
_0 0 C'4 _cO ,-4 4-J m -_ 0 ,--4 m I:L ¢_
0
4n
_J ,.-4 I..4
0 0 m O_ 4.J r4-J C-4 •,-4 _D
L_ 0
e_
u3
,.--i
I
0
I
0 0
I
0 0 0
I
0 0
I
0 0
I
0
[._
ao_aOA
X_TaOlOA
pue
six v
lloN
uaaa_a_
al_u
v
I
0 0 tf_ cq
I
0 0 0
I
0 0 tm
i
0 0 0
I
0 0 h_ 0
I
ZH
-
sau_od
qp
_
oa
RaPTM
I_aaaadg
-
JU
6-5
2 P tGX --> 2 (4_) 2 [50 kft cos 45 ° sec 700] -2 +12.1 dbm
--> --> 70 ° -->
-I00.3 -7.1 -13.2
db db db
Muhleman Muhleman
reflectivity reflectivity
coefficient factor at
Received
power
at
50
kft,
45 ° attitude
=
-108.5
dbm
_--
Misalignment / Plotted
Angle 6-1 Direction Worst Case Beam of DVS
in Fig. _ "_
_ Total Velocity (Abscissa in _ \ _
2_5o _-_
_
.......
Initial i° Velocity ._I ftP_th w _ _-_°ir:8/nme0n 0 Roll Axis
I
Gravity Component 5.3t =
Fig.
6-2.
Relationships
assumed
to
computed
curves
of
Fig.
6-1.
Fig. Before This
6-1
shows
that the
the
worst
case velocity
of
spectral at of which about for
spreading the 1.1 this DVS db. case is
occurs to After (with (line)
for
maximum is 3,000 the
velocity. fps. maximum pass-
burnout, yields a
maximum
operate burnout, the
spectral is i.i at
spreading 850 db. fps.
loss The
velocity band) require is
requirement the same,
loss tests
narrower
Therefore, dbm.
with
narrow
spectra
should
operation
-109.6
Tile situation -111.4 The dbm in the after would problem of
before vendor
burnout and buyer is
is simulated tests, simulated the at beam from [I,
in
STC
i at
levels
of
-106
dbm H andF
and ). of
respectively completely, 3,000 fps and
(see but 850 where
Appendices no fps real
condition
burnout be expected condition The vary at
not
difference
performance Another causes ability then loss is
between occurs
simulations. roll-off which at 34 acquisition kft, and
low
frequencies of 50
preamp for fps
power.
minimum
components 62 fps at
velocity to 29.6
required constant
linearly 29.6 fps
kft
remain
Section
4.6.3.1.7.2].
Since
spreading
losses
6-6
are negligible here, the representative powersto be simulated are (at 70° angle of incidence): Vb eam 62 38 34.4 29.6 fps fps fps fps R 50 40 38.5 34 kft kft kft kft Pr -108.5 -106.6 -106.2 -105.1 dbm dbm dbm dbm
(STC8)
Theworst situation is the -105.1 dbmlevel at 34 kft becausethe preamproll off is about 12 db/octave, while the gain due to range reduction is only 6 db/ octave. The doppler frequency for 29.6 fps is about 800Hz. The closest test condition is STC8, which hasbeamfrequencies at 930Hz. Themaximum range at which this frequency must be acquired is 38.5 kft° Since the difference in preampgain between930Hz and 800Hz is 3.8 db while the difference in altitude is only I.i db, the level for STC8 should be -108.8 dbmin order to check the worst case due to preamproll-off. This is to be compared with -104 dbmand -103 dbm for the vendor and buyer test specifications, respectively. b. RABeam Power Fig. 6-1 showsthat the maximum spreading of the RAreturn spectrum(due to doppler shift) remains somewhat less than the wi_ebandacquisition bandwidth. Consequently, it need not be considered. The high altitude case is computed follows: as
PtGX 2 _> -87.4 dbm
2(4_)2(40kft)2
Muhleman Muhleman
reflectivity reflectivity
coefficient factor at 45 °
_> -->
-7.1 12.1
db db
Received
power
at
40
kft,
45 ° attitude
=
-106.6
dbm
This in
figure the vendor
is
to and
be
compared test
with
values
in STC
2, which
are
-104
dbm
and
-113.3
dbm
buyer
specifications,
respectively.
6-7
The worst case of low frequency acquisition occurs whenboth range and rollaxis velocity are minimal. Specifications require operation at roll-axis velocities downto +i fps [i, Section 4.6.4.1.7]. At a 1,000 ft range, the return power would be -74.6 dbm. No test condition approachesthis combination of range, velocity, and power; a morerealistic check is madein STC7 with Vz = i00 fps, however. c. Returns from Retro-tankage The relative velocity betweenS/Cand ejected retro-tankage can be computed for a numberof vernier thrust profiles if the tankage is assumed be in freeto fall. For this case, the mainproblem is the assignmentof powerdensity levels in possible situations involving near field an_ or minor lobe structures. The effort of such an analysis, however,would not be justified becauseof the doubtfulness of the free-fall assumption. Onereason for questioning this assumptionis becauseof the momentary unlock of beam3 during the descent of Surveyor i. (If the tankagehad been in freefall, the chanceof breaking a DVS beam prior to appreciable attitude correction would havebeen virtually zero.) The fact that unlock occurred so soonafter retro eject makesit appear that the two events are correlated. However,quantization of the telemetered data seems preclude completeknowledgeof what happened to and an analysis of howit happened. For example, if retro entry into the beam 3 did indeed causethe unlock through shadowingor gain-state switching (which might have beenmissedin the telemetered signal), then howdid the retro-tankage enter the beamso shortly after eject (a matter of about two seconds). This might be explained if the retro engine thrust was still "tailing" off. For such a condition, it appears that computationof a velocity-power profile for retro signals into a given beam would be very difficult, and probably would have to be of a MonteCarlo type. The foregoing discussion showsthat the adequacyof present flight-readiness tests cannot be meaningfully evaluated from the available information. Consequently, the situation is reconsidered in view of the special test programin Sections VI.B.3 and VII. Pertinent tests in the current flight-readiness program are listed below for reference.
Vendor DVS: -50 RA: -59 dbm -3.5 fps (ref: kHz, beam STC -113 Tests velocity 3) dbm (ref: STC 3) (-1.6 kHz), Buyer DVS: opening or less, -50 RA: less, receding -113 dbm Tests
velocity of 65 fps dbm or less (ref: RAI30-1) target or of 3.5 (ref: kHz or
less
RAI30-1)
6-8
4.
The
Summary
of
Comparisons discrepancies found in between Table mission other of requirements conflicts 6-1 is have and test
significant are
requirements been discussed VII,
collected discarded.
together
6-1;
already until
and where
Further are
discussion made.
Table
withheld
Section
recommendations
Table
6-1.
Listing
of
Significant and Test
Discrepancies Requirements
Between
Mission
Requirements
MISSION
REQUIREMENT
TESTING
DISCREPANCY
I.
Survive
Van
Allen
belt
I.
Details be need for test
of a in
the to
T-21
test test.
need the
rad ia t ion.
reviewed
determine
special flight
2.
Survive eration
static of
accel-
2o
No
readiness
boost;
program.
operate during static acceleration of retrofire. Operate tion during vibra-
3.
3o
Wideband performed;
vibration narrow spec.
tests band 224800
are vibrais
during
retro-fireo
tion per HAC not checked. 4. Operate on available all 4. Possible frequency simulated.
return signals
levels are
of not
low
return power for situations within specification. Operate retro in rocket
5°
presence tankage
of
5.
Possible
conditions
are
questionable.
separation.
6-9
C. COMPARISON OFPRESENT AND"DESIRABLE" PROGRAMS i. Introduction
Objective it with program analysis. (a) the might evaluation "desirable" have of the present generated its own, in it two program in is can be accomplished although to by the afford comparing "desirable" a thorough program of Section complete IV;
defects is
enough
Comparison Overall that all point
performed of listed the
steps; are compared performed. to determine adequacy of under the assumption
contents tests of
programs
are
adequately is reviewed
(b)
Each
comparison
meeting The IV in and IV, first V. The
requirements. mainly consists of juxtaposing of the tables and details behind and from Sections
step
second
requires of the VI.B. Test
examination test
assumptions in V,
developments of the
consideration in Section
actual
configurations
notice
comparisons 2. Step Tables Table
Flight-Readiness (a) 5-3, 6-2. of 5-4, the and
Programs program 4-3; the comparison consolidated recognized outlined as is handled effect in step is this (b) by overlaying in Their compari-
flight-readiness 5-5 on Table
presented display. of the
Apparent
inconsistencies requires the
are
easily
interpretation, son. Preliminary in almost previously such tests. all
however,
analysis
to
step
(b)
the
STEA tests, of
signal was
simulation examined. spectral G.) This
technique, The basic are must
which finding not be
is was
used that in
flight-readiness characteristics are given
assumed
proper
shape factor
fulfilled considered
(Details test
in Appendix
in determining The will be detailed
adequacy. completing in Section step VII, (b) is listed Test below. Some of these items
review upon
elaborated
Suggested
Modifications.
ENTRY NO.
DISCUSSION
I.
XMTR at
frequency low altitudes in in
coherence: and
Coherence
problems in tests
are the have
not STEA been
very
evident as being
completely G. No
disappear pertinent
technique, or are
discussed performed
2.
Appendix the
Surveyor
program. Unit desirable system level tests for in vibration and but temperacan be
XMTR ture waived
amplitude were in placed lieu
modulation: in of the
program tests.
convenience
thorough
6-i0
ENTRY NO.
3.
DISCUSSION
Other unit spurious testing sweep are outputs and will from be KPSM: This in item is adequately performance G, ranging sweep checked tests. non-linearity might STEA give tests in
implicit As noted
system
4a,
Klystron effects some do not
linearity:
in Appendix current tests are
altitude but
dependent. more extensive
The
tests
indication, check sweep
desirable.
linearity. sweep and rate: Rate measurement tests under because lab is included rate is in not the otherwise are not
4b.
Klystron desired indicated. necessary
average vibration The because power: no The in in
acceleration three is tests directly
extra rate
ambient by
conditions output unit
indicated is called for of
analog
accuracies.
5.
XMTR tests
output because
Measurement other gross
during
environmental operation program two extra lab is
indication test
proper in
klystron the test. present The
obtainable. can be waived tests
unit lieu the
temperature of the
missing temperature are
system program
ambient
6.
current
totally read can (in
redundant. on be a wave expected use), The meter to indicate have
XMTR little
output about
frequency: operation, decreased and
Average except analog
frequencies large changes accuracy sideband
concomitant power istic is levels, most
output spurious of STEA
actual
lowered characteraccuracy,
increased to shift in
generation. analog true in
sensitive
average
frequency, because be
output
completely
insensitive Consequently, EMI, amount. test of stray which The is is
simulations should to lab
propagation all environa
is missing. ments except
frequency not expected extra
checked
change ambient because program might
average tests it is no be
values are so
noticeable the
7,
current
redundant; simple. is the occur.
prelaunch
reasonable, In
though, the overall
Production placed on
fields:
requirement deleted with would at the
EMI that signal for
generation. action would
Therefore, be taken These
testing
assumption
8.
if noticeable tests were tests
problems desirable with
Blanking unit range level
characteristics: convenience test the only. effect testing
Acquisition of blanking.
a realistic
signal
will
9.
Antenna
patterns:
Present
appears
to match
the
"desired"
program.
6-11
ENTRY NO. 10.
DISCUSSION Transmitter-receiver leakage: This portion of the programis as intended. Tests with both antennason an assembled S/C still must be considered. Insertion Ioss/VSWR:Problemscould occur during different environments but they would appearas lower powerlevels or lower sensitivity. Therefore, the unit test is sufficient. Noise figure: The effects of noise figure normally showup in sensitivity measurements.A measurement during acceleration is desired, though, becausethis environmentprobably can be imposedonlyat the unit level. Preamp branchesgain and phasebalance: Present tests are sufficient becauseeffects are also indicated in analog accuracyand false lock measurements. Preampgain stability with time: This test is implicit in the frequent checking of reflectivity calibration and systemsensitivity. Preamp passband shape: Unit level tests are adequatebecausesensitivity tests at various frequencies accomplishsystemlevel checks. Preamp gain selection accuracy: If accuracyand sensitivity tests were run at manydifferent powerlevels, separate systemlevel gain selection tests would be superfluous. Since this probably won't be the case, environmental testing of this item should be complete. Spurious outputs from R/T units: The outputs, in themselves,are secondaryto their effects on false locks, analog accuracy, and sensitivities. Since these effects are to be checkedat the systemlevel, there is no need to check for spurious outputs beyondthe unit level. Reflectivity-output calibration (stability): essentially matchesthe desired one. The present test program
ii.
12.
13.
14. 15. 16.
17.
18a. 18b.
Reflectivity-output ripple: Nodirect specification of ripple exists. Since large values will be evident to the test operator whenmeasuring with the DVM,this phaseneed not be addedto the program.
6-12
ENTRY NO. 19.
DISCUSSION Tracker search range and rate: Theseitems might be covered in thorough acquisition tests at the systemlevel. Nevertheless, since they should be easy to perform and they indicate the time constant of the tracking loop integrator, their addition to the unit test programis reasonable. Signal-tracking threshold operation (sensitivity): Only the lab ambient unit verification tests use spread spectra in the present program. All other tests either apply sinusoids or no signal; the latter is the case in the systemenvironmental tests. Consequentlyfilter bandwidths, threshold circuit operation, klystron AMeffects, preampgain and phase balance, preamp gain selection operation, tracker search range, and blanking circuit operation are not checkedin systemenvironmental simulations. Operation times of DTC circuits: Thesedelay times are not explicit systemrequirements, but they appear to be necessary for proper operation with real signals. (No documented test exists in which they are checked.) Temperature might affect timing without being otherwise evident. Other environmentally induceddefects will be easily discernable in other tests. Delay time for filter BW changein RA: A delay betweendeviation signal and bandwidthchangeof the low pass filter in the RAtracker was evidently found necessary to help insure proper operation in real use. No documented test of this characteristic exists. A check for large variations with temperaturewould be reasonable. Analog output accuracy: The main discrepancy is the lack of a vibration test and an EMI test in the present program. Accuracy tests indicate whether spurious signals and/or noise tending to offset the center of the spectrumbeing tracked are present. Accuracyalso indicates the tracking loop gain value. The fact that realistic spectra are not used in most tests meansthat converter circuitry is not fully checkedfor response range. Analog output noise and ripple: Environmental tests are lacking except for the unit FATvibration test. All tests should be run with spread spectrumsimulation becausethis checks the tracking loop bandwidthand converter responserange. 6-13
20.
21.
22.
23a.
23b.
ENTRY NO. 24a.
Range of the mark mark accuracies: circuitry all to Range and
DISCUSSION
mark accuracy noise imposed. long as indicates is proper on operation inputs. tests being
shows
whether be as
appearing Extra lab
Therefore, do not seem
environments add much
should
ambient is
information
analog
accuracy
checked.
24b.
Range with would the
mark no
susceptibility input realistic in
to the
false present
mark:
These
tests Tests
are with
performed spread spectra on
signal
program. for checking
impose mark
conditions
environmental
effects
circuitry. circuit with might be the logic desired operations: program, to The except EMI. present for A the program EMI test; is in basic logic cir-
25.
Reliability agreement cuitry would
quite
susceptible of the logic
check
during
acceleration
test
integrity sidelobe
circuitry. operation: circuitry threshold Environmental are detectors Operation with two spread test of missing should the tests from be of the the cross-
26.
Cross-coupled coupled program. during the the
sidelobe The many
discrimination gates and
present
checked making
operating test
environments. should be tested The for
circuitry
frequency
spectra. programs time essentially during the agree. EMI
27. 28. 29.
Waveguide System test System be are is
assembly warmup time:
performance: No reason
checking
warmup
evident, power consumption: except, Extra perhaps, tests to help in tbe present that program connections appear to to STEA
superfluous correct.
assure
30.-
37.
Miscellaneous: checks in other at the tests tests: checking enough would rate
The vendor
first level.
four
tests
in
this were
group
are
fitting to be
assurance included
The
others
considered
already The
listed. used signal frequency apparent is in the present ranging The problems tests. tests generally used are
38.
Ranging excludes not due about further great to
procedure
blanking to make be no more
effectiveness. coherence than obtained, test STEA
distances evident. Some This
Problems
AM sweep
information topic will be
and when
linearity
however.
discussed
considering
modifications.
6-14
ENTRY NO. 39.
DISCUSSION Powersupply transits: Although sometests are run at the extremesof supply voltage levels, no checksare madeof transients effects.
3.
Each
Special area a. of
Test the
Programs special test programs Leakage concluded in is Tests that actual testing flight. of The test. plume effects on leakcompared separately as follows:
Transmitterof not
Receiver Section IV
The age is
discussion probably was performed S-8 tests
feasible, to be
except an S-8
favored This
earth-bound latter type of
alternative test was The
decided using
upsidedown vehicle. engine
vibration
the
s_mulated used in
vernier the
vibration test It during coincide data the seemed trackers would
levels either
based be
on based
available on S/C-I retro
information. flight fire data should Processing cedure. or
Levels at
"desired" recent data.
least be
on more if
appears this with to
that phase the lack avoided
levels
during
also of
simulated S-8 data
operation to
is desired. "desired" test proof also the
the
appears of S-8 by
Subsequent
analysis
test which
determination false lock
desensitization were not readily the be
caused. available.
Margins
Finally, ment should b. The in T-2N the
differences
between
the
test
equipment
and
present
flight
equip-
reviewed. Tests test A but small these described amount tests in of were Section additional of main IV appears knowledge value to the to have been fulfilled from the and
Flight
desired T-2H
flight
program. tests,
was
gained
descent engine c. Thorough
flight
control
vernier
systems. Signal in this Tests area was specified is evident in the "desirable" the available test program.
Interfering testing in the tests CCSL flights single buyer RA
No The
such only
effort related (i)
present are: circuitry over negative system 130-1).
program
from
documentation.
The in
was water.
caused
to
operate
in
the
T-2H
tests
(2)
A
-doppler FAT (see
simulation Appendix E,
is
performed
in vendor F,
and test
STC-3;
Appendix
6=15
_ "_" I
,_
-_q._ I_
"qmv qe_I
o 0
*_ m 0
>
°°}
c
"q'"v
q e
"_q!A
.,-_ _" _ C >
\
" qm V
II II
qe]
I-i m
• _ _-
0 .,_
0
_
m
e.-, Dq
C
_
C
I-4
_ta
_D
II
II
II
[---.
'-'d
co
6-16
qauneI°a8 Ie^g^'xns IN_
_, = =a
¢.t'] 0 a...a ,<
"aqt. " qmv
A
,--
qex
'qwv
qel
a.J .,q
Qa >
[--4
O_
4.J 0
¢',1 I
,...4
6-17
Unless further tests are uncovered, the present programmust be regarded as deficient in this area. d. Environmental Overtesting Although environmental overtesting took place in both vendor and buyer TAT, no statistical significance appearsto be ascribable to the results. Levels used seemto be based on estimated flight conditions rather than being varied to determine operational dependences.The amountof instrumentation seemsto have been minimal for assumingthat complete failure would be recognized. Fatigue-type testing is lacking in the present programaccording to available documentation. D. DOCUMENTATION ADEQUACY At all levels of testing (unit verification, unit flight acceptance, and systemflight acceptance), it appearsthat the testing requirementsare clearly defined and documented. However,the test proceduresand equipmentsetups are not oomplete_ documented. With regard to systemstests_ it appearsthat this shortcomingis being remedied(Hughesis in the process of preparing test procedures). A more serious deficiency appears to be in the documentationof proceduresand equipmentsetups for the unit tests; this documentationis believed to be important becausethe unit test equipmentand setups are not consolidated into permanent assembliesas completely as equipment for systems tests.
E. TESTING It One has CONSISTENCY been found that to be testing in the at the various of levels is radar in generally consistent. sensi-
exception The
appears system over
method to the
specifying is given
acquisition terms open model of
tivity. performance possible of
specification of as entry choice
vendor This lunar
altitude of
a range such
angles. of the
method
leaves
a number and
questions, and of
reflectivity precise doppler
extremes be from Since the early a
attitude curve
velocity acquisition of the appear
during
entry.
A more versus and state be
specification frequency mission it and
would
simple current
sensitivity lunar to surface be in a
derived requirements. that at an
knowledge numbers method
anticipated of flux,
sensitivity suggested date.
is believed documented
of
specification
should
employed
6-18
VII.
SUGGESTED MODIFICATIONS TEST
A. FLIGHT-READINESS PROGRAM I. Changes in Content
Suggested on the modifications of Section to the VI. flight-readiness In particular, of the test Table this table program 6-1 are based a mainly of discussions factors. show are which provides recast A listing Table of these
important 7-1 to
The actions
significant are
entries in for
are cases.
into
suggested below;
specific
number VI
results be
discussed
further
additional
comments,
Section
should
consulted° a. In the Unit present while performed listed in Constant program operating during the Acceleration the is the only due Test appreciable to wideband mechanical random stress imposed No upon
the
system are
vibration. or constant
operating
tests
narrow-band
vibration
acceleration [5, Sections stress 3.2.3.3 of each
conditions and 3.2.3.4].
environmental of be which based in the
conditions condition on the
specification creates the
Determination connection and the the must the
greatest
component the is input not
or
mechanical Although characteristic forces and
transfer such
functions
between
points
element of
question. response input and and of
specific be
information reasoned are as made
available, Since
nature
might given
follows: up of many
paths of
between different to the have
the
any the
element
components would be same
sizes poles
materials, zeros
associated over a wide not
transfer frequency normally characterand on the
functions range. be great At
expected time,
spread constants
the
variation
damping
would with
for
solid do
components. not be have
Consequently, sharply defined
because resonances, which
functions both are
the
istics wideband their retro levels
described inputs total descent specified power
narrow-band dependent
would
expected When
to have applying
effects this
mainly
levels.
reasoning to be
to consideration the most severe;
of
phase, are: constant
constant
acceleration
appears
the
acceleration vibration vibration must it be
@ @ @
10.8 1.4 0o2
g --> g rms g rms --> -->
l17g 2 2 2g ms 2 ms statements will be of power
sinusoidal wideband (Driving levels The might For point can be impedances made, but
0.04g
considered that the a
before large
definite difference
is unlikely shows that
noted.)
foregoing induce
discussion the
constant stress the on
acceleration components
environment and connections. is suggested.
easily this
greatest of such
mechanical a test to
reason
addition
flight-readiness
program
7-1
qaungIaa8
II_A_A=ns
IN_
"g
GD O
_, o
<>" =_
• qmv
.iaaag
"-I¢ "q_'I "v= I _
co tO r/) .,-i ,-4 _D
"ID "El
"E
0
I O _D >
IRa "dmo_ •aq_ D O O • ._ _ _.J aJ ii db (narrow-band mode) and be allowed to cause suppression of weaker signals unlikely signal assume about assume 6 db margin the that any non-linearities cause signal significant were suppression capabilities in the be unintentional, if above, that 6 db that is could sidelobe the IF greater the allowed and sidelobe designed might be
18 db in this signal into quite this
extremely
present
suppression
derived
effective. For example, limits heavily on signals the above signal would more foregoing noise discussion, and that a
amplifier rarely than noise. To acquisition order to signals. in CCSL
limits on noise, but be consistent with threshold for CCSL Thus, is I0 db short-term signals
tracker
account
fluctuations
between
mainbeam above the and 14 db
be prevented from than 7 db (wideband be This requirement seen from the burnout and center to CCSL
rising mode)
acquisition (narrow-band
threshold provided they are mode) below the mainlobe
signal. as can persion between IF widened, always
appears more data inTab_sC-i velocity signals include as both can
reasonable andC-2. cause as In signals, large
than those derived previously, In Part IV it is shown that disof doppler in would stronger IF frequencies order have signal for to tbP be would Therefore, bandwidth way in the the
of
lateral to
differences 8kHz. its this
mainbeam or be its
amplifier
bandwidth available
frequency produce
shifted. sidelobe
suppression
amplifier.
Finally, similar manner sidelobe-signal
the to
IF the
mixer mixers
which in can
precedes the be expected.
the
tracker
filter modulator,
will and
operate no
in
a
single-sideband
effective
suppression
C-5
B.
The switching basically be the this tion greater signal is
Gain-State second of as
Switching for the range a that sidelobe ratio switch both signal before same that as value will of that dynamic the of to signal will now the the mainbeam next cannot always suppression be discussed. signal lower to acquisition gain to rises would that RADVS, that simultaneously manage signal signal out For for argued the switch above always this the this (i.e., The sidelobe state? mainbeam question signal If fall the the to threshold signal is always and this within preacquisigain signal range could control mode, threshold, to all con-
possibility will dynamic causes we the see
preamplifier follows: the which
gain-state)
than level
between
tracker
dition amplifier states, suppress spread expressed
satisfied, range; lower thus to a
signals would the range it now 33 in db it be
dynamic threshold.
stronger the
gain-state be assured upon,
sidelobe stronger turns shown. be for
Because weaker be has ratio
applies
approximately
it would the cannot above the
adequately dynamic of spread
signal. a maximum etc.)J45]. in
Unfortunately,
requisite
depended
(allowing might
dispersions signal
gain-state be reduced,
switches,
Although
(C_)shows
order
mainbeam
the gain-state, after burnout).
the value of SNRtracke r must If the acquisition threshold in
be well above 22 db (narrow-band is i0 db above noise, if a 6 db signal fluctuations, the gain-state trip be be the reduced required We, mainbeam signal and to and would and if another power relative about this would therefore, control will the 21 db. would not of often
margin is allowed to account for short-term 3 db is allowed to account for variations preamplifier However, undesirable sufficient answer less trip IV. that than value gain-state, if noise, this for to CCSL the (33 were the 33 of db more range could done, adequate the for ratio gain-states
to be
probably
a number signal
reasons.
Furthermore, of be CCSL cannot the
improvement by
be the the be
ensure
suppression of mainbeam between RADVS, OF
signals. to sidelobe threshold 21 db for
have
suppression range the TO
obtained
because dynamic db
signal acquisition and CCSL about
gain-state system).
present
a modified
SUGGESTED A. The
APPROACH of
CORRECTION
SIGNAL
PROBLEM
Description study
Technique above has the resulted that the problem, in a method and suggestion described require for below only correcting would minor have the CCSL
described It would for
signal siderable
problems. promise
appear
con-
solving
would
modification
to existing Fig. C-I
circuitry. shows a block diagram of the suggested solution. Actually, tracker. as shown this is
just a simplified modification is Also, just the the circuit 90 db
block diagram of a portion that the bandpass filter has shown as inFig.C-I for the should present cut-off in the be
of the present been widened, used for The all
The only in Fig. C-2° rather of the than band-
gain-states, cut-off
state, equal mode by the
system. of the wideband
lower
pass filter is the narrow-band is in determined turn is RADVS by are
to the upper and 1500 Hz maximum by of in the This the order
tracker mode). between upper
filter (B t = 300 Hz in Its upper cut-off frequency all three to DVS be beams, must of this which be freencountered
doppler maximum value planned to C-3,
differences lateral-vdocity of required missions;
determined operation. analysis below given
dispersion cut-off
during quency
frequency estimates capability.
determined
preliminary the circuit's is
illustrate the case
Referring
to Figs. C-2and
illustrated
for
the
CCSL
signal equal signal
present in the low-pass tracker filter, contained in the bandpass filter. The gain. little For low values on the of PM average and PCCSL of effect values
while the correct detector circuits (i.e., E b and low E t. SNR's) However,
mainbeam signal is are assumed to have these as signals the power will in
have each
C-6
Z r I i--i
_
0
_v
E-_
0
c_ c_ °i-4
_j C_
_o
7. r_ P-i E-_ • v _r_ D.o .kJ o _j
_j
t
0 r_
0 ,r-i
6
i-,4
i
o _J
C-7
RESPONSE
(db) Mainbeam CCSL Signal ! | Signal
/
./ _ TRACKER Bb I I #
i I
I
I I I I
I % 1_
r------
BANDPASS
I I
FILTER
/
Fig° C-2.
B t
' \
curves of lowpass and bandpass
LOG
FREQUENCY
Response
filters
in
Fig.
C-I.
/
_0
>
No
Signal
,_ E b
No
Signal
E
I I
PCCSL 2N
O
I I
i --_ PM b i Q--PccsL
B
t
2NoB
Fig.
C-3.
Detector with
outputs as
(Fig. a
C-I)
versus
cross-coupled
sidelobe
power,
PM/PccsL
parameter.
C-8
exceeds It is <
the seen
noise that
contained _E b aver
in > Et
its aver
filter, for the PM
the > I/_2
average PCCSL output
output and of
rises for all
as
shown.
low-level ampli-
(SNR
I) CCSL
signals. for these amplifier
Therefore, cases. output
average
the
differential
fier is positive the differential
Allowing some threshold margin, cannot exceed the threshold for
VT, we see that all cases where
E t - _E b - V T _
0
(C-5)
where (C-5)is CCSL
VT
is
the the will > _E b
threshold circuit When have for all will
setting, not the mainbeam a minor practical
referred CCSL signal effect cases on
to is
the in PM
differential for the is which tracker above the the
amplifier inequality the filter, operation. tracker
input. in smaller In this
Therefore, signal
lock
signals the
satisfied.
only
acquisition
case, Et > threshold. B. The receiver
where
acquisition
Derivation circuit noise is
of
Capability will be analyzed (i.e., doppler in detail, assuming band). The for that the there case are of where no such the spikes spurious of
in Pig. C4 essentially
thermal
transmitter-receiver leakage in the signals will be discussed later. Of fundamental importance
effects
in analyzing
the
circuit
performance signals. between
is
the
band-
width, Bb, determined CCSL based lateral ponent ing on
required to contain both mainbeam by the maximum doppler-frequency This maximum angle of the of plane Bb frequency minimum 150 fps, of burnout and one will
and CCSL difference has velocity that three. and be
This bandwidth is the mainbeam and to a 3o of was maximum be 6.2 kHz, of comfollowoperafps, purpose value value the
signals. burnout passes
difference assuming
been of the
estimated 220 For This lateral
a 45 ° approach velocity through a value
velocity
beams
calculation,
= 8 kHz
assumed.
obtained
from Mr. R. Dibos, Hughes tional spread between the Assume the B t following = = Bb Bd = = = 300 1500 8000 4 Hz ½ (this
Aircraft Company, has an center doppler frequencies radar parameters: mode mode (N.B°) (W.B.) with
estimate of two
of the beams.
Hz, Hz, Hz,
narrowband wideband (from 40
discussion msec will response be shown
R.
Dibos, as for
Hughes) present system)
(for
time to
value
give
discrimination
capability It can be shown that for E b and E t may be expressed Eb _
PM/PCCSL d linear as
-- 6 db) time average values and variances
detector,
of
kI_PM
+ NoBb
Et
_
klVPccsL
+ N
o
B
t
_b 2 2 at
_
k I2 N ° _-'-v BbBd 2 kI ! No_/BtB
(C-6)
d
C-9
wherekI = gain constant N = noise density at input (assumed o uniform throughout bandsBt andBb). Thus, (C-5) can be rewritten Et C_b- VT+ Random Term_ 0 (C-7)
where the randomterm correspondsto the noise fluctuations on Et - C_E . b In the absenceof signals, V will causethe average values of the left side T of (C-7)o be negative. However,false-alarm locks can occur if the random t term goes sufficiently positive to overcome this averagenegative value. Although such false-alarms causeonly a pause in the acquisition search, it is desirable that they occur only infrequently. This can be ensuredby setting IEt(PccsL = 0) - O_b(P = 0) - VTI M (C-8) > >Vut 2 + _2_b2 which is achieved by setting (C-9) For the numbers given above V + 27.5 kI_o T U + 6.0 kl_o r
For low This a factor of 4 in this of inequality, the random the
>> 8.9 kI _o >> ii kl_o
false-alarm beyond rate 4_
(N.B.)
(c-lo)
(W.B.)
should be acceptably false-alarms).
(i.e., value
only values results in
component
would
cause
VT
= =
8 k I_ o 38 kI 0 _N-
(N.B.) (C-II) (W.B.)
Returning derive average CCSL lock and defined by the
now
to
the
general
condition
for
CCSL
lock
given
in
(C-5),
we
will
values no-lock. condition
of PCCSL and PM which In terms of average
define values
the threshold of E t and Eb,
condition between this threshold is
Et or kI_PccsL + NoBt
- °_b
- VT
=
0 (C-12)
- _
kI_PM
+ NoB b - V T
=
0
C-10
or
V
PCCSL NoBt
+
i
8 w_-t
__ i 2
PM " NoB t
Bb + Bt
(N.B.)
(c-13)
_CCSL o N_ t + i 38 x_ t _ I__PM 2 Bb (W.B.) o Bt "_-B--t +--
These radar
equations parameters
have been solved for various given above, and are plotted for PM = 0, acquisition PCCSL t_an the by above = in use the values sensitivity 0 and the is of assuming bandpass 5.6 db smaller signals 5. will
values of PCCSL/NoBt in Fig. C-4.
and
for
the
to the tion
Notice that the circuit's same as filter can be
of PCflNL/NoBt satisfying (expressed as a SNR). that (WoB.) the and mainbeam Fig. C_ 9.1 also db filter. Fig. C-4 shows shows the
(C-13) This is is that that,
correspond of course in the acquisisensithe will Thus, average, a reliably a The lower allow
setting rather of
signal (N.B.). latter.
tracker tivity mainbeam
the This on If we
sensitivity signals
circuit CCSL
improved 6 db
suppress
3 db margin to account suppress CCSL signals. trade-off above values between results may be C. A off inFig._l Use
for signal fluctuations, a The asymptotic ratio shown sensitivity about and the CCSL
signal ratio in Fig. C_is suppression desirable;
of 9 db 1/52 . may and be
acquisition
made.
show that _ = 1/2 is a better compromise.* of a Non-Linear more elaborate with
highest
somewhat
Attenuator circuit and can provide more of flexibility The type linear shown the in in the Fig. tradeC-5.
slightly is
between
acquisition replaced
sensitivity
CCSL
suppression.
attenuator
a non-linear
attenuator
_E b RI Eb _E b __ r Slope Z Slope R 2 i/_ 2
"77T
wl Fig.
r .
---
_yE b
C-5.
Attenuator
characteristic.
This while esting
circuit _2 is to
permits selected that
selection of 51 to to give the desired for _2 > 1 (i.e.,
maximize high-SNR one
the acquisition sensitivity, CCSL discrimination. It is signal may be used to suppress
intera
note
gain),
*For and 7.5 db
example, (N.B.)
5 and
= 1/3 CCSL
gives
values
of
acquisition capability of
sensitivity 9.5 db.
of
5 db
(W.B.)
discrimination
C-ll
10
6 db
I
I I
15
I
20 PCCSL NB ot (db) 25
I
I
30
-5
Acq. (Lock beam nal SNR is
Thres. of mainthis sigfilter) when
values in
tracker
Fig.
C-4.
Sidelobe Bb =
suppression B t = 300
capability Hz (N.B.),
of
circuit Hz apply
in
Fig.
C-I
for Hz,
8 kHz,
1500
(W.B.), when
Bd = 4 PCCSL and
and _i_= 1/2. Acquisition PM are interchanged and
thresholds PCCSL = 0.
C-12
stronger signal. As an exampleof the use of this circuit, for _. = 1/3 and '_2 = I/_2 , the circuit's acquisition sensitivity will be 5 db (W!B.) and 7.2 db (N.B.) and its CCSL discrimination capability for high SNRsignals will be 3 db. D.
The output Actually, band, ripple, components ponents are to be below which tible used the wider tive of the power the than spaced better employs to to provide RADVS filter. merits spurious in present in band would the the for be appear of such is Effects previous smoothly of etc. course, from Some of of Non-Thermal Noise applies as be will noise Components to for the case noise may where noise the arising pre-amplifier in in but the the power-supply will these also combe from there RF noise crystals. discussion distributed, there these
thermal leakage,
spurious components
components crystal be random,
arising
transmitter-receiver essentially periodic.
vibration, arising
which
are
Difficulties
most likely to the acquisition noise-derived when above and two the system It the
occur in the high-gain mode, because they are expected threshold for lower-gain modes. Therefore, any system thresholds a peak are such of from a part noise Notice The of the that major is general Actually, If the a this by in the sampled lesser case doppler in the band the band is RADVS between uses about suscepwhich and the a the are (for will the all (6 both be total noise in Fig. comlikely rather flucin I) relais spurious systems. appears
desentization described
noise-derived the
threshold.
present latter
circuit
difference that the the these statement
present
suggested
modification a broad question. thresholds.
is not is the
possible relevant only suggested
to make to one this
noise-derived that
frequency components band of
spacing
components a manner and If system case the the band. modified the the and RADVS
appears will
noise part
modification), represent component in
modified
system over factor
because
spurious system where argument
component is the in also
a single
spurious
predominates the Even component a greater one single component
superior; is attenuated band for
noise-derived if will several most applies
threshold ponents
in Fig. C-i
sampled the above
is narrower. band, the
simultaneously viewpoint is doppler band will smoothing be a given
sampled
predominate,
given
well. Another throughout the tuations be the smaller standard the which for
as follows: assume a bandpass filter and the detected output observed; in (normalized filter a the given than to for the is from average averaged the mean. detected one; over a narrower
is scanned general, the output) just an this
observed bandpass law the when less
will follows
waveform
interval--the
broader E. A
interval, of the
variations
Summary study of
Technique estimated
Capabilities values given in TablesC-landC-2 cases indicate on For show that to be that the capability
given in Fig. C-4 should be for the present RADVS system. case for would to be encountered, angles vertically of without 25 ° and downward, approach point
adequate Reference RADVS for a and
for all practical 45 data would restrictions angle signals of 45 ° . from
encountered, the worst would beam 2 receive have tn used bebe
imposed
roll-angle, this case, beam
roll CCSL
transmit
2 into
beams I and 3 could cause be avoided (within + i0 °)
trouble. for all
This particular missions, unless on
condition CCSL logic
will probably circuits are
tween these beams. This reference improve as the approach angle moves or toward 45°).
also indicates that either side of 25 °
the PM/PCCSL (i.e., toward
ratio should either 0°
One described the required re-lock lock-on
very
important even of case the
point in the
should event
be of a
made false ratio
regarding lock-up not being
the on
use the
of CCSL
the as
circuit due as to the soon
above: ratio on the (i.e.,
signal,
unlikely
required
signal
exceeded, from the that a be very
is exceeded the mainbeam signal. a continued
circuit will Therefore, will
cause un-lock the probability appears to
CCSL signal and serious false small.
lock-on)
occur
C-13
This tive power inferences obviously was V. done
brief ratio very by the OF
attempt of the the
to
place
bounds and of CCSL
on
the
doppler and is against
separation then false obviously to draw CCSL in
and signal
of
the
relais
mainbeam degree
signals, study
conclusive lock-on, such as order,
regarding Monte
protection detailed
incomplete. Carlo
A more
computation OCCURRENCE
described OF MAINBEAM
in [50]. AND CROSS-COUPLED
EFFECTS SIDELOBE When the
SIMULTANEOUS IN
SIGNALS lateral
TRACKER-FILTER components equal cases for
BANDWIDTH are quite small, the mainbeam signal and
velocity
CCSL signals will has been concerned to both exceed from be prevent We signals the measured almost exact when both now CCSL than are width.
have nearly with those signals from cases this by
doppler frequencies. for which the doppler falling the The mainbeam from that those within which large frequencies the the cases
The previous discussion separation is sufficient the tracker-filter enough will so usually and cases is bandthat are close p_er
simultaneously bandwidth.
consider within power pattern 16 db;
signal
a rather
factor; it appears are
previous ratio for which
discussion will in most one beam
characteristics notable toward this analysis of the the
greater
exceptions lunar
pointing An be example,
vertically analysis of
surface. problem are well is very difficult the and will For not
interference and two signals
attempted
here. the
Past spectra
experience
is very
helpful,
however.
separated,
illustration
inFi_C6 shows to be frequency responds bandwidth 30 is good Hz or referred in to this
that pre-discriminator modulated at the beat beat frequency, 7 cause against Hz). in FM the the
limiting action causes the resultant signal rate. Even if the discriminator bandwidth tracker of effect of two will this, on where pass the only tracker AFC which those signals output. circuit are beats within by effect extremely This provides its doppler the separated
(approximately more to its should as "capture"
Because receivers,
negligible weaker
discrimination
signals
simultaneously
present
discriminator the capture case,
bandwidth. effect will there still will be be present times when even the for overlapping, of the spread small to
Actually, spectra. In
this
however,
amplitude
signal exceeds that of the large signal. the discriminator output, with the result the weaker be biased two spectra Vx signal. slightly is
During these that the VCO
times, it will contribute is driven slightly toward the tracker the separation resulting The occur; on the
Thus, for such fluctuating signal inputs, away from the correct frequency. Because to be lateral velocity to the components, correct
VCO will of the in
proportional Vy will
errors
measuring
and
proportional
velocity. Vy.can nolse the
important the major velocity velocity
point is that no fixed off-set errors effects of the interfering signal will analog outputs, and to cause small components. For is acquired VI. the the relative no mainbeam level of signal. of the the
in measuring V x and be to increase the proportional to
errors
lateral
mainbeam
and
CCSL causing
signals tracker
to
be
incurred, once
there it has
essentially
danger
latter
signals
unlock,
CONCLUSIONS As a result been The of the review and analysis described above, the following conclu-
sions
have (I)
reached: cross-coupled RADVS sidelobe problem is a very serious one for the
present (2)
system.
There are no inherent suppression effects caused by circuit nonlinearities which would be effective to an appreciable degree. The unintentional presence of sufficient non-linearities to do
C-14
Loci of Sum Signal After
Limiting Before Limiting
Input Limiter
/
\
Mainbeam
j Signal, fdM
!
,
.._
¢3
l
0
/
'
,
Fig.
C-6.
lllustration of a large is separated effect of ca1 the
of signal signals to
limiting plus (i.e., a (before
action a small
on
the
resultant For well the
signal.
non-overlapping), limiting) at the center small that
change
non-syrmnetrical
spectrum to a sy_mnetriof the largethe of the larger
resultant spectrum. is do slightly occur
spectrum,
centered For
signal effect signal
fluctuating from
signals, shifts of the
different; center limiting.
resultant-spectrum
after
C-15
(3)
(4)
(5)
(6)
(7)
this would result in noticeable degradation in tracker sensitivity and/or in significant signal clipping. Any suppression of cross-coupled sidelobe signals obtained in this mannershould be explored thoroughly to insure that no performancedegradation occurs in other ways. Gain-state switching can causemainbeam signals to suppresscrosscoupled sidelobe signals below the acquisition threshold, only for powerratios of these signals exceeding33 db, the dynamic range betweengain states. Manycases will occur for which this ratio will not be exceeded,and therefore gain-state switching does not provide effective protection against false lock-on to CCSL signals. The solution to the cross-coupled sidelobe problemby restriction of roll angle is not applicable to all missions. In fact, the technique _ppearsto be most effective for lunar descentsnear 25° from vertical (such as Mission B), and a rather narrow margin appears for this case_5]. The roll angle selected for Mission B doesnot ensure that cross-coupled sidelobe lock-up will not occur, but does give low and approximately equal probabilities for false lock-up on beamsi and 2. In order to eliminate the cross-coupled sidelobe problementirely by antenna improvement, nd not imposeroll-angle restrictions, a each receive antennamust have sidelobes in each of the other two mainbeam directions which are at least 46 db below the mainlobe. This can be inferred from the results in reference 45for 25° approachangle, which is believed to imposeabout the worst requirement. Sucha specification on the antennaswould probably still meanthat certain roll angles for the 25 approachwould have to ° be avoided, in order to avoid having any DVS beampointing within about 5° of lunar vertical. If a partial solution is adopted of rotating the antenna (beams 2 and 3) 180 measurementshould be madeto insure that all patterns °, s relevant to the cross-coupled-lobe problemare measuredor that the cross-coupled product is measured directly. Evenwith this solution, the data contained in reference 63 and the analysis in reference showsthat difficulty could be encounteredfor the 25 approach ° over appreciable intervals of roll angle, assumingno RADVS restrictions on this angle are imposed. Thus, for this solution, each mission mustbe analyzedcarefully to ensure that no serious CCSL problemexists. Reference63contains all the necessarydata on the antennapatterns of S/N i. Limited data on S/N I0 showsgoodrepeatability on the -27 db sidelobe of antenna2 in the mainbeam direction of antenna I. However,the sidelobe of antennai in the mainbeam direction of antenna2, being at a lower level, did not repeat (values are -37 and -46 db). Because sidelobes at this lower level can influence the cross-coupled sidelobe problem, measurements should be madeon eachantennain order to determine the level of the following receive-antennasidelobes in the direction of the indicated transmit mainbeams:
C-16
Receive
Antenna 1 2 i 3 2 3
Transmit 2 i 3 1 3 2
Antenna
These results should then be used to sidelobe problems for each mission.
evaluate
cross-coupled
(8)
It
appears of
that antennas
the
use
of
CCSL the
logic
circuits If
between this is
all done, or care more
pairs should pairs (9) A
can
solve
problem.
be taken that is not allowed method in a the circuit attenuator signals, between for the
simultaneous to result in is described by band the and used,
testing between two false indications. which wide weaker the enough signals complex provide stronger to is
promising signals,
of The
two all
or three
more margin 6 db a
frequency can suppress analyzed. is stronger
contain
mainbeam
signals. circuits,
required in which for A thorough and of
weaker More can
approximately
non-linear
suppression
smaller analysis its
ratios of mainbeam of the bandwidth capabilities of
to sidelobe requirements should unlock be
signal levels. for this circuit made. or false-lock are
suppression serious
(to)
No
problems signal and bandwidth. increase;
tracker
occur
when
the
mainbeam
cross-coupled
signal
simultaneously the analog should be
within velocity tested with
the tracker outputs may spread-spectra
However, the noise on this interference effect
signals.
C-17
APPENDIX D AVAILABLEETAILS D OFVENDOR UNITTESTS I. INTRODUCTION
I
Information referenced in II. VERIFICATION A. KPSM i.
in this appendix section V.C+I. TESTS
was
taken
directly
from
the
Ryan
documents
Klystron Parameter High
Requirements DVS Klystron + 75 vDC RA -800 + Klystron +20 record 0.25% 55 ma 5.0 sec I0 vDC 0._ microamp (max) 2.0 + 0+5 sec. -500 + 45 5 sec vDC ! 1% to _ 65 ma 5. sec. vDC i0 vDC (cathode) vDC (reflector)
Voltage
-2150 +
ripple regulation current time delay Collector ripple regulation current time Filament ripple regulation current time 2. delay Characteristics Rate Signal delay Voltage Voltage
record 0.25% 40 to 20.0 + -500+_ + 1%
record i0 microamps 20 _ 7.2 !
record
20.0 6.3 !
0.3
0.3
record _ 0+15 vDC 0.8 to i.i amp 0 sec. Requirements = = = = Amplitude Width Rise Time = = =
record _ 0.15 vDC 0.9 to 1.3 amp 0 sec.
Modulation
Repetition Flyback Flyback Flyback Start Start Start Noise on AM RF
Amplitude
Pulse Width Rise Time Sweep Sweep Sweep Output Pulse Pulse Pulse
182 + 5 cps -2.0 to I0.0 vDC I0 to 160 microseconds i0 microseconds -3.0 3 to to -ii.0 vDC (max) 30 microseconds
3 microseconds
3.
Requirements noise in in i00 high Hz 125 away, 102 i00 Hz BW 115 away, 92 db db db db below rising at 400 carrler 3db/octave Hz away at 80 to kHz at 80 kHz to
sideband
BW on RA klystron and low dev. AM due Modulation Sweep Average sidebands to power Rates time rate in
below rising at 400
carrier Hz away
supply at -20°C
ripple Requirements = =
3 db/octave
4.
5.0 and
+
0.5
msec + 2.4% + 1.5%
8,000
MHz/sec 800 MHz/se_
D-I
5,
XMTR
Frequencies at times 5½ min.
at
-30°C
Requirements on: 30 sec., 2 min., 3 min.,
Measure 4 min., RA DVS
after
turn
frequency frequency Test KPSM warm up Requirements @ +75 time. !lO°F,
12.9 13.3
GHz GHz
_ +
25
MBz
35 MHz
6o
Thermal-Vacuum Stabilize 4 hours. Check Check system
< 5 x
10 -6 RA
torr, high
for
minimum
of
XMTR's
freq.,
power,
and
deviation
rate.
B.
R/T I.
Units Power Consumption 225 Detector DVS RA Angle E Plane H Plane H Plane Angle with between Lab Test beams Adapters 0°0 ' + 4' 12o30 T + 4' 25o0 ' +_-8' ma from +25 vDC; 15 ma from -25 vDC
Requirements: 2. RF
Bias beams: beams: -3 + 2.0 dbm -2.7 + 2.0 dbm
Requirements:
3.
Beam
Requirements:
4.
Insertion XMT
Loss
Requirements:
Flanges:
I db 4.5 7.0 db db
RCV Flanges: Detectors 5. Two Way Gain 56 Way db min. (3 db) 5.3°max
Requirements: 6. Two
Beamwidth E Plane:
Requirements: 7. VSWR
H Plane:
3.5°max
Requirements: 8. First
1.3:1 Order -30 Figure DVS @
max
at
XMT
and
RCV
flanges
and
detectors
Sidelobe db min. COverall 800 Hz 8 kHz 80 kHz
(@ _ 8 ° )
Requirements: 9. Noise
Receiver) 25.9 19.O 15.8 23.8 17.1 db db db db db max max max max max
Requirements:
RA
@
8 kHz 80 kHz
i0.
Microwave
Isolation between between XMT feeds: XMT 20 db (min) feeds: 55 db (min)
Requirements:
opposite Test at
& RCV
ii,
Thermal-Vacuum Stabilize for
Requirements:
+125 ° + of 4
10°F
(at Check
preamps), system
< 5 x warm up
10 -6 time.
torr,
minimum
ho_rs.
D-2
C. SDC Stabilize at +105 _ 10°F (at the LVPS),< 5 x 10-6 torr, for minimum ° of 4 hours. Checkanalog outputs, range marks, and sensitivity. Checksystem warmup time. III. ACCEPTANCE TESTS A. Laboratory Ambient
i. KPSM Modulation Low Rate Deviation Modulation Rate (KPSM) in the temperature Apply power to the KPSM. klystron flange temperature
Place the Klystron Power Supply/Modulator chamber and allow it to stabilize at _ 30°C. Measure the average after 5½ minutes. Requirements: Hish Place at _ 30°C. deviation at 1 min. the rate time KPSM deviation rate and the
800 Deviation in power the
MHz
sec
+
1.5%
; Rate
5.0
+
0.5
millisec
Modulation temperature the KPSM.
chamber Measure and the
and the
allow average
it
to
stabilize rate, temperature
Apply
to
deviation
at the sweep extremes, intervals thru 4 min. After ist reading, MHz/sec MHz/sec MHz/sec Frequency temperature to the at RA, +
klystron
flange
Requirements: Average: Lower Limit:
8,000 8,000 + 2,000
2.4%
Upper Sweep
Limit: Time:
8,000
MHz/sec
+ 2,000 MHz/sec _.0 + 0.5 Millisec
RA Place 30°C. (undeviated)
and the
DVS KPSM the
Klystron in the
chamber KPSM. 30 12.9 sec., GHz
and
allow the
it
to
stabilize of 4 min., the 5½ +
at RA min. MHz
Apply and
primary DVS (all DVS
power
Measure
frequency 3 min., DVS,
klystrons times) Power the
2 min., ! 25 MHz;
Requirements: RA and
13.3
GHZ
35
Klystron to
Apply primary power record the results. Requirements: 2. R/T Units Gain RA,
KPSM.
Measure
RA
and
DVS
Klystron
power
and
250
mw;
DVS,
8.5
+
1.5
w.
Two-Way Measure
two-way
antenna Beam Beam Beam Beam I @ 2 @ 3 @ 4 @
gain 13.3 13.3 13.3 12.9
on GHz GHz GHz GHz
all
beams. 56 55 55 56 db db db db (min) (min) (min) (min)
Requirements:
D-3
Beam Measure plane angle
An$1es E plane between E and beams Plane Beam Beam i 4 0°0 ' + 0°0 ' + 4' 4' H plane in beam angles for all beams. Measure H
each
unit. E Plane Beam Beam 2 3 0°0 ' + 4'
Requirement:
0°0 ' + 4'
H Plane
O
H Plane 1 4 12 12 ° 30' 30' + 4' + 4' Beam Beam H Plane 25 ° 0' + 8' Between 2 3 17 ° 23' + 4'
Beam Beam
17 ° 23'
+ 4'
H Plane Between
Angle Beams
An$1e Beams 34 ° 46' + 8'
Insertion Requirements: P2 P3 P2 P2 P3 P3 XMIT XMIT Flange Flange
Losses
4.5 4.5 4.5 4.5 4.5 4.5
db db db db db db
(max) (max) (max) (max) (max) (max)
PI P4 Pl PI P4
XMIT XMIT
Flange Flange Flange Flange
1.0 1.0 4.5 4.5 7.0 7.0 7.0 7.0
db db db db db db db db
(max) (max) (max) (max) (max) (max) (max) (max)
/0___ Rec ° /90___ Rec ° / 0 ° Rec /90____Rec °
Flange Flange Flange Flange
/ 0° Rec /-_0° Rec /O ° Ree
Flange Flange P4 /90 ° Rec Flange Flange
A B C D
Record
insertion VSWR
losses
of
adapters
to be
shipped
with
antennas.
Measure
the
VSWR
at 1.3:1
the
points at
given. all RCV & XMIT flanges.
Requirements: Two-Way Take 3. SDC Response Apply shown in apply with Tests. per ten beam Beam
(max)
Patterns measurements on all beams and attach to report.
pattern
Time primary until input voltage. Apply the tracker under test the signals acquires then
22.4 + 0.0, -0.2 VDC the first column below
the step frequency the graphic recorder (Response Apply time input is cent.) times.
shown in the second column. Monitor the results and retain the recorder tapes for the Report on the time at for a reduction of of 20.0 the mv. output Conduct error each by 63 signals level test
D-4
Requireme_ts: V
X
0.115
sec.
max
for R
average Step
Z
of 1
i0
attempts. kHz .... 0.930 kHz .... 0.930 kHz kHz kHz kHz .... 5.880
Step
1
DI=I.60 D2=1.33 DI=1.33 D2=1.60 D2=1.60 D3=1.33 D2=1.33 D3=1.60 DI=1.33 D3=1.33
kHz kHz .... 0.930 kHz .... 0.930 kHz kHz kHz .... 0.930 kHz kHz .... 0.930 kHz .... 0°930 kHz .... 0.930
DI=1.33 D3-1.33 D4=1.60
Step
2
Step
2
DI=I.60 D3=1.60 D4=5o33
V Y
Step
i
Step
2
V z
Cross-Coupled as Gain t___ates S Apply 22.4 +0.0, described in the Freq. (kHz) D2 D3 i0 i0 i0 i0 D2 = D2 = D3 200 200
Side
Lobe
Logic input voltage. Apply the signals
-0.I VDC following
primary steps.
Step 1 2 3 4 5 6 7
8
D2 D3 90190 65 40 40 40 90 90 65 90 90 65 65 65 40 40
Levels !5mv; _5mv; !5mv; 0.5 D 3 = 200 Decrease Increase my; !5mv D 3 from D 2 from D3 ! D2 280mv lOmv from 40mv Both
Requirements track at at at at at at at 200 D2 = 30 _ ! 42my 14 ! 6.2mv .02 42my 14 _ 6.2my 3mv kHz 3my
D 3 dropout D 3 dropout D 3 dropout D 3 acquire
i0 ii0 i0 [i0 I0 vary I0 i0
= 250 30 ! 30 ! 200
D 2 = D 2 = D3 =
Decrease D 3 = 20
0.5mv; ! _ 5mv; 5mv; 0.5mv;
0.5mv from 280mv 10mv 40my
I0. i ! 200 D3 = 30 _ _
Decrease Increase Decrease
D 2 dropout D 2 dropout D 2 dropout
lOll0 I0 I0
D 2 = 250 D3 = 30 !
D 3 from D 2 from
Thermal Record Vibration i. General Vibration serial
Sensor
Data and check continuity and isolation.
numbers
B.
Each unit shall to vibration in
be vibrated separately. Each unit shall accordance with the following schedule.
be
subjected
Nonoperatin$ Sine wave 5 to 16 125 The over two in sine minute two other wave to to 16 125 1500 Hz @ Hz @ 0.45 Inch Da
Hz @ +
6 G Peak
+ 2 G Peak be logarithmically wave to on vibration parallel the thrust unit. each swept shall to the axis from 5 to 1500 of axis total Hz two and of
frequency period. in wave an
shall The axis
a two
minute sweeps sine
sine
consist thrust for a
essentially time
critical
axes
orthogonal
12 minutes
vibration
D-5
Operating Upon completion of the two two minute sine wave vibration sweeps in
each axis, subject the unit a power spectral density of limited between 50 and 2000 minutes ments of taken WGA as on each unit. in described Setup
to white gaussian acceleration (W_A) with 0.002 G2/Ha + 0.002 G 2, - 0.001G ; band Hz. The unit shall be subjected to two The unit shall be operating and measurefollowing paragraphs.
Test Attach obtain
each unit to the vibration exciter in such the desired acceleration without attempting installation. Load each unit as
a manner as to simulate to make
to best the it dynamic-
spacecraft
necessary
ally similar to the flight configuration. on the exciter as near to the supporting otherwise vibrated other axis. 2. KPSM Nonoperatin_ Subject tion of physical the the KPSM two (Sine Wave_ noted in an in axis the detailed axes unit which essentially parallel are
Observe the vibration level bracket as possible. Unless the units axis to shall and the be in two thrust to the thrust
procedure,
critical
orthogonal
perpendicular
to sine wave vibration as two minute sweeps, visually Record any defects noted.
described inspect
above. the unit
Upon complefor any
damage. (WGA) Setup to
Operating Test Attach the KPSM
its
vibration
fixture
by
means
of
its
normal
mounting
provisions. the thrust antenna monitor with the
Mount axis. the
the fixture on the Attach accelerometers. test equipment noted.
exciter head for vibration along Connect the KPSM and an RA/VS to provide the voltages and to
necessary
parameters
Measurements Record the voltage and current from the three power supplies. Monitor DVS
frequency and power and RA frequency and power thirty seconds after turn-on and every thirty seconds through 120 seconds. On a tape recorder, record interference levels from the RA/VS antenna preamplifiers as a function of vibration into the frequency X-Y plotter. (using a spectrum the analyzer). plots and Play retain the for magnetic Report tape on Tests.
Identify
the
Play the magnetic tape into SDC Tracker Lock lamps (illuminated signal-plus-noise the preamplifier at to any 2.5
trackers and record or extinguished).
the condition of the Operate the SDC in the noise.) interference shall be Record peaks limited
to noise mode using normal preamplifier channel and level and frequency of any Lock lamps duration. RA DVS RA DVS illuminate. This
time the Tracker minutes maximum XMTR
test
Requirements:
Freq.,
= = = =
12.9 13.3
GHz GHz
_ +
25 M}{z 35 MHz
XMTR
Power,
250 mw (min) 7 watts (min)
D-6
(at all times) +25 vDCsupply, +25.0 + 0.25 vDC, 60 ma(max) -25 vDCsupply, -25.0 +_0.25 vDC, 5 ma (max) 22.4 vDCsupply, 22.4 + 0.25 vDC, record current
Reference Upon completion below. PARAMETER RA DVS RA XMT XMT XMT Power Power Freq. Tests of vibration tests, measure t_e parameters listed
REQUIRED 350 12.9 13.3 + i00 CC + MW 25MC 35MC
PARAMETER High Dev. Rate Low Dev. Rate Deviation Rate Flyback +25 VDC VDC VDC VDC up (Max) Repetition
REQUIRED 8.0 0.8 182 Time Supply Supply Supply Supply Time Voltage Current Voltage Current Required 0.5 GHz/sec GHz/sec + + 5 Hz 0.025 Millsec
7 WATTS
(MIN)
DVS XMT Freq. 22.4 vDC Supply Voltage 22.4 vDC Current 16.5 vDC Voltage 16.5 vDC Current 3. R/T Supply Supply Supply
CC + Record Amps Record
Record 60.0 Ma (Max) Record 5.0 30 Ma Sec (Max) (Max)
18.0
+25 -25 -25 Warm
23.0
Amps
(Max)
Units (Sine to Attach Wave) its vibration fixture by means the of its normal to mountwave
Nonoperatin$ Attach ing vibration visually noted. the antenna
provisions.
accelerometers.
Subject
antenna
sine sweeps, defects
as described. Upon inspect the unit for
completion of the two any physical damage.
two minute Record any
Operating Test
(WGA_ Setup on the Attach vibration fixture by three accelerometers vibration in the means of its normal mountto each of the three mountaxes:
Mount the antenna ing provisions. ing points I. 2. 3. Attach point Show, of with tain the the ed the for
monitoring
following
Vertical Normal Tangent to to the the Antenna Antenna at most the one of the mounting points or at suitable for equlization control. X-Y charts, at the use the ofthe the number input and location to oband/or control Analyzer accelerometer level Equalizer a
the control accelerometer on the vibration fixture by the the means Analyzer specified of a diagram Monitor Equalizer vibration on and
accelerometers.
equalization adjust by levels
vibration
peak notch vibration at any two Tape
filter of the vibration input level to maintain of three comparable the axes to monitor
system. During vibration, the specified vibration as accelerometer outputs. the Apply outputs. Apply required RF the
adjust monitorrequired
Connect energy
Recorder
preamplifier
operating voltages to the preamplifiers. to the antenna input ports.
D-7
Measu a. Prior to
reme n t s vibrating the antenna, for monitor approximately all preamplifier two minutes outfor the Re-
puts on the Tape reference.
b,
Recorder
Apply the required vibration input levels in-line accelerometers on the X-Y Recorder During corder Scan peaks. Reduce are in the vibration, record for two minutes. the preamplifier in any a I00 Hz the preamplifier with the Record peaks of the a). After output on the until Record
and monitor for seven outputs on Wave of
each of minutes. the Tape
c0
d.
outputs bandwidth. resonant
Noise the the the
and
Spectrum any discrete outputs noted and the
Analyzer e.
levels
discrete I0 db
preamplifier output vibration level level
within
(nominal) test (Step
preamplifier
reference
preamplifier tape recorded Spectrum into noise trackers to noise
output level. preamplifier and record and or mode Tests of vibration record using
completion of tests, play the signals through the Wave and Noise the X-Y Recorder. of the the SDC in Play the the tapes lamps Tracker noise.) Lock
Analyzer
condition (Operate
(illuminated
extinguished). normal
signal-plus-
preamplifier
Reference Upon completion below. Overall
tests,
measure
the
parameters
listed
receiver
noise
figure
@
8kHz:
DVS, RA,
19.0 23.8
db db
max max
Preamp
gain switch 280 mv, after switch max. state output
levels: switch 317 mv,
DVS, before switch I0 +(i0, -0) mv; RA, after 13.5 switch ! 1.0 in 20 vDC 90 db (+i0,
max before -0)mv.
Preamp Max. 4. SDC
gain preamp
signals: amplitude
balance
gain
state:
_
1.0
db
Nonoperating Attach tion as the SDC to
(Sine its
Waye) vibration fixture of by the means the two Record of SDC any two its to minute normal sine mounting wave vibravisually
provisions. inspect the
Attach unit for
accelerometers. Upon any completion physical
Subject damage.
described.
sweeps,
defects
noted.
Operating Attach the SDC
(WGA) to its vibration following fixture tests by means of its normal with after and mounting the test vibra-
provisions.
Attach
accelerometers. the
Interconnect before,
the during
SDC
equipment. Conduct tion in each axis: (a) Apply in the the
Acquisition doppler mode: kHz kHz D3 D4 = = 24.51 75.010 kHz kHz return sine wave frequencies shown below
simulated
signal-to-noise D1 D2 = = 24.51 19.661
D-8
Set the DVSinput signal levels at zero mvrmswith a noise density of 2.5 mvrms in a I00 Hz bandwidth. Set the RAinput signal level to 6.0 mvrmswith a noise desntiy of 0.92 mvrms in a i00 Hz bandwidth. Set the primary input voltage at 22.4 +0.0, -0.2 VDC. Turn the BURNOUT SIGNAL switch to OFF. Turn the DVS PRE-AMP GAINSTATE SIGNAL switch to 90 db. Turn the RANGE PRE-AMP GAINSTATE SIGNAL switch to 80 db. Apply RADVSowerand start the recorder chart p drive motor. Observethe DI, D2, D3, and R Tracker Lock lamps. Slowly increase the signal level on each channeluntil the Tracker Lock lampsilluminate. Recordthe signal level required for each channe i. Requirements: DVS,29.0 to 51.5 my; RA, 6.9 to 9.8 mv Measure analog outputs Vx, Vy, Vz, and Rz with the Digital Voltmeter. Requirements: V = 15.00+ 0.71 vDC, V = -15.00 + 0.71 vDC,V = 50 vDC(satuXated), R -- + 30.00+ _.14 vDC. -z Z
Measure Requirements: Record Graphic the voltage record SDC and current at the primary = 8.5 A input (Max) of I0 seconds on the power source. voltage, current for
analog
outputs
a minimum
Recorder. (b) Analog Wave Noise signals = 170 kHz kHz switch Turn the to ON. the RANGE input shown mv rms below at D4 the = D3 D4 Turn power following mv rms kHz kHz PRE-AMP STATE + 0.0, GAIN SIGNAL -0.2 VDC. STATE levels:
Apply
the
Sine D1
= D2 DI D2
= D3
118
= 4.902 = 5.710 SIGNAL 65 db. Set
= 4.902 = the at 10.124 DVS GAIN 22.4
Turn SIGNAL switch
the to
BURNOUT to db. 60
switch
PRE-AMPS
primary and with _t the
Apply primary input voltage Measure the analog outputs peak-to-peak lO0-second Requirements: analog period Noise: noise containing V V
Z
start the recorder chart drive motor. the Digital Voltmeter. Determine the of S/C simulation filter) Record v p-p v p-p over all (max), (max) the maximum (max) (max), 2.50 excursion. V = 0. i00 ' y R
Z
output
results.
x
= =
0.i00 0.250 vDC, vDC,
v p-p v V y R
Z
p-p
=
0.400
Accuracy:
V x V
Z
= -2.50 = i0.00 Reference
+ 0.140 -+ 0.140
--
= +
-- 3.00
+ 0.140 vDC, -+ 0.120 vDC
--
Test of the vibration check-out and in RADVS I kHz tests, interconnect and the SDC the with the test tests.
Upon
completion at (I) the
equipment
SDC
Unit
station Telemetry BW from
perform
following
LVPS supply
Ripple ripple
Signals 2 kHz to i00 kHz. max; mv max.
Measure Requirements:
power
+25 v supply , 2.0 mv max; -25 v +I00 v supply, 20.0 mv max; -I00 telemetry signals in the "ON" and
supply, 1.0 mv v supply, 20.0 states.
Measure
"OFF"
D-9
Requirements:
ON
OFF
Reliability signals and 5.0, + 2o0/-0.4 vDC 0.0, + 0.4/-1.0 vDC range marks II.0 + 2.0 vDC 0.0, + 1.0/-2.0 vDC D-lock signals 0.0, + 1.0/-2.0 vDC 13.0 + 2.0 vDC R-lock signals (2) Linearity and Accuracy
Apply Measure simulated Vx, DI D2 Vy, = = signals V z, and kHz kHz -5.0 35.0 + 0.47 -+
--
at Rz
the and
frequencies record the
and
level
shown
below.
results. D3 D4 = = 17.157 83.421 + 0.47 -+
--
17.157 18.773 V x V
Z
kHz kHz vDC, vDC
Requirements:
= =
vDC, vDC,
V Y R
Z
= +5.0 = 40.0
0.47
1.52
(3) Same test set (4) Same test set (5)
Linearity up and
and accuracy
Accuracy measurements as test (b) above. (Redundant)
Noise up and
and
Ripple measurement i and as One test (b) above. Foot (Redundant) Mark
noise
Linearity (a)
i Accuracy and the
Thousand
Linearity at
Accuracy and level shown below. Measure
Apply V x , V
simulated V y, z DI D2 , and
signals R z and kHz kHz
frequencies the results.
record
= 2.451 = 1.643
D3 D4 - 20 mv + 0.095 + 0.095
--
=
2.451
kHz kHz BO. 0.095 0.066 vDC, vDC
- 5.062 state, + +
--
Signal Requirements: V x V
Z
Level = +2.5 = 5.0
rms, vDC, vDc,
Low V
gain Y R
Z
= -2.5 = 1.5
(b) With the range Requirements: (6) Apply Vx, simulated Vy, Vz, DI D2 the i000 test foot setup mark output R
g
One the is at = 1.0
Thousand same which + as the
Foot above,
Range
Mark
Accuracy the frequency input of D4 until and
decrease the
generated.
Record mark vDC,
range
frequency
analog
is generated. R-freq. = 4.331 to 4.168 kHz
0.047
Linearity signals
and at
Accuracy the frequencies the results. D3 D4 = = 1.716 16.470 BO, kHz kHz HI DEV. vDC, vDC and level shown below. Measure
and Rz and record = 1.716 kHz = 1.554 Level V V
Z
kHz = 20 m v rms, + 0.068 -0.068 Low vDC, vDC,
Signal Requirements:
gain V R
Z
state, = -0.50 = 18.0
x
= +
0.50 +
--
Y
+ 0.068 -+
--
= 3.50
0.31
D-10
(7)
Apply Measure simulated Vx,
Linearity_ (a)
Accuracy and the
a and Accuracy
Fourteen
Foot
Mark
Linearity at and
signals and = = Vz 0.123 0.123
frequencies the
and
level
shown
below.
Vy, D1 D2
record
results. D3 D4 = = 0.123 0.445 gain kHz kHz
kHz kHz level BO, = 20 mv HI DEV.
Signal
rms,
low
state, Requirements: V x = 0.0
+ 0.049 vDC, -V = 0.25 z Foot Range
V + y 0.049
=
0.0 vDC
+ 0.049 --
vDC,
(b) With until and
Fourteen
Mark
Accuracy of D4 frequency
the test setup the same as above, the 14 foot mark is generated. range analog R
Z
decrease the frequency Record the range input mark is generated. = 387 + 38
output = 0.339
at +
--
which 0.042
the
Requirements: (8) With the
vDC,
R-Freq.
Hz
Acquisition in the
Time signal-to-noise sinewave acquisition signals at the mode (preamp high and
trackers apply below. DI D2 = =
gain state), levels shown
simulated
frequencies
I0 kHz i0 kHz Levels = 21.7 bandwidth. Levels bandwidth. Level = 9.25 bandwidth. Level = bandwidth. 0.92 my mv = 2.50
D3 D4 mv mv
= = rms
I0 80 in in in in a
kHz kHz
DVS Signal a i00 Hz DVS RA RA Noise Hz i00
rms rms rms
a a
Signal i00 Hz Noise i00 Hz
Turn the the time TRACKER
Burn Out Signal ON. between reapplication LOCK lamps. All Record trackers Time time test
Momentarily remove of the signal and the results. signal within
the signals. illumination
Measure of the
Requirements: (9) Same as SDC
acquire
4
seconds.
Response response
in
section
III. A.3
of
this
appendix.
D-II
APPENDIX E DETAILS OFTHE VENDOR
I. INTRODUCTION Information Part II. I. in this tests CONDITIONS test conditions the for the vendor tests are listed in Table and E-l, notes which for appendix and is taken from have Ryan been report omitted.) 51765-2B (change 12), SYSTEM ACCEPTANCE TEST
(Mechanical TEST
inspections
STANDARD The
standard
was the
copied from Table i-i of table are listed below: AV AR BO BBO HD LD M MI M2 NR RV SR SV VV
document
referenced.
Abbreviations
Velocity
Sensor
Accuracy
RA Accuracy After Burnout Before Burnout High Deviation Low Deviation Analog 1,000 14 RA RA Foot Noise Foot Range Mark Required Capability Sensitivity Velocity Capability Not Mark Range
Operation
Range-Velocity Sensitivity Sensor Sensor
Velocity Velocity
Vertical Trajectory Range for DVS Return Based on Lambert Law Return Power 45 ° Attitude A The description Test of the STC's copied provide from the for
Power Scattering Feet Range and
40,000
the
referenced signal
document
follows: required specifirequirethe DVS at the
Standard
Conditions
essential
characteristics
to demonstrate cation. Each ments. and two the It RA will
that the RADVS will meet the Standard Test Condition checks be points noted This on a that in certain a means at provides trajectory
requirements a number of the time. effectively
of the basic product the basic performance equivalent checking The Table range system also of
instances, for the same
differ.
operation the tests is as
simulated
indicates
mode of performed follows: STC i:
operation of for such a
the RA and condition.
the DVS, as well as the The detailed explanation
specific nature of for each condition
The lent
RA of
is
nonoperating. FPS along altitude RA and the
The each and DVS
DVS of
is the
operated three capability angles
at
maximum beams. at
range This,
and angle
the on
equivaall The beams max-
3,000 maximum
doppler
therefore,
checks STC 2:
velocity at pitch
45 ° pitch 40,000
simultaneously. Test both the
of
45 ° and
feet.
imum search requirement FPS at a range of 40,000 linear STC 3: Test the horizontal the RADVS RA and will not velocity DVS with
for the feet. output negative the
RA is tested, Test the DVS capability. simulated retro
which at the input
occurs with a V z of 740 required maximum negative to demonstrate after it is that
signals as a
acquire
main
tankage
target
jettisoned.
E-I
0
0
0
0 0
0 0
0 a ,-,,1
o
0
0
o _o _o
Z a
I
E
a_ u'_ 0 © 0 0 0 0 C u'_ o3 -40 0 0 0 00 0 0 0 0 0 0 0
,..-4
0
2
n_ r_
0 ! u'_ I '1 I I
.,--I I
! 0 n_
I
I
I
i 0c u_
u'_ I
I
e,4 o ,-.i !
_jD
_
0
u_ c_ 0
-,.1
c_c_
0.-.I-
,-4
m 0
.,-I v
0
•_
0
_
4_
0 C 0 ¢,I v
09 c_ C _-_ cxl 0
1..1 C C 0 uC r_ 0 00
0" 4J U'I
J
+
0 0 i 0 0 0 C Oi 0 0 ! L_ 0
0 ,--_ 0
P_
v
+
I
0
N
,--4
v
0 0
0
0 0
0
0
0 !
0
I
0
0
,
+
[,,-a
v
0
_I
on
0
0
oi c
o,
re3
o
0
,-.I ,...-4
E-2
STC
4:
Test
the
DVS
at
the
required
maximum the
positive mid-range
linear of the
horizontal linear
velocity and Tests
output the RA STC 5: Tests the STC 6: DVS
capability. for maximum the at RA for the
Test V z in sensitivity. a near mid-range maximum
requirements mode.
altitude linear mode is tested can of
in
the
low
deviation output
required
vertical at 2,000 for 232902. occur
velocity feet where
capability errors per for effecvermark The may the
and V. and V x at the null. TestsYthe RA in the low deviation be expected foot of and the to occur. curve to the to for of feet its lowest and The HAC DVS
maximum capability
acquisition The on DVS the each
50,000 ranges tively tical STC 7: Tests and
Specification values pitch which angles vector. and
doppler subject beam to foot the of as mark in the
frequencies curve turn lunar
correspond
50,000 the RA
45 ° , with
oriented
most
unfavorable
position of to at at
relative the 1,000 between
relative the from the
velocity generation low and
accuracy mode
range mode.
transition
deviation
the
high
deviation
DVS is tested for accuracy range and zero velocity. STC 8: Tests leading generated the RA in The RA and the the down high for quirements. the range unit
linearity mode is in the
points near
mid-linear linear to DVS been is testre-
deviation frequency lock. placed
mid-range high so foot
the not has The
sufficiently the 1,000 high deviation
require
(Assuming
mode.)
ed for accuracy around the null of V_ and V. and low positive values of V z. x The problem specifically tests the DVS acqulsition capability in accordance with the 40,000 foot curve of the HAC Specification No. 232902. As in STC 6 the of individual for beam each doppler the with worst frequencies condition to the of were lunar DVS selected vertical power as and the and the minimum worst relative to the frequency orientation velocity two different occuring vector. altitudes STC 9: 45 ° pitch angle
respect
The two values indicated for
of received the DVS.
correspond
The STC was constants. for linearity
formerly used, in conjunction These tests are now conducted and accuracy measurements and The after burn-out. observation modes. mode. condition. side lobe under D2 which
with STC 8, to measure on a unit basis. STC only. of the 14 in foot its range wide in in of
9
system time is used mark and mode mode apto simrethese is was 3 to 4 which during
STC STC
I0: Ii:
Tests the RA for the generation the DVS at very low velocities. Tests prior with and ulate quires formerly procedures, more unit plicable the the narrow operation large tested operation CRO logic circuitry. the CRO after allow various the CRO mode to burn-out, test
accuracy DVS is The
operated and sequence and is
band the the the and RO
burn-out,
finally defined made lobe
bandwidths procedures of the in under with angular
measurement at side is tested parameter
timing which
frequency may With occur
variable rejection of using is Change STC tested
dispersion this
burn-out
Cross-coupled rejection which
incorporation this
cross-coupled
compatible tests. VALUES shows
conditions
III.
REQUIRED Table E-2
AND the these not
TOLERANCES total For errors allowable application must in be RADVS to system system in 3_ tests errors using as the with antenna and the alignDigital listed RADVS in Test the
Ryan
document
referenced.
Equipment
(RADVSTE), (I) Errors
adjusted
accordance i.e.,
included
RADVS/RADVSTE terrain error.
tests, bias
ment and Voltmeter (2) Normal
boresight error.
errors,
errors,
RADVSTE
measuring
E-3
I t I
O O •
Lm OO ("4
0 0 u'3 _'_0
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+I
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+1+1 +1 + , +1 +1 ++1
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0
u'_ I_ _o ,,-_ .._0
0 0
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+ +1
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+1
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z
°
o
Z
_D
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_
E-5
The duced add IV.
adjusted testing limits, in such
values and the
are data RADVSTE to
given
in
Table If must their
E-3. RADVS be being
These checked
tolerances falls from to determine
were if
selected these the
to
exreerrors
pedite
evaluation. accuracy justify
performance eliminated
outside RADVS
RADVSTE tolerance.
a manner
the
TESTS A. Check Power that at Consumption power 16.5, from 20.0, the and 22.4 22.4 v DC supply does not exceed 590 watts with the
supply B.
set
vDC.
Thermal
Sensors and isolation of sensors.
Check C. RF
resistance Power these
Conduct + 0.0, Retain all
test
with for
primary the RF this Test
input on
voltage each
at
16.5
+
0. i, for
-0.0
and
22.4 losses.
- 0. I VDC.
Measure
power
beam,
allowing
insertion
computations
Report. w min each beam; RA, 210 mw min.
Requirement: D. XMTR Frequency the at frequencies 16.5 +0. I,
DVS,
1.5
Measure input voltage
using -0.0 RA,
RADVSTE. and 22.4
Perform +0.0, 25 -0.i MHz;
these VDC. DVS,
tests
with
primary
Requirement: E. Standard (I) Test Test Setup the Condition
12.9 Tests
GIlz !
13.3
GHz
_
35
MHa
Interconnect RADVSTE. 16.50 and (2) Test STC i:
RADVS
with input
the
RADVSTE. at
Set
up
the
STC value
on
the
Set primary 26.0 v]. for STC's
voltage
[a specified
between
Listing (a) (b) (c) (d) (e)
Thirty-seconds power DVS Linearity DVS DVS DVS RADVS Maximum Capability RA RA RA Maximum Maximum
Warm-up and Slant Total
at
26
VDC
Primary
Input
Accuracy Range Capability Capability Sensitivity Velocity
Acquisition Linearity Horizontal
Time and
and
STC
2:
(a) (b) (c) (d) (e) (f)
Accuracy Negative Linearity Output
Maximum Maximum Maximum
Slant Range Capability Velocity Capability at Attitude Angle
40,000
Feet
RADVS RADVS DVS RADVS Warm-up RADVS
Acquisition Linearity Maximum and Accuracy Positive Linear Output
STC
4:
(a) (b) (c)
Horizontal
Capability Acquisition Time Linearity and Accuracy Linear Output Capability
STC
5:
(a) (b) (c) (d)
DVS Maximum Vertical RADVS Acquisition
E-6
STC6:
(a) (b) (c)
RADVS RADVS DVS RADVS 1,000 RORA Tracker RADVS DVS
Linearity Acquisition Sensitivity Linearity Foot and Lock Linearity Sensitivity and Range RODVS
and
Accuracy
STC7:
(a) (b) (c) (d)
and Signal Signal and
Accuracy Accuracy Accuracy Accuracy Accuracy
Mark
STC
8:
(a) (b)
STC
9:
(a)
Linearity RADVS Noise
Accuracy and Range Accuracy Mark Accuracy in the Search
STC I0: (a) (b) (c)
STC ii: (a) (b) (3) Typical Make lifier controls changed
Linearity and Ripple Foot
Thirteen CRO and Logic Track
Signal Modes
Accuracy Side-lobe
Cross-coupled,
Rejection
Measurements all which or when Recorder and signal effect tapes signals are identified on each monitored function of the by date, on the test, When note tape. ampRADVSTE are the Retain level, recorded being operational left margin On Test. channel. occurs, Recorder
sure
Recorder
a normal on the Report Accuracy Recorder
time of the event all tapes for the (a) Linearity 1 Start POWER ON. called luminate, operation. Vz, and/or results on drive ing 2 motor range and the Record
chart Turn the the
drive the of RODVS the
motor. TEST and/or tape Burn
Turn Out RORA
the
RADVS to (if il-
switch for).
to ON. the After
ACTIVATE
switch Signal lamps
presence
examine
Recorder
to verify
reliable Vx, the chart Vy,
Turn the VOLTMETER switch to measure R z on the Digital Voltmeter. Record the data sheet. Permit the Recorder to and run for at least i0 seconds while velocity analog outputs.
record-
Turn the BURN OUT Record the absence Repeat the results.
SIGNAL switch to OFF (if called for). of the Burn Out Signal on the recorder. of Vx, Vy, and V z and record the
measurements
(b)
Acquisition I Turn RORA the lamps
Time TEST
and
Sensitivity switch examine to ON. the When the RODVS and/or time OFF.
ACTIVATE
illuminate,
Recorder
tape
to verify
reliable operation. on the data sheet. 2 Turn ments. TEST the TEST Record ACTIVATE ACTIVATE the switch
Record RA and/or DVS Turn the TEST ACTIVATE switch on to OFF. to ON the amd
acquisition switch to the Turn
repeat sheet.
measurethe
results
data
E-7
Turn the TEST ACTIVATEwitch to ONand repeat the s measurement.Recordthe results on the data sheet. Observe(a given) TRACKER lamp. Force loss of LOOP lock of the tracker. Recordthe attenuator setting at which the tracker drops out on the data sheet. Decreasethe attenuation until the tracker locks on. Recordthe attenuator setting at which the tracker acquires on the data sheet. Turn the TEST ACTIVATE switch to OFF. 5 Repeatstep four for other trackers (as indicated). (c)
Warm-up Start ACTIVATE on. Time the Recorder switch the chart to ON. time of drive Turn the motor. the RODVS RADVS and/or Turn POWER of RORA the TEST to
switch spacecraft signals
Record
between
application
power and indication on the data sheet. (d) Analog Set the CHAN CHAN Zero the Transients recorder I 2 SC SC Due
to Preamplifier to the CHAN CHAN
Gain
Switching
channels FIL FIL pens on
following: 3 4 SC SC 1,2,3, FIL FIL and 4 using the
Recorder
channels
Vx, Vy, Vz, lifier gain Adjust ensure i When chart SIGNAL switch ed gain 2 Repeat state Repeat the on
and R OFFSET-SC FILTER levels on channels I,
controls. 2, 3, and
Set the amp4 at 50 MV/LINE. to
the MICROWAVE INPUT SIGNAL that all preamplifiers are all trackers motor the have on 2, acquired each 3, the and beam
ATTENUATION to a level in high gain state. and until Observe the on with the all the the Recorder INPUT
drive to
running,
decrease
MICROWAVE values chart sheet. the low
ATTENUATION channels switch. I,
preamplifiers recordat
mid-gain Record
state. results for gain
4 of
Recorder the data to
the measurements on all channels. the measurements state on
switch
gain
3
for all
gain
switch
from
the
low
to
mid-gain
channels. switch from the all channels. mid-gain
Repeat state (e) Range i Mark
the measurement to the high gain Accuracy the with 1000/14 the
for gain state on
Measure state
foot
range
mark
signal
in
the
OFF
Digital
Voltmeter. on the Range Rate Start Test switch
2
Turn the Simulator START record ment times minimum which
Function to MARK
Selector switch TEST. Turn the
to
SWEEP. When the 1,000/14 foot mark the Electronic Counter indication. ten times. the between primary Interrupt series each voltage of the primary Allow during tests.
lamp illuminates, Take this measurevoltage the data point sheet. two at thirty seconds
input
measurement. is interrupted
Indicate on the
E-8
Measurethe 1,000/Foot RangeMark signal with the Digital Voltmeter and record the results on Data. 4 Perform measurements another [specified] primary at voltage. Checkthat the tracker remains locked at the I000 ft. deviation rate changeat the two primary voltages specified. (f)
Noise Set and the Ripple channels SC SC FIL FIL doppler Set to the CHAN CHAN signal the following: 3 4 SC SC FIL FIL and levels return on pens on controls. a minimum Record data
recorder CHAN CHAN I 2
Apply signal channel channels Permit of 60
the
simulated given.
frequencies gain
levels
recorder
3 and 4 at 50 I, 2, 3, and
MV/LINE. Zero the recorder 4 with the OFFSET-SC FILTER
the Recorder seconds after excursion
chart drive motor to run for this condition is obtained. of the Recorder pens on the
the maximum sheet. (g) CRO i Logic Turn RADVS LAMPS ON. between tion of sheet. on
Sisnal the
Accuracy Recorder switch
in chart
the
Search
and
Track Turn
Modes the to time the data
drive When BURN Burn Record
motor. DI OUT Out the and D3 SIGNAL
POWER When
to ON. turn lamp of the the
TRACKER switch the and on
illuminate, the the CRO indication CRO
illuminates,
measure Signal results
indica-
signal.
Measure the CRO DVS signal and record the results on Increase the frequency of lamp illuminates, measure the CRO D2 Tracker DVS signal.
with the Digital the data sheet. D2 to 1.5 KH z. the time between
Voltmeter
When the RODVS indication of
Lock signal and indication of loss of the Record the results on the data sheet. D2 until the TRACKER LOCK the condition of the CRO data in sheet. the the "OFF" condition on the with data D2 DVS
Decrease the frequency of lamp extinguishes. Record signal Measure (ON the or OFF) DVS on the
CRO
signal
the Digital sheet. (h) Cross-Coupled i
Voltmeter.
Record
results
Side
Lobe
Reiection
Set up STC No. 4 on the RADVSTE with D2 at 3100 H z. Record the actual frequency of DI, D2, D3, and D4 measured with the Electronic Counter on the data sheet.
E-9
2
m
Increase D3 TRACKER
the
RF
signal lamp
level
on
Beam
2 until Record
the the D2
LOCK
extinguishes.
attenuator out on the between cord 3 Change LOCK which data the the
reading at which the data sheet. Compute attenuator on the of reading data D2
D3 tracker drops the difference and -I00 dbm and re-
results
sheet. the D3 TRACKER
the lamp the sheet.
frequency illuminates. D3 TRACKER Compute
until
Record LOCK lamp the
the frequency at illuminates on the between this the the
difference
frequency and 3100 data sheet. Adjust original 3100 Ha out. Record the drops out on the
Hz. the
Record the results on frequency of D2 toward
setting until the frequency at which data sheet.
D3 tracker drops the D3 tracker
Set D2 frequency at 3100 H z. level on Beam 2 until the D3 cord the level at which the data sheet. and -i00 dbm Set D2 signal the Values see record V = V = = 0.125 = 0.500 1.000 Table E-3.
Decrease the RF tracker acquires. tracker acquires between on the
signal Reon the
Compute the difference and record the results level at -i00 in dbm. Step
this level data sheet.
Repeat
measurements
_ using
Beam
3.
(4)
Required (a) (b) (c)
Test
Analog
outputs:
Sensitivities: Analog Vz = 50 ft. noise: 0.300 and 200
v p-p max
max, max @ @ simulated 2000 ft. simulated
v p-pXmax;YRz ft., see
v p-p
v p-p D
(d)
F.
Logic
signals:
Appendix and
Lunar (I)
Reflectivity D1 (a) Tracker Turn on ACTIVATE After out. tion (b) increase
Calibration
Freamp
Gain
State
Signals
the recorder chart drive switch to ON. Turn the RA the and DVS DI RELIABLE on attenuation Beam
motor. RADVS i until signal results and
Turn the TEST POWER switch to ON. LAMPS DI illuminate, tracker this data the on drops condisheet.
the
OPERATE
Measure the for reference. the Beam PI /0 °
reflectivity Record the signal the D1 level
under on the
Observe tion of
decrease just locks
attenuain the
I until
tracker
high gain state. Record the PI /0 ° signal level at DI lock-on on the data sheet. Record the attenuator setting at which the the VOLT tracker SEL locked on the 40 data and P1 sheet. 65 measure the reto a rerange sheet.
(c)
Turn
switch
to P1 Panel Turn the of in
signals with suits on the REFL (DVM) DI reflectivity flectivity conduct the obtainable.
the Display data sheet. and signal adjust level given over the the
Voltmeter. the VOLT VDC. test
Record the SELECT switch on If Beam be on i for of the range signal the data
attenuation 0.5 this
signals test Record
cannot
obtained,
greatest
reflecitivy setting
attenuator
E-10
(d)
Repeat signal results
the on
reflectivity of 1.0, the data
measurement 2.0 and 3.0
in v DC.
Step
(c)
for the
levels
Record
sheet. on Beam I until gain the DI PI preamp preamp the
(e)
Decrease just output
the just
attenuation to prior the at to and
switches
mid-gain after which
state. the Pre-amp
Record swtich. gain
Record switched.
attenuator
setting
Record the results on the data sheet. and PI 65 measurements for 65 db gain the results on the data sheet. (f) Repeat the reflectivity mid-gain state. Record Repeat mid-gain (h) Decrease just ments suits (i) in on the reflectivity Record measurement the results measurements the on db db results Beam gain gain in on
Repeat the P1 40 state and record
Step (c) the data Step the (d)
for the sheet. for sheet. tracker the the measurerethe
(g)
in on
state. the
data DI
attenuation to the 40 for 40 sheet. in on
1 until state. state.
the Repeat
switches
Step (e) the data
Record
Repeat the and record Repeat the and record Observe the
measurement the results measurements the results P1 / 0°
Step (c) the data
for the sheet.
40
db
gain
state
(j)
in Step (d) for the on the data sheet. level and increase
40
db
gain
state
(k)
signal
the
attenua-
tion on Beam I until the DI tracker _ust switches to the 65 db gain state. Record the PI / 0 slgnal level just prior to and after gain switch on the data sheet. Record the (I) attenuator the gain setting at in gain Step the switch (k) on for on the gain the data switch data sheet. to sheet. the
Repeat 90 db
measurements state. Record
results
(2)
Other Repeat
Trackers above steps.
(3)
Required The
Values measurements are taken for should calibration not exceed purposes 5.0 Volts
reflectivity
only. The for the 90
G.
reflectivity analog signal and 65 db gain states. Period POWER switch to Counter Timer ON. Record on the data
Modulation
Sweep
Turn the RADVS on the Universal
the sweep sheet.
period
indicated
E-If
APPENDIX F BUYERATREQUIREMENTS F LISTING Information in this appendix is reproducedfrom HAC document o. 3023926A, N
Surveyor is of to Spacecraft of A-21_ Table System No. In of Test 3-11-g, this Specification. "Test'Requirements table, requirements to by of specific numbers changes of Table This are cannot The first table, Table F-I, a reproduction the referenced without of the same Library," are arranged phases. the second (pp. 155-179)
document. indication requirement column Table
according Different column.
number
applicability are denoted
test in in
aspects The
dash
revision The
letter
allows F-2 of
a means is a the
showing
requirements. "Test shows of Requirein the ap-
second
table, (pp.
reproduction
3-12-g, table in terms be
ments which
Matrix," phases dash of
427-429) test
referenced is evaluated. places
document. Entries tests
each
requirement (X's
plicable because In listed. for
numbers.
indicate
where
conducted
conflicting flight
configurational acceptance are the only. of test
requirements.) requires the passing Tests of (SRT) every test requirement which are
general, The
exceptions
System
Readiness
subphases,
operational Certain
convenience other details the
requirements document.
and
phases
are
given is
below carried
with over
use from
of
excerpts source INITIAL 3.3.1
from
referenced
(Section
numbering
the 3.3
document.) SYSTEMS Test I. CHECKOUT (ISCO) TEST PHASE
Objectives Perform calibration of as required to support and the flight mission. performance test phases. grounding of each tests which cannot be Perform made in Perform Verify craft spacecraft subsequent power TCM and compstibility subsystem. design engineering and data channels this and subsequent test phases
2. 3. 4. 5.
checks. subsystem with the space-
Provide for (i) special tests to verify new features, and (2) interface margin tests. The spacecraft of which shall in These FC/TCM, are such test and
3.3.2.1 Test Description: divided into test groups, each with the telecommunications if the first of the any, will be MS-MA/TCM, RF/CD/SP The with test actions, gration, (TCM) formed tion
shall be functionally be tested in conjunction a manner groups to as that are: (After shall at be the inteperdiscrethe telecommunications mutual inter-
equipment revealed. TV/TCM,
PO-RF/CD/SP,
FC-AM-RA-PR/TCM. requirements in any order
equipments PO-TCM test
referred test tested
subsystem.
integration groups
remaining director.)
F-I
The abbreviations are explained as follows: CD: SP: RF: FC: AM: RA: PR: MA: PO: MS: TV: 3.3.3.1.1
access test to the Test
Cormnand Decoding Signal Processing Radio-Frequency ata Link (or Radio Communications) D Flight Control Altitude Marking Radar RadarAltimeter andDoppler Velocity Sensor Propulsion EngineeringMechanisms Auxiliary Power Mechanical Subsystem Television
Test Signal section tees shall be provide and provided for by shall direct to test be electrical satisfy the cables. on an injection shall The monitoring
Access: of this
spacecraft.
requirements
3.3.3.1.2 external 22 3.3.3.2 ditions. ment 3.4
Power Requirements: volt DC source. Environment: All air tests
spacecraft
operated
shall shall the
be
performed be provided ....
at
room
ambient
con-
Sufficient
circulation below
to maintain
equip-
operating MISSION 3.4.1
temperature
maximum
SEQUENCE/ELECTROMAGNETIC Test Objectives: Interference i. Verify that The Test the
INTERFERENCE objectives shall be of to: in the
(MS/EMI) Mission
TEST
PHASE
Sequence/
Electromagnetic
system
performs
accordance
with
the
System Functional Requirements Specification 224510, and Equipment Specification 224832, when commmnded through all environment. 2. Verify the modes of operation in an ambient of the laboratory Surveyor simulating Launch Pad 36. by
functional radio to b_
compatibility frequency pncountered
spacecraft with the environment 3. Verify that the compatible with the Atlas/Centaur 3.4.2.1 Interference compressed (66 hour) two Mission and voltage into the Test Descriptions: Test Phase shall
interference at AFETR is
Surveyor
spacecraft
functionally
the expected RFI launch vehicle
environment created and its AGE.
be
The Mission Sequence/Electromagnetic divided into (I) and (2) plugs Tests i shall Each and (3) plugs out, Interference Test. have test a constant voltage/time sequence shall supply
in,
Time
(32 hour) Mission Sequence Mission Sequence/Electromagnetic Sequence sequence actual Tests, 2 shall battery segments: sequence have voltage.
real time Of the first supply which divided profile be
power
a power
approximates
following SRT : P/L-L: INJ : C_01 : MC : C_02: TD: POST TD :
System Readiness Test Prelaunch to launch Injection and attitude reference acquisition Coast phase i Midcourse correction Coast phase 2 Terminal descent Post-touchdown F-2
3.4.3.1.1
test during Post access the shall and third objectives
Test be test shall
Access: provided sequence, be performed
For as When the access by r-f
the
first to
two
sequences, with phase sequence in a
hardline the test is true reached flight through
necessary simulated remainder the with
comply Injection of the
requirements.
Touchdown
spacecraft
configuration of no with the spacecraft 3.4.3.].2 simulated sequence following utilize Power
hardline operated
(I00 percent link. During test
plugs-out
configuration)
Requirements: will be battery test power.
sequence
i, a +
19V test
battery voltage 2, the simulated levels spacecraft .... battery
applied voltage sequence
to the spacecraft. shall be adjusted 3, the spacecraft
During for the shall
During
3.4.3.2 formed third the EMI vehicle in an sequence,
Environment: earth the EMI ambient environment until be spacecraft
The
first shall
two be
test to 36 is
sequences the in and EMI a r-f the
shall test. screen At that
be
perthe where Launch the shall to
environment of launch off and
prior pad phase the
During room time,
located
expected is simulation
Atlas/Centaur of the be sequence allowed
simulated shall
Injection
reached. shall
turned
remainder levels
be performed. The EMI simulation stabilize before initiating the
intensity test. TEST
3.5
SOLAR
THERMAL 3.5.1 shall
VACUUM
(STV)
FUNCTIONAL The
PHASE of the Solar Thermal Vacuum
Test
Test be: i.
Objective:
objectives
Verification
of
correct
spacecraft
functional
operations
during a real-time to a range of solar space 2. during 3.5.2.1 shall be Test environment. of Verification
transit mission conditions in spacecraft
sequence while exposed a simulated cislunar thermal performance
correct STV During
simulated
environments. the STV The test flight test phase phase This the shall test spacecraft as defined of consist
Description: in accordance 224550 and
tested
with
the 224555.
mission
program
by HAC specifications 3 subphases. i. shall under
Subphase
A Low
Temperature
Test.
subphase sequence Constant be propower
consist of a 66 hour real-time mission simulated transient and low level Solar A one hour solar eclipse shall shall test
environments. vided from 2. shall during its own consist from
the test. batteries. of a 66
The
spacecraft Test. real-time A This
derive subphase sequence the Solar space-
Subphase continued
B High
Temperature hour subphase chamber
mission interrupting high
A without operation.
Thermal-Vacuum
level The
Constant environment craft derives power 3.
shall be simulated. from the STEA. a plugs-out nominal
Subphase C Nominal temperature test shall be conducted under ment. shall board from Hardline access and test to the be minimized power. The mission through period, assessment. F-3 launch
test. This STV environfor this test from onas a 32-hour operation and a a temperature
spacecraft operated conducted followed terminal
spacecraft shall be
compressed stabilization postlanding
sequence, midcourse, real-time
involving
real-time by descent,
In each of the subphases,the test sequence consists of the following segments. SRT: SystemsReadinessTest MSSEQ (DRYRUN): Mission Sequence-Dry un(Subphases and B only) R A P/L-L: Prelaunchcountdown and launch INJ: Injection C_I: CoastPhaseI MC: Midcoursecorrection C_2: Coast Phase2 TD: Terminal descent POST TD: Post Touchdown SRT: SystemsReadiness Test (Subphase only) C 3.5.3.1.1 Test Access: Hardline test access to the spacecraft
shall be provided through the vacuum chamber penetration plates. During the final test subphase (C) this access shall be minimized to include only spacecraft power access for with cormnunications derived 3.5.3.1.2 utilizing Power emergency solely be shutoff and RF link. Test and mission subphase thermal instrumentation
Requirements: battery power, Three Vacuum to a during
A and B run tests During on
C shall ground shall these be
be
run
spacecraft
subphase sequence
power. conthe the of
3.5.4.2 ducted under spacecraft conditions the mission. or -300°F
Environment: a Solar shall to be The Thermal be subjected encountered simulated a static
environment.
tests
simulated environment approximating all phases of the transit portion shall 5 x consist or of a temperature and solar of 10-6torr, less,
environment pressure
lower, 0.8, (A
radiation of respectively. plane. )
I.i, and 1.0 solar constant for subph_ses A, B, and solar constant is defined as 130 w/ft at the test
C,
3.7
VIBRATION
(VIB)
TEST
PHASE The objective of the Vibration and after of the Test Phase
3.7.1 Test shall be to: I. 2.
Objectives: Verify launch Verify frame
functional vibration proper and all
integrity during environments. and components. Vibration Test
simulated space-
fabrication system The
assembly
3.7.2.1 into two
Test basic i. 2.
Description:
Phase
shall
be
divided
parts: Vibration Environments Earth Ambient Environment alignment Only the tests before or (spacecraft after functional and
exposure and
to vibration). Function checks. Post-
Functional/Pretest RADVS Hardline and along test with
Checkout positional shall this
test 3.7.3.1.1 meeting
Checkout Test test
concerns Access: objectives
access of
be test
minimized phase.
while The
the
requirements
testing shall be accomplished No con_nands shall be sent to 3.7.3.2 conditions. Environment: Vibration All levels
primarily the vehicle tests are
in a plugs-out test configuration. during the shake periods. be performed at room ambient
shall
specified...
F-4
3.8 VERNIERNGINEIBRATION E V (VEV)TEST PHASE 3.8.1 Test ObJectives: The objective of
Vibration Test Phase shall the of RADVS be to: beams do not Verify a result 3.8.2.1 Test shall be divided i. 2. 3.8.3.1 assembled Among
the
Vernier
Engine
produce
a
false
lock
as
vernier
engine
vibrations. Engine Vibration test phase
Description: into two basic Flight tion
The Vernier parts: open tests The in
control/RADVS environment. functional
loop
operation and
in after
a vibravibra-
Spacecraft tion test.
before spacecraft a flight
Spacecraft mecahnically exceptions i. 2. 3. The The
Configuration: and electrically shall retro be the
shall be fully configuration...
required
following: shall not shall be be not installed. be mounted with on
inert
rocket
altitude
marking
radar
the spacecraft. Fuel and oxidizer gas to minimum i0 + 5 PSlG differential
tanks
shall
filled
Helium
inside the bladder with 2 PSIG across the bladder, positive shall be removed and re-
4. 5.
pressure inside. Thrust Chamber Assemblies placed RADVS with equivalent feed horns shall
masses. be terminated
in microwave for the
loads to simulate a free space RF transmitters and receivers. 6. 7. The The and ASPP shall be in the transit omni spacecraft B shall be legs and extended. access and the with
environment position.
directional
antennas
A
3.8.3.1.1 minimized
Test to meet
Access: the test
Hardline objectives
to STEA
the
spacecraft The the
shall omni
be directional
requirements. through
spacecraft
shall be operated antenna RF command
in conjunction link.
3.8.3.1.2 Power Requirements: battery power during vibration use external ground power. The
The spacecraft testing. Pre-and
shall utilize Post-vibration
on board tests
shall
3.8.3.2
Environment:
spacecraft
shall
be
mounted Pressure, Vibration
on
the
system
test and shall
stand, utilizing vibration humidity conditions shall be as specified by the
isolation airmounts. be laboratory ambient. subparagraphs. shall be applied
temperature, environment
following Vibration
3.8.3.2.1 dummy axis. litude
Vibration:
simultaneously
through
vernier engines The excitation (band at each limited dummy
in a direction force shall be between 84 vernier engine
parallel to random noise and 20 2000 Ibs of RMS.
the spacecraft roll (Z) having a gaussian discps) and an average amp-
tribution
cps
3.8.3.2.2 jected to 3.8.3.2.3 force within
the
Period of vibration Tolerance:
Vibration Exposure: The environment for a period Spectral density cps of shall the in
spacecraft shall of 240 seconds. summed general and be
be
sub-
averaged maintained
RMS
input between 84 cps and 2000 +3 db of their nominal level.
F-5
3.10 AIRFORCE EASTERN RANGE TEST (AFETR) EST T PHASE 3.10.1 Test Obiectives:
i. Perform is 2. 3. 4. ready (J-FACT). Demonstrate vehicle Perform are during J-FACT for that and the spacecraft and to on and check launch pad launch critical pad in and compatible balance, flight. alignment, transport checks subsystem for and system test to verify spacecraft Test a Joint-Flight Acceptance Composite
weight,
functions prior Verify spacecraft perform functional tion for launch.
to encapsulation. is ready for and
operational
prepara-
3.10.2
General 3.10.2.1 Test test Description: test The subphases AFETR test phase in shall the order be comof the
prised following
of
nineteen brief
separate
performed
descriptions: Tests: subsystem and This level cannot PVT-4; did for launch (VPS) tests, be test subphase which adequately test subphases damage Acceptance shall system perform Control leak spacepad a shall are verify vital at a shall in ship-
i.
AMR-FC-SP-Subystem performance to mission system level. PVT-2, that to AFETR Test Functional system and the PVT-3, the and with and of success
parameters tested These not
2.
PVT-I, verify ment Composite
and ready the System
spacecraft is
suffer
any
a Joint-Flight vehicle. test and low subphase Jet Gas tests.
3.
VPS
Leakage:
This
Vernier (GJAC) tests, 4.
Propulsion high
Attitude
functional
pressure
pressure
decay
SRT (Post Encapsulation): Test to demonstrate craft is adequately prepared for transfer to after encapsulation. (LP): prior A to system start Readiness check of test checks of the J-FACT. shall be performed and confirmation Test shall be SRT link
that the the launch performed the as
5.
system 6.
functional
spacecraft
via
telemetry
CD (LP): Countdown system operational be placed J-FACT: shall and compatible flight and in The in
to provide that system
can
7.
launch configuration prior to start of Joint-Flight Acceptance Composite Test that Centaur During and be the spacecraft system this and launch test, the be a simulated readiness subphase, shall retro-rocket the final
J-FACT. subphase are
demonstrate thru Align:
vehicle countdown, initial performed. installa-
retromaneuver. test with omitted. shall be spacecraft All critical checked after Facility. cannot be verifications
8.
Weigh Those tions
spacecraft and
alignments fueling test shall
requirements This
associated subphase
9.
PVT-5:
testing at the functions shall spacecraft
Spacecraft Checkout be verified which
encapsulation.
F-6
i0. ii. 12.
13. 14. 15. 3o10.3.1.1
be of provided the test as
WB&A, Fuel Load, Pressure: Final Weight, balance, and alignment after retro rocket installation and fueling operations shall be performed during this test subphase. PVT-6: This test phase shall consist of connector pin retention tests to demonstrateconnector mating integrity, squib circuit, verification, and SSandAD checks. SRT(Post-Encapsulation): A SystemReadinessTest shall be performedduring this test subphaseto deomonstratethat the spacecraft is adequately prepared for transfer to the launch pad after final encapsulation. SRT(LP): A SystemReadinessTest - (LaunchPad) test shall be performed to verify that the spacecraft is adequately prepared to be launched. CountdownLP): A countdown(LaunchPad) test shall be ( performed to allow final spacecraft operational checksand to place the spacecraft systemin a launch configuration. SRT LP-Final and CDLP-Final: Same tests as SRT(LP) and CD(LP) which are performedat the appropriate time in the launch vehicle countdown procedure.
Test Access: from matrix, Hardline test Table test F-2. Ground power and test requirements Table F-2. and spacecraft set forth spacecraft battery in the power AFETR access and set RF link in control the AFETR shall zone determined requirements forth
requirements
3.10.3.1.2 Power shall be provided as zone of the Test requirements. 3.10.3.2 shall either operating be case
Requirements: determined from Matrix,
Requirements
configuration
Environment: encapsulated air below sufficient temperature
During or the on
the a
various test stand shall
test in be
subphases room ... to
the
spacecraft conditions. equipment In
either
ambient
conditioning maximum
provided
maintain
specified
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F-26
APPENDIX DISCUSSION OF THE
G STEA SIGNAL
S IMULTATION
TECHNIQUE
I.
INTRODUCTION The simulated of the equipment developed oscillator return uses from is signal a provided transmitted method.) oscillator with the in accordance return signal transit the by STEA signal, In or open from with signal is obtained as shown loop by in tests, in single-sideband Fig. the 5-1. (Some varisimulamotion. a (The trackS/C.) modulating loop
modulation unit signal able tion, In both single ing is An test is the
spacecraft's
similar For
a crystal controlled the
a manually computed spectrum
operated simulated is
frequency cases,
oscillator. therefore,
tests
spacecraft
closed
simulated
essentially
line which tracks the transmitted delayed, of course, by the signal actual return signal also tracks
frequency variations. time between STEA and frequency delay delay short of
transmitted the transit propagation This by the
variations, using from STEA is high altiappreciable frequency modulation
but with a much less than about tudes effect of the is on in the the
greater delay. For instance, 10 -7 seconds while the actual range same of 10 -5 of -10 -4 problems seconds. caused to seriousness
difference
has
an
transmitter effects
term
incoherence. altimeter
The
situation
pertains
nonlinear
klystron. differences and echo time will between rather actual closely signals The match and the simulated power ones also gain exist spectrum pattern, lunar scatde-
Important in of both a true lunar
spectral
characteristics. is nearly a random, the
long-term two-way
density
antenna
while the echo will
simulated signal spectrum also fluctuate in time in of the rough
a single noise-like simulated
line. The expected manner due to the signal is essentially
tering properties terministic. The quency to II. main
surface;
purpose and have spread
of
this already
Appendix been
is
to discuss affect in
how
differences results. VI.A.3 and
due
to
fredue Id.
coherence
nonlinear
modulation
testing Sections
Problems VII.A.
doppler
discussed
TRANSMITTER Undesirable
INCOHERENCE transmitter spectrum. "coherence frequency fluctuations result in a spreading of the
mixing-product filters, called
This spreading can loss," and possible
cause signal false locks the one can
power loss in subsequent and tracking errors. result delayed be of mixing in time is as two and sig-
To determine the seriousness of the effect, nals from a sinusoidally frequency-modulated shifted perturbed in frequency by the (by doppler). Assuming these sinusoidal modulation,
consider source, that the signals
doppler
shift expressed
negligibly
et(t,_c)
= E 1 cos
[_ct
+
# sin
_rt]
(G-l)
and
er(t,_c)
= Ket(t
- T d ' _ c + _d )
= KE I cos
[(_c
+ _d ) (t
- Td)
+
_ sin
_r(t
-Td) ]
(G-2)
G-I
where K = a constant, c= transmitter carrier frequency, _d= doppler frequency shift, r = frequency of the modulating sinusoid, = modulation index of the transmitted signal, Td= time delay betweentransmission and reception. The low frequency component the mixing of these two signals is of _r Td e3 = E3 cos [_dt + 2_Isin --_-]sin _r(t - Td + _)] whereE3 is a constant, and _ is a constant dependenton _rTd [75, p.89]. The one-sided Fourier spectrumof the waveformin eq. (G-3) is composed of lines at frequencies led _ n _rl, n=O,_ i, _ 2..... with amplitudes proportional to the Bessel functions r Td Jn(2_Isin--_I). The power level of eachcomponentelative to the total signal power is, therefore, r Sn = 20 log [I Jn (2_Isin _rTd I)I] _ decibels (G-4) (G-3)
Representative numerical values for (G-4) will be obtained for typical causesof frequency incoherence. Thesenumbers directly indicate the magnitudeof the incoherenceproblem, which would go unnoticed in STEA type simulation testing. A. Power Supply Ripple
for The maximum allowable sensitivity the DVS klystron is I00 kHz/volt is ripple that component likely will to be I0 times occur the normally applied to anode voltage supply variations specified [56]. (The RA klystron, being a reflex sensitive.) at the value, the DVS The major component 2.4 of index is of kHz, the power or major converter frequency,
klystron, supply twice ripple
more either
frequency.*
Choosing
latter
modulation klystron
to a maximally
sensitive
i00 =
x 4.8
103 x
V x_2 = 103 30 V (G-5)
where ripple
V
is the rms value of the ripple component components are assumed to be negligible°) minimum that value the of total V for return which false is lock 28 db
at
4.8
kHz.
(The
effects
of
other
The by
could above
occur the
is
easily
computed level; any
assuming
signal
acquisition
Later
model
KPSM's
might
operate
at
3.8
to 4.0
kHz.
G-2
higher signal would causea preampgain switch to effectively band signal by about 25 db. The solution to 20 log [ Jl (60V)] = -28 db
suppress the side-
is V . = 1.33 my. (Such a situation could occur at a beamslant range of 50 kft and mlnangle of incidence of about 20 off the lunar vertical. an ° The corresponding spacecraft slant range could be anywhere between46 kft and 90 kft.) A more serious problem occurs whenthe ripple is high enoughfor lock of a sideband to persist an appreciable time. For example,a i0 mv ripple is sufficient to keep the upper beamof a spacecraft at 25° attitude locked over slant ranges of 50 kft to I0 kft, where gain state switching would occur. (The after-burnout sensitivity wasused for this computation.) As another example,one of the worst situations involves the upper beamof a spacecraft at 5° attitude. A ripple of about 8 mvbefore burnout or 2 mvafter burnout would be sufficient to maintain lock on the first sidebanddownto about 20 kft. Some other levels are shownin Fig. G-I.* It can be shownthat the first sideband levels (in db) relative to the acquisition threshold vary approximately as 20 log V and independently of range for situations of interest. A consequence the independence of toward range is that false locks will not normally be broken unless gain states are switched or appreciable attitude changeoccurs° The amountof power lost due to sideband generation should also be considered. In the case of the DVSklystron, the fundamentalcomponent reduced less than one is db for a ripple of less than about 16 mv. This amount,of course, is not serious. As a final consideration of ripple, it should be noted that there is no pertinent test requirement specified. The vendor test simply requires that ripple be recorded. It is not measured anywhere else nor are its frequency modulation effects observed. B. Vibration
The frequency between expected probably that eq. G-4 vibration modulation I0 Hz and g levels, be is 2_r,_ --_--)J less [61]. sensitivity must not 2 kHz [56]. therefore, than For 1.5 kHz cases specification exceed 200 kHz for the DVS klystron for 25 is that the at the would within to peak-to-peak g vibration deviation modulation most likely approximation
A reasonable value for frequency would be i kHz. The frequency of because of mechanical in this resonances range, are interest a good
region
Sn =
20
log
_[Jn(2_f
(G-6)
Computations following
are
based
on
an
expected
return
power
of
-94
dbm
at
50
kft
and
the
acquisition
thresholds: Typical Sensitivities BBO ABO -118 -117 -116 -114 -112 dbm
R 50 40 30 20 i0 kft
-iii -Ii0 -109 -107 -105
dbm
G-3
GAIN
STATE
1 AFTER BURNOUT Ist SIDEBANDS RIPPLE
SWITCH
+i0
I BEFORE BURNOUT
i0 MV
2
r./'l
Ir II
0 ACQUISITION THRESHOLD
-
o
I_
-_P"
B_J_OB--_q_---T
_
2n---'_ SIDEBANDS
5"-"
_I -i0 I r.
Z
_K_TER
BURNOUT
--
i
_
i0 MV
RIPPLE
f
BEFORE
BURNOUT s SIDEBANDS
< _
m
-20
t
i MV ist AFTER BURNOUT
RIPPLE SIDEBANDS
I MV
RIPPLE
I i0
I 20
I 30
! 40
I 50
SLANT
RANGE
(kft)
Fig.
G- i.
Sidebsnd threshold DVS beam vertical.
levels
relative
to
the
signal
acquisition
vs roll axis of a spacecraft
slant range for the upper at 5 ° angle to the lunar
G-4
O
0
FUNDAMENTAL
O
_'-10
_
M > M
-20
0
20
40 BEAM SLANT
60 RANGE (kft)
80
i00
Fig.
G-2.
Signal deviation to about
component at 1.5
levels
vs.
range of
for 1 kHz up
sinusoidal
frequency modulating kHzo
modulation
frequencies
G-5
where flf is the frequency deviation, c is the velocity of propagation, and r is the beamslant range. (Notice that eq. G-6 is independentof modulating frequency.) Equation G-6 is plotted in Fig. G-2 (for n =0, i, and 2) for the caseof I kHzmaximum deviation. The high altitude effects are similar to those caused by about 50 mvripple° As mentionedin the discussion of ripple effects, situations can be found for which false lock can occur on any sidebandwithin about 28 db of the total power. Loss of powerin the fundamentalcomponents also i seen to be a problem for the vibration case. In fact, if the frequency deviation were greater than about 2 kHz, the fundamentalwould disappear completely at some beamrange below 90 kft. C. EMI High frequency EMIis not likely to causeproblemsbecauseassociated modulation indices would probably be low. The contrary is true for low frequency EMI, however. (Suchfrequencies con_nonly rise from converters and conmmtators.) a Also, shielding against these lower frequencies is generally found to be moredifficult. No special casesare considered here becauseEMIeffects can cover a wide range. For inputs which essentially have a single frequency, the results would be similar to those described in the preceding sections. EFFECTS OFNONLINEAR MODULATION OFTHE RAKLYSTRON Common sweepnonlinearities can often be adequately modeledby addition of a quadratic term to the frequency function. For example,supposethe transmitted function is et(t) = EIR(t)* (i#(t)cos where E1
LD O
III.
[(_o-
mt + ,_ t2)t]}
(G-7)
= a
constant, center rate, of the quadratic defined -T + 2 nonlinearity, as frequency,
= undeviated = linear _ W(t) sweep
= coefficient = = the "window"
function < t < ---
i for
-T + _ 2
= 0 elsewhere R(t) = the "repeat" function defi_ed as
Z g (t m= - oo T m sweep = *
- mT)
repetition period,
period,
"flyback" indicates
convolution
G-6
(The shapeof the waveformduring "flyback" is not important to this example.) The ideal return function would then be er(t) = Ket(0_ + COd, - Td) ° t (G-8)
where_d and T have the same meaningsas in previous sections and K is a constant. Consequently,_he low frequency mixing component after blanking) would be ( eB(t) = E3R(t)* {B(t) cos [eat + 2mrdt + 3_ Tdmt- 3_ Tdt2 + _]_ (G-9)
where E_ and _ are constants and B(t) is a "window"function which provides the blankin_ effect. The third frequency term in eq. G-9 represents the steady range error due to the assumed nonlinearity. The percent error is
150 • =
m
¢_ T d 7o (G-10)
This equation simulations; would actually
shows one of the indicated exist over
the effects error would most of the
of be
the small about i00
delay times
factor (Td) in STEA less than that which
descent. 180 Hz It can IV, Ch. apart by be shown 2] the R(t) that the
term
The spectrum of with an envelope width for
e_(t) is composed of lines spaced _etermined by the other factor. case is approximately [76, pt.
spectral
this
6_
W _"
TdT s 211
(G-11)
where
T
S
= sweep
period
after
blanking. the spectral Then, from width, (G-10) 0.01m 150 suppose _T d is such a value as to
To produce
obtain a number for a 0.01% range error.
_Td 0.06m and W = T
S
=
2_( 150) If T
S
is
roughly
5 x
10 -3
seconds
and
m
= 2_
(8
x
108 )
(at
low
deviation),
then
W
=
1.6
kHz. At the same time,
This the
shows
how
the width
true would
spectrum still
can be
easily
become
quite
wide.
simulated
very
narrow.
G-7