NATIONAL TRANSPORTATION SAFETY BOARD
OFFICE OF AVIATION SAFETY
WASHINGTON, D.C. 20594
September 25, 2008
POWERPLANTS GROUP CHAIRMAN’S FACTUAL REPORT
NTSB ID No.: DFW08MA076
Location: Oklahoma City, Oklahoma
Date: March 4, 2008
Time: 1515 central standard time
Aircraft: Cessna Citation 500, N113SH, Southwest Orthopedic and Sport Medicine
B: POWERPLANTS GROUP
Investigator-in-Charge Timothy J. LeBaron
National Transportation Safety Board
Group Chairman: Gordon J. Hookey
National Transportation Safety Board
Member: John F. Dargin III
Federal Aviation Administration
Member: Jan R. Smith
Technical Advisor: Paul F. Crosby
Pratt & Whitney of Canada
Bridgeport, West Virginia
On March 4, 2008, about 1515 central standard time, a Cessna Citation 500 airplane,
N113SH, was destroyed upon impact with terrain following a loss of control shortly after takeoff
from the Wiley Post Airport (PWA), Oklahoma City, Oklahoma. The airplane was equipped
with two Pratt & Whitney of Canada (PWC) JT15D-1A turbofan engines. The two pilots and
three passengers were fatally injured. The airplane was registered to Southwest Orthopedic and
Sport Medicine Clinic of Oklahoma City, Oklahoma. Visual meteorological conditions
prevailed and an instrument flight rules flight plan was filed for the flight from PWA to Mankato
Regional Airport, Mankato, Minnesota.
The airplane’s engines were partially disassembled in the presence of the Powerplants
Group. Most of the left engine’s fan blades were fractured adjacent to the blade root platform
with the fractured ends bent opposite the direction of rotation, but there were four fan blades that
were full length that were also bent opposite the direction of rotation. The left engine’s fan exit
vanes and compressor inlet vanes were packed with finely chopped wood. The left engine’s high
pressure compressor (HPC) impeller vanes were broken and bent opposite the direction of
rotation. The right engine’s fan blades were all in place in the disk, but were bent rearward
against the fan exit vanes and compressor inlet vanes. There was a quadrant of fan blades that
had soft body impact damage that was coincident with a cluster of fan exit vanes that were
splayed apart. The right engine’s HPC impeller vanes were also broken and bent opposite the
direction of rotation.
D: DETAILS OF INVESTIGATION
1.0 Engine information
The PWC JT15D-1A engine is a dual-spool turbofan that features a one-stage fan, a one-
stage centrifugal flow HPC, reverse flow annular combustor, 1 a one-stage high pressure turbine
(HPT) that drives the HPC, and a two-stage low pressure turbine (LPT) that drives the fan. The
JT15D-1A engine has a takeoff thrust rating of 2,200 pounds at standard day temperature and
pressure conditions. 2
1.2 Installed engines
The engines installed on the airplane were PWC JT15D-1A turbofans. The engine
identification data plates were not found in the engine wreckage. According to the airplane’s
maintenance records, the left engine was serial number (SN) PCE-77179 and the right engine
was SN PCE-76084. The left engine was confirmed as being in the left-hand position by the
A reverse flow combustor has the airflow change direction so that the airflow and combustion flows from back to
Standard day temperature and pressure conditions are 59° Fahrenheit and 29.92 inches of mercury, respectively.
engine mounts that were located on the right side of the engine. 3 The right engine was
identified by process of elimination.
2.0 Left engine
The recovered left engine extended from the inlet duct to the exhaust duct. The engine
did not have any indications of an uncontainment or case rupture. The engine mount structure
was on the right side of the engine. There were pieces of cowling on the upper right forward part
of the engine. The pieces of cowling had random burn and soot patterns on the exterior and
interior surfaces. (Photo 1)
Photo 1: View of left engine.
The inlet duct was in place on the engine, but was crushed back against the low
compressor case between 4 and 9 o’clock. The inlet duct did not fluoresce when illuminated
with an ultraviolet (UV) light. 4
The low compressor case was attached to the intermediate case forward flange between
10 and 2 o’clock. The remainder of the low compressor case was pushed rearward so the low
compressor case’s rear flange partially overlapped the intermediate case. The low compressor
case was intact. The low compressor case had circumferential rub marks in the fan blades’ plane
of rotation between 3 and 8 o’clock that corresponded to the circumferential rub marks on some
of the fan blade tips.
All references to position or directions, as referenced to the clock, will be as viewed from the rear, looking
forward, unless otherwise specified.
Organic proteins such as bird remains and blood will fluoresce green when illuminated with a UV light.
The fan hub was intact and still in place. All of the fan blades except four were fractured
transversely across the airfoil directly adjacent to the blade root platform and the roots remained
in the fan hub. The fracture surfaces on the broken fan blades were all coarse and at an angle to
the airfoil surface. Most of the fractured ends of the broken fan blades were bent opposite the
direction of rotation. There were four full length fan blades in a continuous sector that also
remained in the fan hub and the airfoils were bent opposite the direction of rotation. (Photo 2)
Several broken fan blades were recovered from within the fan case. The full length fan blades
and the broken fan blades had numerous nicks and dents on the leading and trailing edges. There
were several fan blade tips that had circumferential rub marks, heat discoloration, and material
displaced opposite the direction of rotation. When the full length fan blades and the broken fan
blade pieces were illuminated with a UV light, only two broken fan blade pieces fluoresced. The
fan rotor could not be rotated.
Photo 2: View of left engine’s fan rotor showing blades broken adjacent to root platform
and four blades bent opposite the direction of rotation.
The spinner was still attached to the fan hub, but was partially crushed against the hub.
There was material that appeared to be wood and smelled of jet fuel trapped in a fold of the
partially crushed spinner that fluoresced when illuminated with a UV light.
Almost all of the fan exit vanes were pulled out of the inner and/or outer shroud slots.
The fan exit vanes did not fluoresce when illuminated with a UV light. The compressor inlet
vanes were all in place, but had numerous nicks and dents on the leading edges. The compressor
inlet vane area was packed full of coarsely chopped wood material.
The intermediate case was cracked and broken. The forward inner bypass duct was in
place on the intermediate case and the aft inner bypass duct was attached to the forward bypass
duct. Both inner bypass ducts were buckled around the entire circumference. The inner bypass
ducts had random burn and soot patterns on the inside of the duct on the upper right side of the
engine. The unburned, unsooted inside areas of the inner bypass ducts did not fluoresce when
illuminated with a UV light.
The HPC impeller was intact. All of the impeller vanes had circumferential rub marks on
the forward parts of the vanes and had pieces broken off the aft section of the vanes. The
fracture surfaces on the vanes were all coarse and irregular shaped. (Photo 3) Almost all of the
impeller vanes had an approximately 1-inch long section in a circumferential ring in the center of
the impeller that was either bent opposite the direction of rotation or broken away. The HPC
impeller could not be rotated. The HPC impeller did not fluoresce when illuminated with a UV
light and there were no feathers or bird remains in the vanes. The HPC impeller shroud had
circumferential rub marks on the inner diameter.
Photo 3: View of left engine’s HPC impeller showing vanes
broken and bent opposite the direction of rotation.
The combustor plenum was partially cut away to facilitate examination of the diffuser
tubes and combustor liner. The visible diffuser tubes were in place and did not have any thermal
distress. There was one diffuser tube that had a ¼-inch by ¼-inch hole with the edges pursed
outward. The inside of the tube with the hole in it did not have any marks.
The combustor plenum was buckled adjacent to the front and rear flanges. The
combustor plenum did not have any thermal distress. The inside of the top of the combustor
plenum was sooted except where the diffuser tubes were touching the plenum. The combustor
liner was buckled and had been displaced off of the fuel nozzles, but the combustor liner did not
have any thermal distress. There were no feathers or bird remains around the combustor liner,
fuel nozzles, and igniter plug ports.
The 2nd stage LPT blades were all in place on the disk, full length, and did not have any
apparent damage. (Photo 4) The LPT rotor could not be rotated.
Photo 4: View of left engine’s 2nd stage LPT blades. (PWC)
The exhaust duct was crushed flat and the aft end of the duct was bent upward.
3.0 Right engine
The recovered pieces of the right engine extended from the inlet duct to the exhaust duct.
The inlet duct and fan case and rotor were separated from the remainder of the engine. (Photo 5)
The engine did not have any indications of a case rupture, uncontainment, or inflight fire
although most of the engine had indications of having been exposed to a ground fire.
Photo 5: View of right engine.
The inlet duct was twisted and torn and was crushed back against the fan case.
The spinner was partially crushed and torn. The spinner was recovered separated from
the fan hub. The spinner did not fluoresce when illuminated with a UV light.
The low compressor case was intact, but was buckled around its circumference and bent
inward slightly at one location. The fan case inner diameter did not have any circumferential rub
marks. There were several areas on the inner diameter of the low compressor case that had metal
The fan hub was intact except for an approximate 30 degree arc of the forward fan blade
retaining ring flange that was bent radially inward. All of the fan blades except three were full
length. There were three fan blades at locations No. 1, 5 14, and 17 that were fractured adjacent
to the blade root platform and 3 ½- and 3-inches above the blade root platform, respectively. All
of the fan blade airfoils were displaced axially rearward, with blades Nos. 2 through 7 being
pushed rearward more than all of the other fan blades. Many of the fan blades were pressed
against each other with the adjacent midspan shrouds shingled 6 or sticking through the adjacent
The fan blade fractured adjacent to the blade root platform was arbitrarily identified as the No. 1 fan blade. In
accordance with gas turbine engine industry convention, all of the remaining fan blades were numbered
consecutively in a clockwise pattern as viewed from the rear looking forward.
Shingled is the condition of the mid span shroud overlapping the shroud of an adjacent blade in lieu of abutting at
the contact surfaces.
airfoil. (Photo 6) Fan blades Nos. 8, 9, 10, and 11 had soft body impact damage 7 with the
leading edge curled back opposite the direction of engine rotation as well as the whole airfoil
bent back opposite the direction of engine rotation. (Photo 7) The fan blades did not fluoresce
when illuminated with a UV light.
Photo 6: View of right engine’s fan rotor. (Cessna)
Photo 7: Close up of right engine’s fan showing soft body impact damage.
Soft body impact damage is characterized by the large radius of curvature of the deformation to the blade,
typically a fan blade. Soft body impact damage can result from impacts with pliable objects such as birds, ice slabs,
tire rubber, and plastic objects.
The fan exit vane outer shroud was rolled radially outward in one quadrant. Most of the
fan exit vanes remained attached to the outer shroud although the vanes’ inner foot was pulled
out of the inner shroud and the vanes were pressed flat against the shroud. Most of the vanes
that remained were pressed against the outer shroud. There was a 6-inch wide group of about
eight adjacent exit vanes that were missing. These vanes were about 90 degrees away in the
counterclockwise direction from the fan blades with the soft body impact damage. The vanes on
either side of the missing vanes were bent away from the location of the missing vanes. (Photo
8) The fan exit vanes and compressor inlet vanes in the area of the missing exit vanes fluoresced
when illuminated with a UV light.
Photo 8: Close of right engine’s fan exit vanes showing splayed vanes.
The LPT drive shaft was broken about 3-inches forward of the No. 2 bearing. The
fracture surface on the drive shaft was on a very slight spiral. The broken end of the shaft was
also bent directly in front of the bearing.
The intermediate case was completely missing from the engine and was not received with
the remainder of the engine. The forward and aft inner fan bypass ducts were still joined
together and were buckled. The fan ducts were pressed against the core of the engine. The
bottom of the forward fan duct was burned away completely and the aft fan duct was burned on
the bottom from the front flange aft about 1-foot between 5 and 7 o’clock. There were several
areas on the aft inner bypass duct that fluoresced when illuminated with a UV light. The inside
of the outer bypass duct did not fluoresce when illuminated with a UV light.
The HPC impeller was intact. Most of the HPC impeller vane tips had circumferential
rub marks and scoring as well as small pieces missing. (Photo 9) The flowpaths between the
HPC impeller vanes were filled with a white ash. The HPC impeller shroud had circumferential
rub marks. There was a 1 ¼-inch long by ¼-inch wide diagonal hole in the impeller shroud.
Photo 9: Right engine’s HPC impeller. (PWC)
The combustor plenum was cut away to facilitate the examination of the diffuser pipes,
combustor liner, and fuel nozzles. The inside of the combustor plenum was sooted. Several of
the diffuser pipes were pulled out of the forward connectors. The diffuser pipes at the bottom of
the engine were crushed flat. The combustor liner was buckled in several places, but did not
have any thermal distress. There were no bird remains in the diffuser pipes or around the
The 1st stage HPT blades were all in place in the disk, the blades’ tips were curled
opposite the direction of rotation.
The 2nd stage LPT disk was intact. All of the 2nd stage LPT blades were in place in the
disk and did not appear to have any damage. The end of the LPT stub shaft had an approximate
60 degree arc of a circumferential rub mark that corresponded to burnish mark on the emergency
fuel shut off (EFSO) valve actuation button. The LPT stub shaft had three circular marks
approximately opposite the circumferential rub mark.
Photo 10: View of right engine’s 2nd stage LPT blades showing no damage. The two cuts in the
fan blades were from the saw that was used to facilitate the examination of the engine. (PWC)
The EFSO valve housing was broken at the aft end. The EFSO valve plunger was
extended about twice as much as if it had normally been tripped. (Photo 11) The EFSO valve
actuation button had a burnish mark that corresponded to the circumferential rub mark on the
end of the LPT shaft.
Photo 11: Right engine’s emergency fuel shutoff valve showing actuator extended
4.0 Ornithological examination
The examination of a piece of the right engine’s inlet duct by the Smithsonian Institute’s
Feather Identification Laboratory revealed a feather fragment embedded in a rivet hole. The
laboratory’s examination of the feather revealed that it was consistent with feathers from a
Previous DNA testing by the Feather Identification Laboratory of material that was found
splattered on the right side of the vertical stabilizer determined the material was from an
American white pelican (Pelecanus erythrorhynchos). According to the Cornell University Lab
of Ornithology website, an American white pelican is a large white waterbird with a long bill
that has an extensible pouch. An American white pelican can be 50 to 65-inches tall, have a
wingspan of 96 to 114-inches, and weigh up to almost 20 pounds. (Photo 12)
Photo 12: American white pelican. (Cornell University Ornithology Laboratory)
5.0 Engine certification large bird ingestion test
As part of the development and certification of the JT15D-1A engine, PWC 8 was
required to conduct bird ingestion tests including large bird ingestion tests. PWC’s report to the
Canadian Department of Transport 9 Air Transportation Administration stated in Section 3
Conclusions & Recommendations, that the engine had satisfactorily demonstrated the ability to
meet the foreign object ingestion requirements in conformity with the Federal Aviation
Administration’s Advisory Circular (AC) 33-1A Turbine Engine Foreign Object Ingestion and
Rotor Blade Containment Type Certification Procedures. In addition to the large bird ingestion
test, AC 33-1A required the engine ingest a number of items that included small birds, sand and
gravel, cleaning cloth, water, hailstones, and an ice sheet. With regards to the large bird
ingestion test, AC 33-1A required a freedom from catastrophic effects without any reference to
PWC’s report stated that for the large bird ingestion test, a freshly killed 4-pound chicken
was propelled into the front of the engine at a speed of 495 feet per second (293 knots) while the
engine was at maximum continuous cruise power conditions. PWC’s report stated that the
engine stopped abruptly immediately after the chicken was propelled into the engine. PWC’s
report further stated that following the ingestion, five fan blades broke and penetrated the
aluminum fan case. PWC’s report stated that the aluminum fan case was inadequate and was
replaced with a steel fan case.
At the time the engine was certified in 1971, PWC was referred to as United Aircraft of Canada.
The Canadian Department of Transport is now referred to as Transport Canada.