Honeywell EngineTeardown Report-T53-L-13B, SN LE-24073

Transcription

TEARDOWN REPORT
OF ONE MODEL T53-L-13B
TURBOSHAFT ENGINE
SERIAL NUMBER LE-24073
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May 12,2000
Approved By:
Product Safety and Integrity
Honeywell
TEARDOWN REPORT
OF ONE MODEL T53-L-13B
TURBOSHAFT ENGINE
SERIAL NUMBER LE-24073
1.0
INTRODUCTION AND SUMMARY
1.1
Pumose
This report presents the findings of a engine examination conducted on a Honeywell
International (AlliedSignal/Lycoming) model T53-L-13B turboshaft engine, serial number LE24073, at the Honeywell International Product Safety & Integrity facility in Phoenix, Arizona on
February 24 and 25,2000.
The examination was conducted at the request of the National Transportation Safety
Board (NTSB), and under the cognizance of the Federal Aviation Administration (FAA).
1.2
Backmound
The engine, serial number LE-24073, was installed in a Bell UH-1H helicopter,
registration N853M, serial number 68-16087. The aircraft crashed near Ft. Myers, Florida on
January 20, 2000. The NTSB reported that the pilot ‘heard a loud “pop”, the engine fail light
illuminated, and engine rpm decayed to zero.’
Summary
The engine damage observed was the result of separation of a section of the second-stage
GP turbine disk.
The separation of the section of the second-stage GP turbine disk was due to peak strain
low-cycle fatigue. No material defects were identified which would contribute to the
separation of the section of the second-stage GP turbine disk.
The service history of the second-stage GP turbine disk could not be determined fiom
records provided by the operator.
All other engine damage observed is attributable to the unbalance and structural damage
induced by the separation of a section of the second-stage GP turbine disk.
No conditions were observed which would have caused or contributed to the separation
of a section of the second-stage GP turbine disk.
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2.0
FINDINGS OF T53-L-13B9SERIAL NUMBER LE-24073
NOTES
1. All references to position are aft looking forward. All references to
angular measurements are clockwise.
2. All observations reported herein are based on visual examinations
with the unaided eye, unless otherwise noted.
2.1
General
(a)
The engine was received in a military-style engine shipping container. The engine
nameplate (Figure 1) indicated that the engine had been supplied to the U.S. Army under
contract DAAJO1-73-D-0036-004 in October, 1974. A Corpus Christi Army Depot
(CCAD) overhaul tag was attached to the exterior of the engine inlet (Figure 1). The
blocks on the CCAD overhaul tag for ‘Overhaul Date’ and ‘Time Since New’ were
marked “/A’.
The gas generating (NG) spool rotated freely. Continuity of the NG spool and NG
accessory drive geartrain was verified by rotating the starter drive pad and observing
rotation of the compressor, and accessory gearbox tachometer generator output drive
shaft.
The power turbine (NP) spool assembly was initially not free to rotate, and metallic
fragments were observed visually in the power turbine. After removal of the
combustor/turbine module, the power turbine shaft rotated freely. Continuity of the NP
spool and NP accessory drive geartrain was verified by rotating the power turbine shaft
and observing rotation of the output shaft, and overspeed governor tachometer generator
gearbox output drive shaft.
There was no evidence of fire damage (Figure 3). The engine mounts were intact (Figure
3).
2.2
Output Reduction Carrier and Gear Assembly
(a)
The output reduction gearbox assembly, part number 1-030-350-12, serial number P908,
was undamaged (Figure 4). All planet gears, bearings and the output shaft rotated freely.
There was residual oil on all internal surfaces and components of the reduction gearbox.
The reduction gearbox input sun gear was undamaged. The sun gear retaining bolt
washer was fractured (Figure 5). There was residual oil on all surfaces of the sun gear
retaining bolt washer.
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2.3
Accessory Drive Carrier Assembly
The accessory drive carrier assembly assembly, part number 1-030-350-12, serial number
P908, was undamaged (Figure 6). All gears and bearings rotated freely. There was residual oil
on all internal surfaces and components of the reduction gearbox.
2.4
Inlet Housing Assembly
The inlet housing assembly was undamaged (Figure 7). There was residual oil on
internal surfaces of the inlet housing.
2.5
Overspeed Governor and Tachometer Drive Assembly
The overspeed governor and tachometer drive assembly (Figure 8), part number 1-160500-04, serial number not observed, was undamaged and rotated freely. The overspeed governor
and tachometer drive assembly driveshaft was undamaged.
2.6
Accessory Drive Gearbox Assembly
The accessory drive gearbox assembly (Figure 9), part number 1-070-220-03, serial
number 248A, was undamaged and rotated freely. The accessory drive gearbox assembly
driveshaft was undamaged.
Compressor Section
The variable inlet guide vanes (VIGV) were observed to be in the fully closed position
upon receipt of the engine. The VIGV’s rotated freely and in unison when actuated
manually using the VIGV connector rod.
The VIGV unison ring was undamaged.
The VIGV connector rod was undamaged and travelled freely.
The NG accessory drive pinion gear was undamaged. There was residual oil on the NG
accessory drive pinion gear teeth.
The first-stage axial compressor rotor was intact (Figure 10). There was dirt adhering to
the pressure and suction sides of the first-stage compressor rotor blades (Figure 10).
Approximately half of the first-stage compressor rotor blade tips displayed rotational
scoring (Figure 10). The inter-stage seal land aft of the first-stage axial compressor rotor
displayed rotational scoring, through approximately 180 degrees, with corresponding
rotational score marks on the first-stage axial compressor stator forward ID vane
platform.
The second-stage axial compressor rotor was intact (Figure 10). There was dirt adhering
to the pressure and suction sides of the second-stage compressor rotor blades (Figures 10
and 11). Approximately half of the second-stage compressor rotor blade tips displayed
rotational scoring and material accumulation on both pressure and suction sides (Figures
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10 and 11). The inter-stage seal land forward of the second-stage axial compressor rotor
displayed rotational scoring, through approximately 180 degrees, adjacent to the aft ID
vane platform on the first-stage axial compressor rotor stator (Figure l l ) , with
corresponding rotational score marks on the aft ID vane platform on the first-stage axial
compressor rotor stator. The inter-stage seal land aft of the second-stage axial
compressor rotor displayed rotational scoring, through approximately 180 degrees, with
corresponding rotational score marks on the forward ID vane platform on the secondstage axial compressor rotor stator.
The third-stage axial compressor rotor was intact (Figure 10). There was dirt adhering to
the pressure and suction sides of the third-stage compressor rotor blades (Figures 10 and
11). Approximately half of the third-stage compressor rotor blade tips displayed
10 and 11). The inter-stage seal land forward of the third-stage axial compressor rotor
rotational score marks on the aft ID vane platform on the second-stage axial compressor
rotor stator. The inter-stage seal land aft of the third-stage axial compressor rotor
rotational score marks on the forward ID vane platform on the third-stage axial
compressor rotor stator.
The fourth-stage axial compressor rotor was intact (Figure 10). There was dirt adhering
to the pressure and suction sides of the fourth-stage compressor rotor blades (Figures 10
and 1 I). Approximately half of the fourth-stage compressor rotor blade tips displayed
10 and 11). The inter-stage seal land forward of the fourth-stage axial compressor rotor
rotational score marks on the aft ID vane platform on the third-stage axial compressor
rotor stator. The inter-stage seal land aft of the fourth-stage axial compressor rotor
rotational score marks on the forward ID vane platform on the fourth-stage axial
compressor rotor stator.
The fifth-stage axial compressor rotor was intact (Figure 10). There was dirt adhering to
the pressure and suction sides of the fifth-stage compressor rotor blades (Figures 10 and
11). Approximately half of the fifth-stage compressor rotor blade tips displayed
10 and 11). The inter-stage seal land forward of the fifth-stage axial compressor rotor
rotational score marks on the aft ID vane platform on the fourth-stage axial compressor
rotor stator. The inter-stage seal land aft of the fifth-stage axial compressor rotor
rotational score marks on the forward ID vane platform on the compressor exit guide
vane.
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The first-stage axial compressor rotor stator was intact (Figure 12). There was dirt
adhering to the pressure and suction sides of the first-stage axial compressor rotor stator
vanes (Figure 12). The forward and aft ID vane platforms of the first-stage axial
compressor rotor stator displayed rotational scoring, with corresponding rotational score
marks on the inter-stage seal land aft of the first-stage axial compressor rotor and the
inter-stage seal land forward of the second-stage axial compressor rotor.
The second-stage axial compressor rotor stator was intact (Figure 12). There was dirt
adhering to the pressure and suction sides of the second-stage axial compressor rotor
stator vanes (Figure 12). The forward and aft ID vane platforms of the second-stage axial
marks on the inter-stage seal land aft of the second-stage axial compressor rotor and the
inter-stage seal land forward of the third-stage axial compressor rotor.
The third-stage axial compressor rotor stator was intact (Figure 12). There was dirt
adhering to the pressure and suction sides of the third-stage axial compressor rotor stator
vanes (Figure 12). The forward and aft ID vane platforms of the third-stage axial
marks on the inter-stage seal land aft of the third-stage axial compressor rotor and the
inter-stage seal land forward of the fourth-stage axial compressor rotor.
The fourth-stage axial compressor rotor stator was intact (Figure 12). There was dirt
adhering to the pressure and suction sides of the fourth-stage axial compressor rotor stator
vanes (Figure 12). The forward and aft ID vane platforms of the fourth-stage axial
marks on the inter-stage seal land aft of the fourth-stage axial compressor rotor and the
inter-stage seal land forward of the fifth-stage axial compressor rotor.
The compressor exit guide vanes were intact (Figure 12). There was dirt adhering to the
pressure and suction sides of the compressor exit guide vanes (Figure 12). The forward
and aft ID vane platforms of the compressor exit guide vanes displayed rotational scoring,
with corresponding rotational score marks on the inter-stage seal land aft of the fifthstage axial compressor rotor.
The upper axial compressor housing was intact (Figures 13 and 14). The lower axial
compressor housing was intact, and was not removed for examination.
The centrifugal compressor was intact (Figure 1 9 , and was not removed for examination.
The inlet and transition shroud-line edge of approximately half of the blades of the
centrifugal compressor displayed rotational scoring and material accumulation on both
pressure and suction sides (Figure 15), with corresponding rotational score marks on the
adjacent surfaces of the centrifugal compressor shroud.
The upper centrifugal compressor shroud was intact (Figure 16). The upper centrifugal
compressor shroud displayed rotational scoring through approximately 180 degrees on
surfaces adjacent to inlet and transition shroud-line edges of the centrifugal compressor
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blades (Figure 16), with corresponding rotational score marks on the inlet and transition
shroud-line edges of the centrifugal compressor blades. The lower centrifugal
compressor shroud was intact, and was not removed for examination.
Combustor Section
The 1 and 11 o’clock ignitor plugs were intact (Figure 17). There was fretting on the
surfaces of both ignitor plugs at the combustion chamber interface (Figure 17). There
was carbon on the tips and electrodes of both ignitor plugs (Figure 17). The electrodes of
both ignitor plugs were eroded (Figure 17).
The 4 and 8 o’clock ignitor plugs were intact (Figure 18). There was fretting on the
surfaces of both ignitor plugs at the combustion chamber interface (Figure 18). There
was carbon on the tips and electrodes of both ignitor plugs (Figure 18). The electrodes of
both ignitor plugs were eroded (Figure 18).
The combustor drain valve appeared to be undamaged, but was not removed for
inspection.
The combustion chamber liner was intact (Figure 19). There were carbon deposits on
external (Figure 19) and internal surfaces of the combustion chamber liner dome. There
was an outward puncture through the combustion chamber liner inner wall (Figure 20)
and outer wall (Figure 21) at approximately 8 o’clock.
The combustion chamber plenum was intact (Figure 22). There was an outward bulge in
the combustion chamber plenum wall at approximately 8 o’clock (Figure 23).
The combustion chamber outer curl appeared to be intact (Figure 24), but was not
removed for examination. There was metallic debris in the dished area of the combustion
chamber outer curl (Figure 24).
The diffuser housing appeared to be undamaged, but was not removed for examination.
The v-band clamp securing the exhaust diffuser support cone assembly was undamaged.
The exhaust diffuser support cone assembly was intact (Figure 25). There was corrosion
on external surfaces of the exhaust diffuser support cone assembly.
The exhaust diffuser assembly was intact (Figure 26). The exhaust diffuser struts were
intact. The exhaust diffuser strut at 12 o’clock was damaged (Figure 27).
The fire shield was undamaged (Figure 28).
The start fuel nozzles were intact (Figure 29) and appeared to undamaged, but were not
functionally tested. There were carbon deposits on the start fuel nozzle tips (Figure 29).
Portions of the gaskets under the start fuel nozzles had separated and were missing
(Figure 30).
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The right and left fuel manifold assemblies were intact (Figures 31 and 32) and appeared
to undamaged, but were not functionally tested. There were carbon deposits on the right
and left fuel manifold assembly fuel injector air shrouds (Figure 33).
Gas Producer (GP) Turbine Section
The first-stage GP turbine nozzle was intact (Figure 34), but was not removed from the
engine due to the severe corrosion of the retaining bolts (Figure 35). The trailing edges
of several first-stage GP turbine nozzle vanes were fractured and deflected forward
(Figure 36). There was a rub, with what appeared to be oxidation covering the rub, over
360 degrees, on the first-stage turbine rotor shroud (Figure 37). Several bolts securing
the second-stage GP turbine cylinder to the first-stage GP turbine nozzle were found to be
fractured (Figures 35,36 and 37).
The first-stage GP turbine rotor was intact (Figure 38). There was rotational scoring,
with what appeared to be oxidation covering the score marks on the tips of all first-stage
GP turbine rotor (Figures 39 and 40). There were nicks and dents on the trailing edges of
the first-stage GP turbine rotor blades (Figure 40).
The second-stage GP turbine nozzle was heavily damaged (Figures 41 and 42). The
leading edges of the second-stage GP turbine nozzle vanes were cracked (Figure 43). The
trailing edges of the second-stage GP turbine nozzle vanes were deformed and fiactured
(Figure 44). Sections of the trailing edges of the second-stage GP turbine nozzle vanes
were missing (Figures 42 and 44). There was rotational scoring, over 360 degrees, on the
aft face of the second-stage GP turbine nozzle inner support (Figures 42 and 44).
Sections of the aft face of the second-stage GP turbine nozzle inner support were missing
(Figures 42 and 44). The second-stage GP turbine rotor shroud was scored and deformed
(Figures 42 and 44). Sections of the aft face of the second-stage GP turbine nozzle outer
support were missing (Figures 42 and 44).
The second-stage GP turbine rotor assembly (Figure 45) was identified as part number 1101-360-04, serial number 5 15. The second-stage GP turbine rotor disk was identified as
part number 1-100-063-05, serial number 515, manufacturer code LE-72. A portion of
the second-stage GP turbine rotor disk rim had separated (Figures 45 and 46). Six
second-stage GP turbine rotor blades were missing (Figures 45 and 46). The remaining
second-stage GP turbine rotor blades were fi-actured at approximately mid-span (Figures
45,46 and 47).
The second-stage GP turbine rotor assembly was submitted for Material Analysis (MA)
to determine the reason for the separation of a portion of the second-stage GP turbine
rotor disk rim. The results of the MA indicate that the fracture of the second-stage GP
turbine rotor assembly was due to peak strain low-cycle fatigue (LCF). No material
defects were identified which would contribute to the separation of the section of the
second-stage GP turbine disk. The MA is included in Appendix I.
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The GP spool aft nut was intact (Figure 48) and the lock cup properly secured. The
break-away torque of the aft nut was approximately 50 ft-lb.
The forward and aft GP turbine cones were intact (Figure 49). The aft GP cone displayed
uneven serration patterns corresponding to the aft spline on the GP shaft. The aft GP
turbine cone was submitted for MA to determine the reason for the uneven serration
pattern. The MA is included in Appendix I.
Power Turbine
The first-stage PT nozzle was heavily damaged (Figures 50 and 5 1). All of the first-stage
PT nozzle vanes had separated and were missing (Figures 50 and 51). There was
rotational scoring, over 360 degrees, on the inner diameter of the first-stage PT nozzle
outer support (Figures 50 and 52). There was an outward puncture of the outer support
approximately three inches in length (Figure 52). The first-stage PT nozzle inner support
was cracked in four places (Figure 51). There was rotational scoring, over 360 degrees,
on the forward face of the first-stage PT nozzle inner support (Figure 5 1).
The first-stage PT rotor (Figure 53) was identified as part number 1-190-010-02, serial
number 24273, CAGE code 91547. The leading edges and tips of all first-stage PT rotor
blades were battered, and sections of all blades were missing (Figure 53). Five of the
first-stage PT rotor blades were fractured near the disk OD and had separated (Figures 53
and 54).
The second-stage PT nozzle was heavily damaged (Figure 55). One of the second-stage
PT nozzle vanes was missing (Figure 55). The leading edges of all second-stage PT
nozzle vanes were battered and dented (Figures 55 and 56). The trailing edges of all
second-stage PT nozzle vanes were battered and dented (Figures 56 and 57). Sections of
the trailing edges of all second-stage PT nozzle vanes had separated and were missing
(Figures 55 and 57).
The second-stage PT rotor (Figure 58) was identified as part number 1- 140-550-07, serial
number 000078. The leading edges of all second-stage PT rotor blades were battered,
and sections of five blades were missing (Figures 58 and 59). There were cracks on the
trailing edges of several second-stage PT rotor blades (Figure 59).
The No. 3 (roller) and No. 4 (ball) bearing housing was intact (Figure 60) and appeared to
be undamaged, but was not disassembled for inspection. The bearing housing rotated
freely, and there was residual oil on internal surfaces of the bearing housing. There were
soot and carbon deposits on external surfaces of the bearing housing (Figure 60). The
heat shield enclosing the bearing housing was corroded, and the aft section of the heat
shield had separated and was missing (Figure 60).
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2.11
Fuel
(a)
The fuel control and overspeed governor unit (Figure 61), part number 1-170-240-60,
serial number 12AS11221, appeared to be undamaged. The servo (Figure 62) and inlet
(Figure 63) fuel filters were intact and uncontaminated. A fuel sample was obtained from
the fuel control and submitted for analysis. No anomalies were observed. The results of
the analysis are included in Appendix 11.
The fuel control and overspeed governor unit were functionally tested at the Pueblo
Airmotive facility in Tucson, Arizona on April 19, 2000. The test data sheet is included
in Appendix 111. The following conditions were observed:
1.
The fuel flow recorded at the sixth point of the 59" accel schedule test, at 3700
rpm NG and 11 in. Hg inlet pressure (Pl) was 295 pph, 5 pph above the
maximum limit of 290 pph specified.
2.
The fuel flow recorded at the eighth point of the 59" accel schedule test, at 4200
rpm NG and 29.92 in. Hg P1 was 950 pph, 5 pph above the maximum limit of 945
pph specified.
3.
The fuel flow recorded at the second point of the deceleration schedule and
minimum flow test, at 2800 rpm NG and 29.92 in. Hg P1, was 146 pph, 5 pph
above the maximum limit of 141 pph specified.
4.
The NG speed recorded at first point of the air bleed trigger line test, at 800 pph
with the actuator closed, was 3898 rpm, 67 rpm below the minimum limit of 3965
rpm specified. This condition is the result of a field adjustment allowed to match
bleed band operation to engine characteristics.
5.
The NG speed recorded at second point of the air bleed trigger line test, at 400
pph with the actuator closed, was 3303 rpm, 64 rpm below the minimum limit of
3367 rpm specified. This condition is the result of a field adjustment allowed to
match bleed band operation to engine characteristics.
6.
The NG speed recorded at fourth point of the air bleed trigger line test, at 800 pph
with the actuator open, was 3836 rpm, 66 rpm below the minimum limit of 3902
7.
The NG speed recorded at fifth point of the air bleed trigger line test, at 400 pph
with the actuator open, was 3262 rpm, 53 rpm below the minimum limit of 33 15
8.
The NG speed recorded at the first point of the steady state and face cam test, at
23" throttle angle, was 2145 rpm at 137 pph, 35 rpm below the minimum limit of
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21 80 rpm specified. This condition is the result of a field adjustment allowed to
adjust fuel flow to engine characteristics.
9.
The NG speed recorded at the third point of the steady state and face cam test, at
38" throttle angle, was 2843 rpm at 210 pph, 47 rpm below the minimum limit of
2890 rpm specified. This condition is the result of a field adjustment allowed to
adjust fuel flow to engine characteristics.
10.
The NG speed recorded at the fifth point of the steady state and face cam test, at
100" throttle angle, was 4320 rpm at 800 pph, 21 rpm above the maximum limit
of 4299 rpm specified. This condition is the result of a field adjustment allowed
to adjust fuel flow to engine characteristics.
The flow divider and dump valve (Figure 64) appeared to be undamaged, but was not
functionally tested. There was corrosion on external surfaces of the flow divider and
dump valve. No component identification markings were observed.
The start solenoid valve (Figure 65), part number 1-300-191-05, serial number G07485,
appeared to be undamaged, but was not functionally tested. The attaching ears of the
start solenoid valve were fractured (Figure 65).
The VIGV actuator (Figure 66), part number 1-1SO- 150-01, serial number 68, appeared to
be undamaged, but was not functionally tested.
Oil
The oil pump (Figure 67), part number 1-300-212-04, serial number C8168, appeared to
be undamaged, but was not functionally tested. There was residual oil within the oil
pump. The oil pump rotated freely when the drive shaft was manually rotated, and
pumping and scavenge operation was observed when the oil pump was rotated.
The oil filter was intact and uncontaminated (Figure 68). The oil filter was submitted for
backflush and analysis. The filter backflush showed that 6 milligrams of contaminants
were present in the filter, consisting mostly of dirt and fibers, with trace amounts of silver
and iron. The results of the analysis are included in Appendix 11.
The magnetic chip detector was intact and uncontaminated (Figure 69).
The accessory gearbox oil strainer (Figure 70) and rear turbine oil screen (Figure 71)
were intact and uncontaminated.
During disassembly the rear turbine oil scavenge fitting (Figure 72) was observed to be
hand tight.
An oil sample was obtained from the air inlet housing and submitted for analysis. 13
ppm of iron was detected during SOAP analysis. The results of the analysis are included
in Appendix 11.
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Pneumatic
The hot air valve, serial number 6031434 appeared to be undamaged, but was not
functionally tested (Figure 73). No component part number was observed.
The bleed band actuator, part number 1-170-050-12, serial number X347 appeared to be
undamaged, but was not functionally tested (Figure 74).
Electrical
The ignition exciter, Unison part number 10-383225-1, serial number not observed,
appeared to be undamaged, but was not functionally tested (Figure 75).
The ignition leads appeared to be undamaged, but were not functionally tested (Figure
76).
The exhaust gas temperature thermocouple harness was marked as part number ‘8 1996/1300-177-4V’, ‘MFR 24733’, NSN 1680-00-732-0636,, serial number 107023, appeared
to be undamaged (Figure 77). The exhaust gas temperature thermocouple harness was
functionally tested. Functional test results demonstrate that the exhaust gas temperature
thermocouple harness was capable of providing an accurate signal representing the
exhaust gas temperature. The results of the functional test are included in Appendix 111.
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3.0
Analvsis and Conclusions
3.1
Analvsis
The turbine damage observed was initiated by separation of a section of the second-stage
GP turbine disk due to peak strain low-cycle fatigue.
Separation of the section of the second-stage GP turbine disk resulted in liberation of disk
and blade fragments which caused lesser secondary damage to the upstream second-stage GP
turbine nozzle, first-stage GP turbine rotor, and first-stage GP turbine nozzle, and more extensive
secondary damage to the downstream first-stage PT nozzle, first-stage PT rotor, and second-stage
PT nozzle and rotor.
Separation of the section of the second-stage GP turbine disk resulted in significant
imbalance of the GP spool, which caused the interference of the axial and centrifugal compressor
stages with the respective shrouds. The interference of the axial and centrifugal compressor
stages with the respective shrouds resulted in the tip rubs observed on the first- through fifthstage axial compressor blades, and the centrifugal compressor blades.
Imbalance of the GP spool, or the initial imbalance load caused by separation of the
section of the second-stage GP turbine disk, resulted in relative movement between the aft GP
turbine cone and the GP spool. This relative movement resulted in interference of the aft GP
turbine cone with the aft internal splines of the GP spool, and the impressions observed on the aft
GP turbine cone, and the fracture of the sun gear retaining bolt washer.
No material defects were identified which would contribute to the separation of the
section of the second-stage GP turbine disk. The service history of the second-stage GP turbine
disk could not be determined from records provided by the operator.
Qualitative bearing evaluation and freedom of the oil filter and magnetic chip detector
from metallic particles suggests that a lubrication related bearing degradation and resulting
vibration did not contribute to the separation of the section of the second-stage GP turbine disk.
Functional testing of the fuel control demonstrated that the fuel control and overspeed
governor were operable. The fuel control had been adjusted to accommodate engine
characteristics, resulting in conditions observed which did not meet functional test specifications.
No fuel control conditions were identified which would have prevented satisfactory engine
operation, or contributed to separation of a section of the second-stage GP turbine disk.
3.2
Conclusions
(a)
The engine damage observed was the result of separation of a section of the second-stage
GP turbine disk.
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(b)
The separation of the section of the second-stage GP turbine disk was due to peak strain
low-cycle fatigue. No material defects were identified which would contribute to the
separation of the section of the second-stage GP turbine disk.
(c)
The service history of the second-stage GP turbine disk could not be determined from
records provided by the operator.
(d)
All other engine damage observed is attributable to the unbalance and structural damage
induced by the separation of a section of the second-stage GP turbine disk.
(e)
No conditions were observed which would have caused or contributed to the separation
of a section of the second-stage GP turbine disk.
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PI
ooz I so:22 I 0: ??a:IVsd
Hone-well
Figure 35.
First-Stage GP Turbine Nozzle Bolts (Photo 6-8).
Figure 36.
First-Stage GP Turbine Nozzle Bolts (Photo 6-13).
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Figure 43.
Second-Stage GP Turbine Nozzle Vanes (Photo 5-9).
Figure 44.
Second-Stage GP Turbine Nozzle Vanes (Photo 5-12).
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Figure 45.
Figure 46.
Second-Stage GP Turbine Rotor (Photo 4-21).
Second-Stage GP Turbine Rotor Disk Rim
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(Photo 4-22).
Honeywell
Figure 51.
First-Stage PT Nozzle Inner Support
(Photo 4-17).
Figure 52.
First-Stage PT Nozzle Outer Support
(Photo 7-9).
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Figure 53.
Figure 54.
First-Stage PT Rotor
(Photo 5-31).
First-Stage PT Rotor Blades
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(Photo 5-32).
Honeywell
Figure 55.
Figure 56.
Second-Stage PT Nozzle
(Photo 7-1).
Second-Stage PT Nozzle Vanes (Photo 7-4).
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Figure 57.
Figure 58.
Second-Stage PT Nozzle Vanes (Photo 7-3).
Second-Stage PT Rotor (Photo 7-17).
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Figure 59.
Figure 60.
Second-Stage PT Rotor Blades (Photo 7-18).
No. 3 and No. 4 Bearing Package
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(Photo7-20).
Honeywell
I
COMPONENT INVESTIGATIONS
- For Internal Use Only -
D. Looper
PHONE 365-2365
2nd GP Turb. Rotor AssyICone Rear compressor
UESTER
T NAME
ERIAL
D979/aluminurn bronze/C101
IP~ENGINE T53-L-13B
PRELIMINARY MATERIALS ANALYSIS 102406
Page 1 of 23
Engines & Systems
Phoenix
4009-21224
1- 10 1-360-04/1-110-14 1-01 PART SIN
CHARGE NO.
PART NO.
LE-24073
03/02/00
5 15
M3603lAMS4640lM3617 SUPPLIER Not Reported
MAT'L SPEC
EQUIP SIN
DATE
OPERATING TIMEICYCLES
Reportedly 962 TSO
IITIONAL INFORMATION
irbiiie Rotor assembly and aft cone removed from engine following aircraft accident on 1/20/00 near Ft. Myers, Florida.
igine records indicate 962 hours since overhaul, cycles unknown, total time unknown.
verhaul nameplate indicates CCAD overhaul, however CCAD has no record of overhaul work on this engine
.igine was run at CCAD test cell on 4/8/93. Engine build records being pulled.
iarkings on Disk: MFR LE72, S/N 5 15,91547, 1- 100-360-04 (corrected to 1- 101-360-04), RCVOR S/N 5 1 5
I54711- 100-063-05L.
isk P/N 1-100-063-05 & Blade P/N 1-100-1 18-07.
3BLEM STATEMENT
iocument rotor and cone. Determine nature of fracture of rotor.
xamine aft cone and determine if evidence of vibration is present.
MMARY/CONCLUSION(S)
'he primary separation of the 2"dGP Turbine Rotor was the result of Sustained Peak Strain Low Cycle Fatigue.
Three of the disk post separation fractures and the exposed fractures of two pin slot cracks exhibited similar heavily
oxidized intergranular features. Evidence of fracture progression was observed as variations in the surface oxide. These
features are typically associated with Sustained Peak Strain Low Cycle Fatigue.
The fourth disk post separation was the result of secondary overload.
'he Rear Compressor Cone exhibited thick oxides on the surface and evidence of deformation.
Evidence of fret wear was not detected.
~ESTIGATOR F.
Krempski
IISTRIBUTION
>.Looper
. Hanenburg
VI.
Morgan
2. Henry
vi. Woodhall
'. GourleylR. Sines
'
.
(2)
REVIEWER
/f*&-.
DEPT
MIS
72-20
93-23
93-23
93-64
93-32
2 102- 12 1
404-282
404-282
404-257
404-287
93-33
301-227
I
DATE
NAME OF PART
2"d GP Turb. Disk
Cone, aft
2"dGP Turb. Blade
sb!/&REVIEWER
PART NO.
1- 100-063-05
1-1 10-141-01
1-100-1 18-07
Gour'ey
DATE 'OF
MATERIAL
PART CONDITION
D979
Aluminum bronze
ClOl
FA SPS Low Cyc
OT deform. crack
DO no OT
PRELIMINARY RESULTS
REPORTED TO:
Distribution
REFERENCES
No similar previous MAS.
J.
DATE:
03/17/00
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Engines & Systems Phoenix
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Overview
Three “sets” of hardware were submitted for analysis.
0
The 2”d GP Turbine Rotor with remnants of all but six of the blades still attached.
0
The Rear Compressor Cone.
0
A bag of multiple blade and vane pieces.
The results will be reported for each set of hardware in the order listed above.
2nd GP Turbine Rotor
Macroexamination
The rotor exhibited uniform separations of all of the attached blades at a height of
approximately 0.2-inch to approximately 0.4-inch outboard of the disk rim (Figure I A).
The fractures were indicative of impact overload and were the result of secondary
damage (not shown).
The firtree slots of two of the six missing blades exhibited evidence of deformation on
the firtree dovetails at the forward side of the slot (not shown). One slot exhibited
approximately 0.15” of deformation and the other slot exhibited approximately 0.35” of
dovetail deformation. The locking pins in both slots were fractured in shear. Several
other attached blades were displaced forward on the disk rim. These blade separations
were also believed to be secondary damage.
The rotor exhibited separations of the forward side of four adjacent posts (tenons) as
shown in Figures 1A & B. Two fractures initiated at the sharp corners of the same pin
slot (thick white arrows, Figures 2A, 3A & 3B). Both fractures were heavily oxidized
and coated with dirt and debris. Close examination revealed intergranular fracture
features with some evidence of fracture progression indicated by differences in
oxidation (black arrows, Figures 3A & B).
The intergranular appearing portions of these two adjacent fractures were similar in size
measuring approximately:
0.2-inch along both of the pin slots,
0
0.3-inch at the center of the fracture and
0
0.26-inch along the forward face.
The remaining portions of the post separation fractures were the result of overload
(brackets, Figures 3A & B). This portion of the fracture was also discolored due to
oxidation.
The one partial fracture, although smaller, exhibited similar intergranular features and
initiated at the corner of the pin slot (arrow, Figure 3C). The third fracture was the
result of overload (Figure 3D).
Additional similar appearing cracks were observed emanating from the corners of eight
other pin slots (reference black arrows, Figure 1A). Some pin slots exhibited only one
crack (Figures 4C & D) and others exhibited two (Figures 4A, 4B, 5A & 5C).
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Two cracks emanating from the corners of pin slot "D" (Figures 5A & C) that was
situated approximately 180" opposite the separation fractures (reference Figure 1A)
were exposed for examination. These fractures were also heavily oxidized and
intergranular in appearance (Figures 5B & D). These two fractures were measured to be
approximately:
1. 0.12-inch in the pin hole & 0.14-inch on the forward face.
2. 0.15-inch along the pin hole (including the tunneled portion) & 0.1 1-inch on the
forward face.
Macrohardness test results
The hardness of the disk was measured on the forward hub rim and on the forward post
endface to be 71HRA and 71HRA, respectively. Specified was 69HRA to 73HRA.
Scanning electron microscope examination
One of the separation fractures (Figures 6 & 7) and both of the exposed fractures were
examined using the scanning electron microscope (SEM). The fractures all exhibited
intergranular features. Intergranular features and heavy oxides were evident near the
origins (Figures 6B, 6C, 8B & 9B). The extent of oxidation decreased with increased
distance from the origins (Figures 6A, 7A, 7B, 7C, 8C, 8D, 8E, 9C & 9D).
Metallographic examination
A metallographic section was prepared to view one separation fracture by lightly
grinding and polishing parallel to the pin slot into the side of the fracture with the
fatigue origin (Figures 10 & 11A).
The fracture was primarily intergranular with some scattered clusters of transgranular
fractures. The fracture features were very heavily oxidized near the fatigue origin
(white arrows, Figure 1OA). The oxidation products on the fracture surface decreased
with distance from the origin (Figures 1OC & D) and some of the intergranular cracks
exhibited heavy oxidation products (arrow, Figure 1OC). The grain boundaries in the
cross section exhibited heavy precipitation (Figure 11).
The grain size on the aft side of the platform was distinctly smaller than in the
remaining portion of the cross section. The grain size within the aft disk rim (Figure
11C) was measured to be ASTM 8 and the grain size within the for'ward disk and rim
was measured to be ASTM 7 (Figure 1 IB). This variation in grain size was believed to
be from the forging.
The Rear Compressor Cone
Macroexam ination
The cone was relatively dark in appearance with a dull gray color. The thin portion of
the cone exhibited distinct contours on the outer and inner surfaces indicative of close
contact with the mating parts. The outer cone-shaped surface exhibited contours
suggestive of the ends of a spline (Figure 12A & B). These contours were not uniform
around the entire circumference. One approximately 180' segment (Figures 12 and
13A) exhibited much more pronounced contours than the opposite 180" segment
(Figures 13B & E). The more pronounced contours exhibited features suggesting that
the contours were the result of deformation of the part (Figures 12 and 13A).
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Narrow bands on the inner and outer surfaces of the cone exhibited score marks
indicative of axial movement between the cone and the mating parts. The white arrow
in Figure 12B denotes the location on the outer diameter. The marks on the inner
diameter are in approximately the same location (not shown). These score marks were
relatively bright in appearance compared to the remaining portions of the cone. This
bright appearance indicated that the marks were made late during service compared to
the other damage on the cone.
Scanning electron microscope (SEM) examination
SEM examination of several of the contours on the outer and inner surfaces of the cone
revealed thick accumulations of oxidation/corrosion products (Figures 14, 15A, 15B
and 16). Distinct evidence of fret wear damage was not observed.
The evidence of axial movement was confirmed at the one location on the outer surface
of the cone (Figures 15C & D). The location on the inner diameter was not examined.
Metallographic sections were prepared to view the contours in the thin portion of the
cone (Figures 17A & B). The thick oxides on the surface were confirmed (white
arrows, Figures 17C & D). In addition, a tight crack (black arrow, Figure 17E) and
some deformed microstructural features (bracket, Figure 17E) were observed at the
corner on the inner diameter and at other locations along the surface (not shown).
Multiple blade a n d vane pieces
Macroexamination
The multiple assorted pieces submitted in the bag were cleaned, sorted (Figure 18) and
examined at up to 40X magnification. All of the damage was secondary impact
damage.
Metallographic cross sections were made through the tip of one of the longest blade
pieces (reference Figure 18C) and one of the blade tip shrouds. No evidence of over
temperature exposure was observed (Figure 19).
Energy dispersive x-ray examination
The two metallographic cross sections (reference Figures 19A & C) were examined
using energy dispersive x-ray (EDX) analysis to identify the alloy of each part. The
EDX spectrum obtained from the blade corresponded to C 10 1 alloy, the alloy specified
for the 2"d GP Turbine Blade. The EDX spectrum obtained from the tip shroud.
corresponded to IN7 13C alloy, the alloy specified for other blade stages.
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Figure 1. (A) An overall view of the forward side of the 2”d GP Turbine Rotor
Assembly, (B) a close-up view of the adjacent post (tenon) separations and
(C) a close-up view of representative fret wear observed around the bolt holes.
(A) - 0.4X, (B) - 2X
&
(C) - 3X.
Approximate Magnifications:
The black arrows denote pin slots with cracks. The pin slots were identified “A”
through “H” counterclockwise as shown in View (A).
The small b/w arrows denote the slots with missing blade pieces.
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Page 6 of 23
+d
$
H
tr Figure 2. Views. of the adjacent pin slots showing (A) two primary .fracture origins (thick white arrows), two sec0ndar.y fracture
H
..
6\ 2
R,
!?
0
wl
L
R,
0
0
origins (thin white arrows) and (C) two overload fractures on the outboard side of each of the pin slots.
All - 15X.
..
,
c
(
&
.
b
..
2
N
..
N
0
wl
w
N
0
0
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Page 7 of 23
Figure 3. Views of the adjacent post fractures that initiated at the pin slots showing
(A & B) the two primary fractures and their origins (thick white arrows) and one small secondary fracture origin (thin white arrow),
(C) a secondary fracture and origin (thin white arrow) and (D) an overload fracture.
All - 1OX.
The black arrows denote differences in oxidation indicative of fracture progression.
The black brackets denote the overload portions of the fractures.
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. Page8of23
Figure 4. Views of representativ.e cracks (arrows) emanating from the corners of pin
slots.
(A & B) The cracks on the outboard side of pin slot “H” as identified in Figure 1A.
(C & D)’The cracks on the outboard sides of pin slots “G” & “E”, respectively, as
identified in Figure 1A.
(A, B & D) - 30X &
(C) - 15X.
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Figure 5. Views of the cracks and laboratory exposed fractures emanating from the
corners on the outboard side of pin slot “D” as identified in Figure IA.
(A & C) - 30X
&
(B & D ) - 2 O X .
The brackets in Views (B & D) denote the pin slot. Note that the origins are not shown
balanced around the pin slot because one fracture is the inboard side of the fracture and
the other is the outboard side.
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Figure 6. Scanning electron microscope (SEM) photographs of one of the primary post
separation fractures .
(A) An overall view.
(B & C) Close-up
photographs of the two
locations indicated.
Arrows “A”, “B”, “C” &
“D” in View (A) indicate
the locations shown in
Figure 7, Views (A, B, C
& D), respectively.
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Figure 7. (A, B, C & D) SEM close-up photographs of the post separation locations
denoted by Arrows “A”, “By’,“C” & “D” in Figure 6A, respectively.
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Figure 9. SEM photographs of one of the exposed fractures at pin slot “D” (reference
Fiiure 5B).
An overall view.
(B, C & D) Close-up photographs of the locations indicated by arrows “B”, “C” & “D”,
respectively, in View (A).
(4
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Figure 15; SEM photographs of (A & B) one of the locations on the outer diameter of
the cone with heavy damage and (C & D) the score marks on the outer diameter.
Magnifications:
(A) - 30X, (B) - lOOX (C) - 200X &
(D) - 500X.
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Figure 14. Scanning electron microscope (SEM) photographs of one of the locations
on the outer diameter of the cone with light damage.
Magnifications:
(A) - 5 0 X , (B) - 200X &
(C & D) - 500X.
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Figure 1’3; Close-up views of the Rear Compressor Cone.
(A) The outer diameter at location “B” as noted in Figure 12A.
(B) The outer diameter, (C) the edge and (D) the inner diameter at location “C” as noted
in Figure 12A.
(E) The outer diameter at location “D” as noted in Figure 12A.
All - 7.5X.
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Figure 12. (A) Overall view and
close-up views of the Rear Compressor
Cone showing:
(B) the outer diameter,
(C) the edge and (D) the inner diameter.
(B, C & D) - 7.5X.
(A) - 1.3X &
Letters “B”, “C”& “D” in View (A)
denote the locations shown in Figure 13.
The white arrow in View (B) denotes
the location of the score marks on the
outer diameter.
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Figure 11. (A) Optical photomicrograph of a cross section adjacent and parallel to one
of the pin slots showing the fracture with (B & C) photomicrographs showing the grain
sizes at the forward and aft locations, respectively, of the disk rim.
Arrow “0” in View (A) denotes the fracture origin.
Magnifications:
(A) - 25X
&
(B & C) - 1OOX.
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Figure 10. (A, B, C & D) Optical photomicrographs of a cross section adjacent and
parallel to one of the pin slots showing successive locations along the length of the
fracture.
Arrow “0”in View (A) denotes the fracture origin.
The white arrows in Views (A & B) denote the thick oxides at and near the fracture
origin.
The white arrow in View (C) denotes the oxidation associated with grain boundary
cracks.
Not etched.
Magnification: All - 200X.
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.
Figure 19; Optical photomicrographs of cross-sections (A & B) through the remaining
tip portion of one of the longest blades and (C & D) through one of the blade tip
shrouds. Electrolytically etched with 10% Oxalic.
(A & C) - 500X
&
(B & D) - 1OOOX.
Magnifications:
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(E) Unidentified pieces @ I .2X.
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Figure 17. Photomicrographs of (A) a
circumferential cross section through the thin end
of the cone showing the deformed shape and
(B, C, D & E) an axial section through a heavily
damaged portion of the cone showing the thick
oxides (white arrows), a tight crack (black arrow,
View E) and some deformed features in the
microstructure (bracket, View E).
Etched with 90% NaOH + 10% H202.
Magnifications:
(A & B) - 25X,
(C) - 200X, (D) - IOOX &
(E) - 500X.
“I” denotes the inner diameter.
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Figure 16. SEM photographs of the inner diameter of the cone with (A & B) heavy
damage and (C & D) light damage.
(A) - lOOX, (B) - 300X (C) - iOX
&
(D) - 200X.
Magnifications:
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1-20
A : C102372
FS
: 6426
EDX @ red crosshair
Particle-D
Prs t :None
LSec : 53
11: 1 3 :24
7-Mar- 0
-
---
\."L
0.90
1.80
2.70
SI
3.60
L
4.50
5.40
6 30
I
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7.20
8.10
9.00
09/08/00
10:38
'IP002 385 2429
HONEYWELL PS&I
@
003
I
Looper, Dave
From:
Sent:
To:
Cc:
Subject:
Welch, Jim
Thursda March 02.20007:16 AM
Looper, gave
Hanenbur , Jason; Elliott. Paul; Morgan, Mike
T53L138 &N LE24073
Records at CCAD show this englne was run in test cell at CCAD on 8 april 1993-----There is no record of it having been
repaired or overhauled, only tested
1

Honeywell EngineTeardown Report-T53-L-13B, SN LE-24073

Transcription

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