Here are the Navy's report about the recent V-22 engine fire. It requires a major redesign. I wish our military would modernize and stop using ALL CAPS.

HMR-EI FINAL REPORT

UNCLAS 161451Z FEB 07 COMNAVAIRSYSCOM PATUXENT RIVER MD(UC)

TO VMMT 204(uc)
DCMA BELL HELICOPTER TEXTRON(UC)
AIG 423
AL 423(UC)
CC CG 2ND MAW(uc)
COMMARFORCOM(uc)
MALS 29(UC)
MALS 26(uc)
NAVAIRDEPOT CHERRY PT NC(UC)
PEOASWASM PATUXENT RIVER MD(UC)
NAVICP PHILADELPHIA PA(UC)
CENNAVAVNTECHTRA DET NEW RIVER NC(UC)
COMOPTEVFOR NORFOLK VA(UC)
MAG 26(uc)
MAG 29(uc)
ASC(UC)
CG II MEF(uc)
HQ AETC(UC)
HQ AFSOC(UC)
MV 22 FIT(uc)
HURLBURT BDP(UC)
COMNAVAIRFOR SAN DIEGO CA
DCMA BELL HELICOPTER TEXTRON(UC)
DCMA BOEING PHILADELPHIA(UC)
412OG(UC)
VMM 263(uc)
VMM 162(uc)
USSOCOM MESSAGE CENTER(MC)
COMNAVAIRSYSCOM PATUXENT RIVER MD(UC)

PASS TO OFFICE CODES:
FM COMNAVAIRSYSCOM PATUXENT RIVER MD//DRPO// TO VMMT TWO ZERO
FOUR//QA/AMO/AAMO/QAO// COMNAVAIRSYSCOM PATUXENT RIVER
MD//4.0/4.1.1.2/4.3/
4.3.5/PMA275//
DCMA BELL HELICOPTER FORT WORTH TX//DCMDW-RKTB// AIG 423 INFO CG SECOND
MAW//ALD/ALD-B// COMMARFORCOM//ALD-B// MALS TWO NINE//AMO// MALS TWO
SIX//AMO/ASO/CLASSDESK/FIT/SUPPLY//
NAVAIRDEPOT CHERRY PT NC//4.1.10.1/V-22FST/V-22FST.1/
V22FST.2//
PEOASWASM PATUXENT RIVER MD//4.1.1.2/PMA-275// VMMT TWO ZERO
FOUR//AMO/QA/QAO// NAVICP PHILADELPHIA PA//0322/0322A// CENNAVAVNTECHTRA
DET
NEW RIVER NC//QA// COMOPTEVFOR NORFOLK VA//585// MAG TWO SIX//ADJ// MAG
TWO
NINE//ADJ// ASC WRIGHT PATTERSON AFB OH//ASC-LU/LUXL// CG II
MEF//ALD-B// HQ
AETC RANDOLPH AFB TX//LGXX// HQ AFSOC HURLBURT FLD FL//V-22/XPQV/XPRX//
MV
TWO TWO FIT 18FLTS HURLBURT FLD FL//DET2// COMNAVAIRFOR SAN DIEGO
CA//N412C/N421H// DCMA BELL HELICOPTER FORT WORTH TX//DCMDW-RKOA/
DCMDW-RKTB//
DCMA BOEING HELICOPTERS PHILADELPHIA PA//REO/RETA/RETC// HQ USSOCOM
MACDILL
AFB FL 418FLTS EDWARDS AFB CA VMM TWO SIX THREE//QA// HMM ONE SIX
TWO//QA/AAMO/MO// COMNAVAIRSYSCOM PATUXENT RIVER MD//DRPO// //N04790//
MSGID/GENADMIN/COMNAVAIRSYSCOM PAX DRPO// REF/A/DOC/COMNAVAIRFORINST
4790.2
CH-1/01MAY2006// REF/B/MSG/COMNAVAIRSYSCOM PATUXENT
RIVER/130640ZDEC2006//
REF/C/MSG/COMNAVAIRSYSCOM PATUXENT RIVER/191646ZDEC2006//
REF/D/MSG/COMNAVAIRSYSCOM PATUXENT RIVER/211911ZDEC2006//
REF/E/MSG/COMNAVAIRSYSCOM PATUXENT RIVER/212200ZDEC2006//
REF/F/MSG/COMNAVAIRSYSCOM PATUXENT RIVER/112228ZFEB2007//
REF/G/MSG/COMNAVAIRSYSCOM PATUXENT RIVER/292006ZJAN2007//
REF/H/MSG/NAVAIRDEPOT CHERRY PT/291538ZJAN2007// NARR/REF A IS
INSTRUCTION
FOR THE NAVAL AVIATION MAINTENANCE PROGRAM.
REF B IS VMMT-204 HMR-EI REQUEST FOR BUNO 166482 LH INBOARD EAPS
PRESSURE
TUBE, RCN V52842-06-0208. REF C IS V22 FST PRELIMINARY EI REPORT LH
INBOARD
EAPS PRESSURE TUBE, RCN V52842-06-0208. REF D IS
VMMT-204 EI REQUEST FOR LH OUTBOARD EAPS BLOWER, RCN N52842-06-0212.
REF E IS VMMT-204 HMR-EI REQUEST FOR LH INBOARD EAPS BLOWER, RCN
V52842-06-0215. REF F IS V22FST FINAL EI REPORT FOR LH EAPS START
CONTROL
VALVE, RCN N52842-06-0213. REF G IS PMA275 FLIGHT CLEARANCE RESTRICTION
ENGINE AIR PARTICLE SEPARATOR (EAPS) V-22 EAPS FCR-07.
REF H IS AFB-73 TCTO 1V-22(C)B-548 EAPS BLOWER RADIAL PLAY INSPECTION.//
POC/GRAVES, MICHAEL/-/NAVAIRDEPOT CHERRY PT NC/LOC:V-22FST.1
/DSN:451-/TEL:252-464-8451// RMKS/THIS MESSAGE WAS AUTO GENERATED FROM
THE
NAMDRP WEBSITE FOR NON-WEB SITE CAPABLE ORGANIZATIONS. THE REPORT WAS
ORIGINATED BY:
------ NAVAIRDEPOT CHERRY PT NC/V-22FST.1.
IF RESPONSE VIA WEB SITE IS NOT POSSIBLE, TO: LINE RECIPIENTS SHOULD
ADDRESS
RESPONSE DIRECTLY TO:
------ NAVAIRDEPOT CHERRY PT NC/V-22FST.1 WHEN APPROPRIATE. THIS
DISCREPANCY
REPORT WILL BE PROCESSED VIA THE NAMDRP WEBSITE. FOR FURTHER DETAILS OR
REAL
TIME STATUS VISIT THE NAMDRP WEB SITE AT:
NAMDRP.NAVAIR.NAVY.MIL.
1. VMMT-204/52842
2. V52842-06-0208
3. AIRCRAFT T/M/S: MV-22B, BUNO: 166482, NOMENCLATURE: TUBE ASSY,
PRESSURE,
METALLIC, P/N: 901-083-374-101, S/N: N/A, LOT/BATCH NR:
N/A, NSN: 4720 - LLV220152, CONTRACT NR: N/A, WUC: 71810343 4.
NAVAIRDEPOT
CHERRY PT NC 5. ICN: WC2EI-V22-0106-06M 6. TIME SINCE NEW: 419.3
FLIGHT
HOURS TIME SINCE REWORK: N/A 7. LAST REPAIR DATE: NA 8. BACKGROUND:
A.
IAW REF A, REF B REPORTED UPON LANDING AT MCAS NEW RIVER, THE AIRCREW
RECEIVED HYD 3 FAIL ON THE CDU/EICAS, IMMEDIATELY FOLLOWED BY A HYD 3
SYSTEM
PRESS LOW. TWO SECONDS LATER, A LANDING GEAR - NOSEWHL STEER FAIL
POSTED.
SIX SECONDS LATER, A FCS UTIL SYSTEMS INOP CAUTION POSTED. TEN SECONDS
LATER, LH NACELLE FIRE POSTED. THE TOTAL ELAPSED TIME FROM LANDING TO
FIRE
WAS 54 SECONDS.
THE FIRE WAS VISUALLY CONFIRMED BY THE AIRCREW. POST FLIGHT MAINTENANCE
DETERMINED THE LH INBOARD ENGINE AIR PARTICLE SEPARATOR
(EAPS) HYDRAULIC PRESSURE TUBE, P/N 901-083-374-101, RUPTURED. THE
RUPTURED
TUBE DEPLETED HYDRAULIC SYSTEM 3 AND HYDRAULIC FLUID FROM THE RUPTURED
TUBE
DRAINED INTO THE IR SUPPRESSOR RESULTING IN THE FIRE. REF C ASSIGNED EI
CONTROL NUMBER AND THE FAILED TUBE WAS SUBMITTED TO NAVAIR CHERRY POINT
MATERIALS LAB FOR FAILURE ANALYSIS.
B. REF B EI IS ONE OF SEVERAL ENGINEERING INVESTIGATIONS RELATED TO THE
EAPS
PRESSURE TUBE FAILURE ON BUNO 166482. REFER TO REF D FOR THE ONGOING EI
ON
THE LH OUTBOARD EAPS BLOWER, REF E FOR THE ONGOING EI ON THE LH INBOARD
EAPS
BLOWER, AND REF F FOR THE FINAL EI REPORT ON THE LH EAPS START CONTROL
VALVE.
9. DESCRIPTION OF FINDINGS: A. INSPECTION OF THE LOWER NACELLE FIRE
DETERMINED THE LH INBOARD EAPS PRESSURE TUBE, P/N 901-083-374-101,
RUPTURED
AT THE UPPER END OF THE TUBE. THE FAILURE OCCURRED IN THE TUBE UNION
FITTING IMMEDIATELY ADJACENT TO THE ORBITAL WELD. THE FRACTURED END OF
THE
TUBE WAS NO LONGER LOCATED IN LINE WITH THE UNION FITTING, AND WAS
DISPLACED FORWARD OF THE UNION FITTING. THE FRACTURED END OF THE TUBE
WAS
RESTING AGAINST A VERTICAL STRUCTURAL INTERCOSTAL IN THE LOWER NACELLE.
INSPECTION OF THE TUBE AS INSTALLED ON THE AIRCRAFT NOTED THE FRACTURED
END
OF THE TUBE WAS FREE TO MOVE IN ANY DIRECTION. INSPECTION OF THE
REMAINDER
OF THE PRESSURE TUBE AS INSTALLED ON THE AIRCRAFT DID NOT REVEAL ANY
ANOMALIES. INSPECTION OF THE HYDRAULIC BLOCK CLAMPS WHICH MOUNT THE
PRESSURE, RETURN AND CASE DRAIN TUBES TO THE FORWARD VERTICAL FIREWALL
IN
THE LOWER NACELLE ALSO DID NOT NOTE ANY ANOMALIES, SUCH AS DEFLECTION OR
DISPLACEMENT OF THE BLOCK CLAMP MATERIAL. DURING THE PROCESS OF
REMOVING
THE FAILED TUBE FROM THE AIRCRAFT, CARE WAS EXCERCISED TO DETERMINE IF
THERE
WAS ANY NOTABLE PRELOAD ON THE TUBE.
NO SIGNFICANT TUBE INSTALLATION PRELOAD WAS NOTED ON THE FAILED TUBE,
ALTHOUGH ANY CONTRIBUTION TO PRELOAD FROM THE FAILED FITTING COULD NOT
BE
DETERMINED. A NEW REPLACEMENT LH INBOARD HYDRAULIC PRESSURE TUBE WAS
SUBSEQUENTLY INSTALLED ON THE AIRCRAFT AND NO NOTABLE PRELOAD WAS NOTED
DURING INSTALLATION.
B. THE LH INBOARD EAPS PRESSURE TUBE, P/N 901-083-374-101, IS 0.25 INCH
DIA
X 0.022 INCH WALL THICKNESS 3AL-2.5V TITANIUM TUBE PER AMS 4945. THE
TUBE
IS 31.75 INCHES LONG WITH 10 BENDS. A 90 DEGREE RYNGLOK ELBOW END
FITTING,
P/N L50221T04, IS INSTALLED ON THE LOWER END OF THE TUBE. A MALE BEAM
SEAL
TO WELD UNION FITTING ADAPTER, PER BELL CONTROL STANDARD 50-095, IS
INSTALLED ON THE UPPER END OF THE TUBE. THE UNION FITTING IS WELDED TO
THE
TITANIUM TUBE THROUGH AN ORBITAL WELDING PROCESS PER BELL PROCESS SPEC
4549
AT BELL HELICOPTER. FOLLOWING INSTALLATION OF THE RYNGLOK AND UNION
FITTINGS, THE TUBE ASSY IS PROOF PRESSURE TESTED TO 10000 PSI AT BELL
HELICOPTER. THE UNION FITTING IS CONSTRUCTED OF 6AL-4V TITANIUM PER AMS
4965 WITH AN OPERATING PRESSURE RATING OF 5000 PSI, PROOF PRESSURE
RATING OF
10000 PSI AND BURST PRESSURE RATING OF 20000 PSI.
ONE OF THE WRENCH FLATS OF THE UNION FITTING WAS MARKED WITH P/N
R42542T-04,
WHICH THE BELL CONTROL STANDARD 50-095 INDICATES IS A QUALIFIED END
FITTING
FROM RESISTOFLEX CORP. EXTERNAL MARKINGS ON THE PRESSURE TUBE WERE AS
FOLLOWS: P/N 901-083-374-101, HEAT LOT 383625-1.
C. FOLLOWING REMOVAL FROM THE AIRCRAFT, THE FAILED TUBE ASSEMBLY WAS
INSPECTED ON A COORDINATE MEASURING (VECTOR) MACHINE AT NAVAIR CHERRY
POINT
TO VERIFY THE TUBE WAS PROPERLY BENT IAW THE BEND DATA REQUIREMENTS
SPECIFIED ON BELL ENGINEERING DRAWING 901-083-374. THE VECTOR MACHINE
ANALYSIS VERIFIED THE TUBE MET THE BEND DATA REQUIREMENTS.
D. THE FOLLOWING IS A SUMMATION OF THE MATERIALS LAB FAILURE ANALYSIS
REPORT (AVAILABLE UNDER THIS RCN ON NAMDRP). THE TUBE ASSEMBLY WAS
MARKED
WITH A RED TICK MARK TO REFERENCE THE 12 OCLOCK POSITION AS INSTALLED IN
THE
LOWER NACELLE. EXAMINATION OF THE OPPOSING ENDS OF THE FRACTURED TUBE
AND
TUBE FITTING AT HIGHER MAGNIFICATIONS REVEALED THE ORBITAL WELD STOP
POSITION CORRESPONDED, WITHIN PLUS MINUS 5 DEGREES, TO THE 12 OCLOCK
POSITION AS INSTALLED ON THE AIRCRAFT. THE FRACTURE SURFACE WAS ANGLED
ROUGHLY 25 TO 30 DEGRESS TO THE TUBE BASAL PLANE, WITH THE FURTHEST AND
CLOSEST POINTS OF THE FRACTURE 0.230 INCH AND 0.175 INCH FROM THE FRONT
PLANE OF THE FITTING, RESPECTIVELY. MACROSCOPIC EXAMINATION OF THE TUBE
AND
FITTING FRACTURE SURFACES REVEALED A RELATIVELY PLANAR FRACTURE REGION
NORMAL TO THE TUBE AXIS. THE FITTING FRACTURE WAS APPROXIMATELY 0.230
INCH
FROM THE INLET SIDE OF THE FITTING.
E. THE FITTING HAD AN OUTER DIAMETER OF 0.255 INCH AND AN INNER DIAMETER
OF
0.200 INCH. THE SURFACE FINISH VALUES ON THE INTERIOR AND EXTERIOR
SURFACES
RANGED BETWEEN 63 AND 32 MICRO-INCH. WITH THE EXCEPTION OF A STRAW
COLORED
BAND OF CONTAMINATION AROUND THE FITTING OUTER CIRCUMFERENCE, THERE WERE
NO
OTHER DEFECTS NOTED ON THE FITTING. NO WELD BEAD MATERIAL WAS OBSERVED
ADJACENT TO THE FRACTURE AROUND EITHER THE INNER OR OUTER CIRCUMFERENCE
OF
THE FITTING. THE TUBE INNER DIAMETER BEYOND THE WELD BEAD WAS MEASURED
AT
0.204 INCH AND THE OUTER DIAMETER WAS MEASURED AT 0.251 INCH. THE WELD
BEAD, WHICH AVERAGED 0.125 INCH WIDTH, WAS COMPRISED OF RELATIVELY LARGE
COLUMNAR GRAINS. RELATIVELY LARGE GRAINS WERE ALSO OBSERVED ON THE
INTERIOR
SURFACE OF THE TUBE ADJACENT TO THE WELD BEAD. THE HEIGHT OF THE WELD
BEAD
REINFORCEMENT ON THE INNER DIAMETER OF THE TUBE WAS MEASURED AT
APPROXIMATELY 15 MILS, WHICH MET THE REQUIREMENTS OF THE BELL PROCESS
SPECIFICATION, ALTHOUGH THE REINFORCMENT OF THE WELD BEAD WAS CONSIDERED
TO
BE EXCESSIVE WHEN COMPARED TO AWS WELDING SPECIFICATIONS. AWS WELDING
SPECS
SPECIFY A TYPICAL WELD BEAD REINFORCEMENT OF ONE-THIRD THICKNESS (0.007
INCH
FOR A 0.022 INCH WALL THICKNESS TUBE). INSPECTION OF THE UNION FITTING
INNER AND OUTER SURFACES REVEALED CIRCUMFERENTIAL MACHINING LINES
ADJACENT
TO THE FRACTURE.
F. THE TUBE AND FITTING FRACTURE SURFACES WERE EXAMINED UNDER SCANNING
ELECTRON MICROSCOPE WHICH INDICATED THE PRESENCE OF A FATIGUE CRACK.
THE
CRACK ORIENTATION INDICATED FATIGUE CRACKING HAD INITIATED ON THE INNER
CIRCUMFERENCE OF THE TUBE FITTING. THE FATIGUE CRACK EXTENDED
APPROXIMATELY
0.225 INCH OR THROUGH AN 85 DEGREE ARC AROUND THE INNER CIRCUMFERENCE OF
THE
FITTING. THE RELATIVE POSITIONS OF THE TWO ENDS OF THE FATIGUE CRACK
ZONE
TO THE
12 OCLOCK (UP) POSITION AS INSTALLED ON THE AIRCRAFT WERE DETERMINED TO
BE
BETWEEN APPROXIMATELY 55 DEGREES AND 140 DEGREES ON THE AFT SIDE OF THE
TUBE. SUBSEQUENT EXAMINATIONS AT HIGHER LEVELS OF MAGNIFICATION
CONFIRMED
THE PRESENCE OF MULTIPLE FATIGUE CRACK INITIATION SITES (INDICATED BY
RATCHET AND RADIAL MARKS) ALONG THE INNER CIRCUMFERENCE OF THE TUBE
FITTING.
MULTIPLE FATIGUE CRACK INITIATION SITES IS INDICATIVE OF FATIGUE CRACK
GROWTH AT THE HIGH END OF THE SN CURVE, WHICH WOULD INDICATE PRESENCE OF
HIGH STRESSES.
EXAMINATIONS AROUND THE INNER CIRCUMFERENCE ADJACENT TO THE PLANAR
REGION
REVEALED MACHINING LINES ASSOCIATED WITH THE INNER DIAMETER OF THE
FITTING
ONLY AT ONE END OF THE PLANAR REGION. A SMALL AMOUNT OF CRACKING ALONG
THE
FITTING SIDE OF THE WELD BEAD WAS EVIDENT. A SHALLOW SOMEWHAT
DISCONTINUOUS
CIRCUMFERENTIAL GROOVE WAS PRESENT WITHIN THE WELD BEAD ADJACENT TO MOST
OF
THE PLANAR REGION, HOWEVER, CRACKING DID NOT APPEAR TO HAVE INITIATED
FROM
THE GROOVE. WHILE MOST OF THE FRACTURE SURFACE CONSISTED OF SMALL
PLATEAUS
OR FACETS APPEARING TO EXHIBIT DISCONTINUOUS STRIATIONS, A FEW AREAS
APPROACHING THE OUTER CIRCUMFERENCE OF THE TUBE ASSEMBLY CONTAINED
CLASSIC
FATIGUE CRACK STRIATIONS. THE STRIATIONS APPEARED TO BE RELATIVELY
COARSE
(0.3 MICROMETER). THE FRACTURED UNION FITTING WAS SENT TO BELL
HELICOPTER
FOR A MORE ACCURATE STRIATION SPACING/COUNTING ASSESSMENT. BELL
HELICOPTER
EXAMINATION INDICATED APPROXIMATELY 24000 STRIATIONS. THE FRACTURED
UNION
FITTING WAS ALSO EXAMINED TO DETERMINE IF ANY DEFECTS WERE PRESENT
WITHIN
THE FITTING ALONG THE FUSION LINE. EXAMINATION OF THE FITTING REVEALED
NO
EVIDENCE OF WELD BEAD MATERIAL ALONG EITHER THE INNER CIRCUMFERENCE OF
THE
FATIGUE CRACK OR THE REST OF THE FRACTURE SURFACE. THE SAME MULTIPLE
ORIGINS ALONG THE INNER CIRCUMFERENCE THAT WERE PRESENT ON THE FRACTURED
TUBE SURFACCE WERE OBSERVED ON THE FRACTURED FITTING.
CRACKS OR GAPS WERE OBSERVED NORMAL TO THE DIRECTION OF CRACK
PROPAGATION
NEAR ONE END OF THE FATIGUE FRACTURE. THIS TYPE OF CRACKING IS OFTEN
OBSERVED DURING THE LATER STAGES OF FATIGUE CRACKING. SMALL CRACKS
PARALLEL
TO THE FATIGUE CRACK WERE ALSO OBSERVED WITHIN THE FITTING MATERIAL,
WHICH
ARE TYPICALLY OBSERVED ADJACENT TO TENSILE FRACTURE SURFACES. THE
IMPLICATION HERE WOULD BE THAT THE OPPOSING FRACTURE SURFACES WERE
SEPARATING RATHER QUICKLY AT THIS LOCATION.
G. VICKERS MICROHARDNESS MEASUREMENTS WERE MADE ACROSS THE FAILED TUBE
AND
FITTING END CROSS-SECTIONS AS WELL AS BELL AND BOEING EXEMPLAR TUBE AND
FITTING ASSEMBLY CROSS-SECTIONS. THE AVERAGE TUBE HARDNESS VALUES
APPROXIMATELY 0.25 INCH FROM THE WELD BEADS WAS 259 HV300, 252 HV300,
AND
250 HV300 FOR THE FAILED, BELL EXEMPLAR, AND BOEING EXEMPLAR TUBE
ASSEMBLIES, RESPECTIVELY. THE HARDNESS VALUES WITHIN THE WELD BEADS
AVERAGED 342 HV300, 331 HV300, AND 337 HV300 FOR THE FAILED TUBE
ASSEMBLY,
BELL EXEMPLAR, AND BOEING EXEMPLAR, RESPECTIVELY. THE AVERAGE FITTING
HARDNESS VALUES APPROXIMATELY 0.25 INCH FROM THE WELD BEADS WERE 368
HV300,
371 HV300, AND 374 HV300 FOR THE FAILED, BELL EXEMPLAR, AND BOEING
EXEMPLAR
TUBE ASSEMBLIES, RESPECTIVELY. THE AVERAGE HARDNESS VALUES WITHIN THE
HEAT-AFFECTED ZONES ON EITHER SIDE OF THE WELD BEADS DID NOT VARY
SIGNIFICANTLY FROM THE VALUES OF THE TUBE AND FITTING APPROXIMATELY 0.25
INCH AWAY.
H. THE FAILED TUBE AND UNION FITTING WERE CHEMICALLY ANALYZED THROUGH
DISPERSIVE X-RAY ANALYSIS. THE RESULTS OF THE ANALYSIS INDICATED THE
TITANIUM TUBE MATCHED THE REQUIRED MATERIAL COMPOSITION OF 3AL-2.5V
TITANIUM
PER AMS 4945H. THE FITTING ALSO MATCHED THE REQUIRED CHEMICAL
COMPOSITION
OF 6AL-4V TITANIUM, PER AMS 4965H. NO ANOMALIES WERE NOTED WITH THE
CHEMICAL COMPOSITION OF THE TITANIUM TUBE OR THE TUBE FITTING.
I. THE FAILED FITTING AND ONE OF THE BELL EXEMPLAR WELDED TUBE
ASSEMBLIES
WAS SENT TO LAMBDA TECHNOLOGIES FOR RESIDUAL STRESS ANALYSIS. RESIDUAL
STRESS MEASUREMENTS WERE MADE ADJACENT TO ONE END OF THE FATIGUE CRACK
ON
THE INTERIOR OF THE FITTING END OF THE FRACTURED TUBE. THE RESIDUAL
STRESS
MEASUREMENTS, WHICH WERE MADE AT INCREASING DEPTHS FROM THE SURFACE,
REVEALED RELATIVELY LOW (-2.0 TO
-4.1 KSI) COMPRESSIVE RESIDUAL STRESSES DOWN TO 0.0015 INCH DEPTH AND A
-13.1 KSI RESIDUAL COMPRESSIVE STRESS 0.0032 INCH BELOW THE SURFACE.
THE
RESULTS OF THE RESIDUAL STRESS MEASUREMENTS INDICATE PRESENCE OF
RELATIVELY
SMALL RESIDUAL STRESSES.
J. BELL HELICOPTER PERFORMED A STRESS ANALYSIS ON THE LH INBOARD EAPS
PRESSURE TUBE. THE ANALYSIS INDICATED PRESSURE SPIKES (2000 TO 3000 PSI
MAGNITUDE ABOVE THE NOMINAL 5000 PSI SYSTEM PRESSURE) COMBINED WITH
BENDING
LOADS IN THE TUBE WOULD CAUSE AN AREA OF HIGH STRESSES IN THE UNION
FITTING
ADJACENT TO THE WELD. REF E EI OF THE LH INBOARD EAPS BLOWER FOUND THE
BLOWER WAS OPERATING IN A DEGRADED CONDITION, AND THE BLOWER WAS
PRODUCING
2000 TO 3000 PSI PRESSURE SPIKES (ABOVE 5000 PSI) AT THE BLOWER MOTOR
INLET
PORT. THE BELL STRESS AND FATIGUE ANALYSIS INDICATED THAT AS LITTLE AS
0.06
INCH DISPLACEMENTOF THE TUBE AT THE CLAMPING LOCATIONS OR A 1.4 DEGREE
OR
LARGER MISALIGNMENT OF THE UNION FITTING IN THE LONGERON WILL RESULT IN
ENOUGH PRELOAD AND BENDING STRESS, WHEN COMBINED WITH THE 2000 TO 3000
PSI
(ABOVE 5000 PSI) PRESSURE SPIKES, TO INITIATE AND GROW FATIGURE CRACKS
IN
THE TUBE. THE BELL ANALYSIS INDICATED A COMBINATION OF BOTH THE
PRESSURE
SPIKES AND THE BENDING LOADS WERE REQUIRED TO RESULT IN FATIGUE CRACK
INITIATION.
10. CONCLUSIONS: A. THE LH INBOARD EAPS PRESSURE TUBE FAILED AS A
RESULT OF
FATIGUE CRACK INITIATION AND PROPAGATION FOLLOWED BY OVERLOAD. THE
FATIGUE
CRACK INITIATED AT MULTIPLE LOCATIONS ON THE INNER DIAMETER OF THE
FITTING
IN THE HEAT AFFECTED ZONE ADJACENT TO THE FITTING SIDE OF THE WELD BEAD.
MULTIPLE FATIGUE CRACK ORIGINATION SITES IS INDICATIVE OF THE PRESENCE
OF
HIGH STRESSES.
METALLOGRAPHIC EXAMINATION REVEALED NO APPARENT DEFECTS IN THE FAILED
TUBE
AND FITTING MICROSTRUCTURE, AND THE MICROSTRUCTURE OF THE FAILED TUBE
MATCHED THE EXEMPLAR BELL AND BOEING TUBES THAT WERE SUBMITTED FOR
ANALYSIS
DURING THIS EI. THE CHEMICAL COMPOSITION OF THE FAILED TUBE WAS IN
ACCORDANCE WITH AMS 4945H AND THE FITTING WAS IN ACCORDANCE WITH AMS
4965H.
MICROHARDNESS PROFILE MEASUREMENTS RECORDED ACROSS THE FAILED TUBE AND
FITTING WERE COMPARABLE TO MICROHARDNESS MEASUREMENTS FROM THE BELL AND
BOEING EXEMPLAR TUBES.
THE RESIDUAL STRESS ANALYSIS OF THE FAILED TUBE AND FITTING INDICATED
THE
PRESENCE OF RELATIVELY MINOR RESIDUAL STRESSES IN THE TUBE FOLLOWING THE
WELDING PROCESS. NO SIGNIFICANT METALLURGICAL OR MANUFACTURING ISSUES
WERE
NOTED DURING THE FAILURE INVESTIGATION OF THE TUBE OR UNION FITTING.
B. THE ROOT CAUSE OF THE LOADING AND STRESSES THAT RESULTED IN FATIGUE
CRACK
INITIATION AND GROWTH WAS MOST LIKELY THE RESULT OF SIGNIFICANT
HYDRAULIC
PRESSURE SPIKES AND EXTERNALLY APPLIED BENDING
LOADS AT THE UNION FITTING. THE SOURCE OF THE PRESSURE SPIKES WAS
DETERMINED TO BE THE LH INBOARD EAPS BLOWER, WHICH WAS OPERATING IN A
DEGRADED CONDITION. AS PART OF THE ONGOING EI OF THE INBOARD EAPS
BLOWER,
REFER TO REF E, AN ACCEPTANCE TEST PROCEDURE (ATP) TEST INDICATED THE LH
INBOARD BLOWER WAS ONLY OPERATING AT ONLY 10 PERCENT OF THE NORMAL 12500
RPM, SIGNIFICANT 2000 TO 3000 PSI PRESSURE SPIKES ABOVE THE NORMAL 5000
PSI
PRESSURE WERE NOTED AT THE BLOWER PRESSURE INLET PORT, AND SIGNIFICANT
INLET
AND CASE DRAIN FLOW VARIATIONS WERE NOTED AT THE INLET AND CASE DRAIN
PORTS.
DISASSEMBLY OF THE LH INBOARD EAPS BLOWER REVEALED A SEVERELY WORN
IMPELLER
SHAFT AND JOURNAL BEARING. THE SEVERELY WORN SHAFT AND JOURNAL BEARING
ALLOWED THE IMPELLER TO CONTACT THE SHROUD, WHICH RESULTED IN THE RAPID
PRESSURE AND FLOW VARIATIONS WITHIN THE EAPS BLOWER. THE POTENTIAL
CAUSES
OF THE BENDING LOADS IN THE TUBE UNION FITTING MAY BE FROM ONE OR A
COMBINATION OF THE TWO FOLLOWING CAUSES:
(1) THE BLOWER PRESSURE SPIKES, FLOW VARIATIONS AND EXCESSIVE BLOWER
VIBRATIONS MAY HAVE INDUCED RELATIVE MOTION IN THE TUBE WHICH WOULD BE
REACTED AT THE WELDED UNION FITTING. SINCE THE UNION FITTING IS SECURED
TO
THE INTERCOSTAL FRAME IN THE LOWER NACELLE, ANY RELATIVE MOTION OF THE
TUBE
WOULD NOT BE TRANSLATED TO THE FITTING AND WOULD RESULT IN AN AREA OF
BENDING STRESS AT THE FITTING.
(2) TUBE INSTALLATION AND PRELOAD FACTORS MAY HAVE ALSO CONTRIBUTED TO
BENDING LOADS IN THE UNION FITTING. BELL CONDUCTED A STRESS AND FATIGUE
ANALYSIS THAT INDICATES AS LITTLE AS 0.06 INCH DISPLACEMENTOF THE TUBE
AT
THE CLAMPING LOCATIONS OR A 1.4 DEGREE OR LARGER MISALIGNMENT OF THE
FITTING
IN THE LONGERON WILL RESULT IN ENOUGH PRELOAD AND BENDING STRESS, WHEN
COMBINED WITH THE 2000 TO 3000 PSI (ABOVE 5000 PSI) PRESSURE SPIKES, TO
INITIATE AND GROW FATIGURE CRACKS IN THE TUBE. THE BELL ANALYSIS
INDICATED
A COMBINATION OF BOTH THE PRESSURE SPIKES AND THE BENDING LOADS WERE
REQUIRED TO RESULT IN FATIGUE CRACK INITIATION.
C. THE EAPS PRESSURE TUBE RUPTURE RESULTED IN A SIGNFICANT AND RAPID
LOSS
OF HYDRAULIC FLUID FROM THE UTILITY HYDRAULIC SYSTEM. THE HYDRAULIC
FLUID
DRAINED THROUGH THE LOWER NACELLE AND INTO THE IR SUPPRESSOR, WHERE IR
SUPPRESSOR CENTERBODY TEMPERATURES ARE
SUFFICIENT TO IGNITE HYDRAULIC FLUID. BELL BOEING AND NAVAIR
PERFORMED A NACELLE FLUID DRAINAGE TEST ON A PRODUCTION AIRCRAFT IN
AMARILLO
ON 06 JAN 2007. THE FLUID DRAINAGE TEST INDICATED THE PRESENCE OF
MULTIPLE
FLUID FLOW PATHS THAT ALLOW FLAMMABLE FLUIDS (HYDRAULIC FLUID, FUEL,
OILS,
ETC) TO DRAIN FROM THE LOWER NACELLE COMPARTMENT INTO THE IR SUPPRESSOR.
WHILE THE LOWER NACELLE HAS FIRE DETECTION AND SUPPRESSION, THE IR
SUPPRESSOR DOES NOT HAVE FIRE DETECTION AND SUPPRESSION INSTALLED.
11. RECOMMENDATIONS:
A. REDESIGN THE EAPS HYDRAULIC SYSTEM TO WITHSTAND SYSTEM LEVEL EFFECTS
OF
A DEGRADED OR FAILING EAPS BLOWER. THE REDESIGN MAY INCLUDE BUT NOT BE
LIMITED TO REROUTING THE HYDRAULIC TUBE
INSTALLATIONS OR CHANGING TUBE MATERIALS AND CONSTRUCTION. THE
EAPS HYDRAULIC SYSTEM MUST BE ROBUST ENOUGH TO WITHSTAND THE SYSTEM
LEVEL
EFFECTS OF EAPS BLOWER FAILURES.
B. REDESIGN THE VMS SOFTWARE LEAK DETECTION AND ISOLATION SYSTEM TO
PREVENT
COMPLETE LOSS OF HYDRAULIC SYSTEM 3 IN A FAST LEAK SCENARIO. RAPID
DETECTION AND ISOLATION OF HYDRAULIC SYSTEM LEAKAGE WILL REDUCE THE
AMOUNT
OF FLUID THAT MAY BE EXPELLED INTO A POTENTIALLY FLAMMABLE AREA OF THE
AIRCRAFT, SUCH AS THE MIDWING OR THE LOWER NACELLE.
C. REDESIGN THE NACELLE DRAINAGE SYSTEM TO ROUTE FLUID LEAKAGE AWAY
FROM
THE HOT SECTIONS IN THE LOWER NACELLE AND IN THE IR SUPPRESSOR. THE
NACELLE
DRAINAGE SYSTEM SHOULD BE ABLE TO CONTROL AND DIRECT A LEVEL OF FLUID
THAT
WOULD BE EXPECTED TO BE EXPELLED IN THE LOWER NACELLE AS A RESULT OF A
HYDRAULIC SYSTEM FAILURE, FUEL FEED HOSE FAILURE, OR ENGINE OIL SYSTEM
FAILURE.
D. REDESIGN THE IR SUPPRESSOR SECTION TO ADD FIRE DETECTION AND
SUPPRESSION.
E. REDESIGN THE EAPS BLOWER FAILURE DETECTION SYSTEM TO PROVIDE AN
EARLY
DETECTION OF IMPENDING BLOWER FAILURES. THE CURRENT FAILURE DETECTION
SYSTEM, CONSISTING OF 500 PSI PRESSURE SWITCHES ON THE CASE DRAIN TUBES,
IS
NOT EFFECTIVE SINCE THE CASE DRAIN FLOW RESTRICTION ORIFICES WERE
REMOVED AS
PART OF INTERIM AFC 058/TCTO TCTO 1V-22(C)B-514.
12. RELATED INFORMATION: A. IN 1998, HISTORICAL RECORDS INDICATE A LH
INBOARD EAPS PRESSURE TUBE, P/N 901-081-346-103, RUPTURED ON EMD
V-22 D0010, BUNO 164942. THE FAILED TUBE FRACTURED NEAR THE WELD AREA
IN
THE HEAT AFFECTED ZONE DUE TO FATIGUE CRACKS INITIATED FROM MULTIPLE
ORIGINS
ON THE INSIDE DIAMETER OF THE TUBE (REF BELL HELICOPTER REPORT NO.
90198M-108). NO MANUFACTURING OR MATERIAL IMPERFECTIONS WERE OBSERVED
ON
THE FAILED TUBES. THE FAILED EMD TUBE WAS REPLACED WITH A NEW TUBE AND
THE
NEW REPLACEMENT TUBE SUBSEQUENTLY RUPTURED IN ONE OF THE BENDS.
TROUBLESHOOTING EVENTUALLY DETERMINED THE EAPS BLOWER HAD A CRACKED
CYLINDER
BLOCK WHICH CAUSED THE TUBE RUPTURES ON BUNO 164942. THE EAPS BLOWER
CYLINDER BLOCK WAS REDESIGNED TO ELIMINATE THE FAILURE MODE THAT
OCCURRED ON
THE BUNO 164942 EAPS BLOWER FAILURE.
B. REF G EAPS FLIGHT CLEARANCE RESTRICTION ALLOWED EAPS OPERATION UPON
COMPLIANCE WITH REF H. IN ADDITION, ERAC 408 (35 FLIGHT HOUR EAPS BLOWER
RADIAL PLAY INSPECTION) AND ERAC 409 (POST HYDRAULIC SYSTEM 3 HOT
CAUTION
CONDITIONAL INSPECTION) WERE ALSO RELEASED.
C. THE NAMDRP WEBSITE HAS THE CHERRY POINT MATERIALS LAB REPORT AND THE
RESULTS OF THE BELL TUBE ANALYSIS ATTACHED UNDER THIS RCN.
D. TO ASSIST US IN MEASURING OUR PERFORMANCE, PLEASE PROVIDE FEEDBACK ON
THE
QUALITY OF THIS PRODUCT BY ACCESSING
HTTP:WWW.NADEPCP.NAVY.MIL/CUSTSATSURVEY/CUSTSATSURVEY.CFM.
13. PENDING ACTIONS:
A. COMNAVAIRSYSCOM/BELL HELICOPTER: REDESIGN THE EAPS HYDRAULIC SYSTEM
TO
WITHSTAND SYSTEM LEVEL EFFECTS OF A DEGRADED OR FAILING EAPS BLOWER. THE
REDESIGN MAY INCLUDE BUT NOT BE LIMITED TO REROUTING THE HYDRAULIC TUBE
INSTALLATIONS OR CHANGING TUBE MATERIALS AND CONSTRUCTION. THE EAPS
HYDRAULIC SYSTEM MUST BE ROBUST ENOUGH TO WITHSTAND THE SYSTEM LEVEL
EFFECTS
OF EAPS BLOWER FAILURES.
B. COMNAVAIRSYSCOM/BOEING: REDESIGN THE VMS SOFTWARE LEAK DETECTION AND
ISOLATION SYSTEM TO PREVENT COMPLETE LOSS OF HYDRAULIC SYSTEM 3 IN A
FAST
LEAK SCENARIO. RAPID DETECTION AND ISOLATION OF HYDRAULIC SYSTEM LEAKAGE
WILL REDUCE THE AMOUNT OF FLUID THAT MAY BE EXPELLED INTO A POTENTIALLY
FLAMMABLE AREA OF THE AIRCRAFT, SUCH AS THE MIDWING OR THE LOWER
NACELLE.
C. COMNAVAIRSYSCOM/BELL HELICOPTER: REDESIGN THE NACELLE DRAINAGE
SYSTEM
TO ROUTE FLUID LEAKAGE AWAY FROM THE HOT SECTIONS IN THE LOWER NACELLE
AND
IN THE IR SUPPRESSOR. THE NACELLE DRAINAGE SYSTEM SHOULD BE ABLE TO
CONTROL
AND DIRECT A LEVEL OF FLUID THAT WOULD BE EXPECTED TO BE EXPELLED IN THE
LOWER NACELLE AS A RESULT OF A HYDRAULIC SYSTEM FAILURE, FUEL FEED HOSE
FAILURE, OR ENGINE OIL SYSTEM FAILURE.
D. COMNAVAIRSYSCOM/BELL HELICOPTER: REDESIGN THE IR SUPPRESSOR SECTION
TO
ADD FIRE DETECTION AND SUPPRESSION.
E. COMNAVAIRSYSCOM/BELL HELICOPTER: REDESIGN THE EAPS BLOWER FAILURE
DETECTION SYSTEM TO PROVIDE AN EARLY DETECTION OF IMPENDING BLOWER
FAILURES.
THE CURRENT FAILURE DETECTION SYSTEM, CONSISTING OF 500 PSI PRESSURE
SWITCHES ON THE CASE DRAIN TUBES, IS NOT EFFECTIVE SINCE THE CASE DRAIN
FLOW
RESTRICTION ORIFICES WERE REMOVED AS PART OF INTERIM AFC 058/TCTO TCTO
1V-22(C)B-514.
14. NAMDRP WEB SITE HAS A SUPPORTING DOCUMENT ATTACHED. ACCESS WEB SITE
TO
VIEW SUPPORTING DOCUMENTS.
15. THIS IS CONSIDERED CLOSING ACTION ON HMR/EI RCN: V52842-06-0208,
INVESTIGATION CONTROL NUMBER WC2EI-V22-0106-06M.//