Announcement

Collapse

Terms of Use Agreement

1. You agree, through your use of these public Forums, not to post any material which is unlawful, libelous, defamatory, obscene, vulgar, sexually orientated, abusive, hateful, harassing, threatening, harmful, invasive of privacy or publicity rights, inflammatory or otherwise objectionable. You also agree not to post any copyrighted material unless the copyright is owned by you. You further agree not to use these public Forums for advertising or other commercial enterprise purposes. Any questions directed to, or concerning the administration of this website, will be sent to admin@popasmoke.com and not posted to the public Forums.

2. All postings express the views of the author, and neither the administrators nor POPASMOKE will be held responsible for the content of any postings submitted by the Members or anyone else. The administrators of these Forums reserve the right to remove, edit, move or close any postings for any reason. Members who make postings on the Forums which are not in accordance with the Terms of Use Agreement, risk having their posting privileges withdrawn.
See more
See less

Major V-22 Redesign Underway

Collapse
X
  • Filter
  • Time
  • Show
Clear All
new posts

  • Major V-22 Redesign Underway

    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:http://WWW.NADEPCP.NAVY.MIL/CUSTSATS...TSATSURVEY.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.//

  • #2
    V-22

    Stan:

    You are correct about the caps...what is with that anyway?

    That is a long read but I cannot see where there is a timeframe mentioned for the fixes...unless I missed it?

    It is more than a passing fancy with the V-22 as our son will be flying in it. I have mixed feelings about the aircraft. There are times I think it is a great aircraft and at other times, I think it is going to be a problem child. I had the opportunity a couple of years ago to climb all over one down in New River. I have to admit it was impressive...would have loved to have taken a spin in one... ...

    Guess if they do get deployed to Iraq, we will see if they can do everything being claimed.

    S/F Gordon

    Comment


    • #3
      V-22 Modification

      So...

      ...does this mean that one AK47 round can take it out?
      /s/ray

      Raymond J. Norton
      1513 Bordeaux Place
      Norfolk, VA 23509-1313

      (757) 623-1644

      Comment


      • #4
        The reason the post is in all caps is because this is official naval message traffic sent across wireless mediums and does not comform to normal windows based fonts. This was not written or formatted to be published for the internet.

        Yes, one round can take it out the engine air particle system (EAPS). Just like one round can take out a utility hyd pump in the H46, or the main gearbox oil cooler in an H53. Any medium sized round, or lighter high velocity round (7.62 or 5.56) at normal velocity will rupture all of these systems on all of these aircraft.

        In a hot LZ, pilots will probably secure the EAPS system. Though it is designed to keep dust and debries out of the engine intake, the EAPS is not required for landing in confined dusty/FOD'ed LZs. The aircraft has been operated for an extended time without using EAPS. Over that time engine degradation was noted over time, but it can be tracked and mitigated.

        The EAPS lines and pump are now being inspected now at regular maintenance cycles. New lines and drains are being designed to keep pressure spikes from causing rupture, and to keep hyd fluid from pooling in the event of rupture. Re-read paragragh 13.a through 13.e to see what they're doing about it.

        Comment


        • #5
          I wish our military would modernize and stop using ALL CAPS.
          Naval Message format is all CAPS. System gives no options when you type the message, automatically goes as caps, so when someone copies it- yup, it's all caps
          Slick

          Comment


          • #6
            V-22

            I'm surprised this problem didn't show up sooner. Does anyone have any idea the number of hours on any of the airframes? Are there any 22's with 200, 300, 1,000 hours on them? In areas like desert testing (elevated temps/lots of dust) how did the engines fair, seals, rotorprops, etc.? What kind of hours did they get out of engines?

            How much redundency is there in the fly by wire systems?

            S/F Gordo

            Comment


            • #7
              Originally posted by Eyedohnoh View Post
              Yes, one round can take it out the engine air particle system (EAPS). Just like one round can take out a utility hyd pump in the H46, or the main gearbox oil cooler in an H53. Any medium sized round, or lighter high velocity round (7.62 or 5.56) at normal velocity will rupture all of these systems on all of these aircraft.

              In a hot LZ, pilots will probably secure the EAPS system. Though it is designed to keep dust and debries out of the engine intake, the EAPS is not required for landing in confined dusty/FOD'ed LZs. The aircraft has been operated for an extended time without using EAPS. Over that time engine degradation was noted over time, but it can be tracked and mitigated.

              The EAPS lines and pump are now being inspected now at regular maintenance cycles. New lines and drains are being designed to keep pressure spikes from causing rupture, and to keep hyd fluid from pooling in the event of rupture. Re-read paragragh 13.a through 13.e to see what they're doing about it.
              Yeah- a round could rupture all of these systems on all of these aircraft. Good thing they can at least attempt a safe autorotation! Oh wait, thats right... V-22's cannot do that.

              http://www.v22forum.com/v22forum/for...sts.asp?TID=12

              Josh

              Comment


              • #8
                Gentlemen!

                I am giving fair warning. This thread has been deleted before and can be again. We all have our own opinions of this aircraft, but, we will not hash out what is wrong or correct about it. The A/C has been, is and will remain in production and has an active squadron flying. We all hope for the best and hope that it will prove itself without any more catastrophes, but we are just spinning our gears and causing a lot of hot air.

                Please consider what you post about this topic, because if it continues the way it is going, I can see the thread once again having the plug pulled.

                Brook Stevenson
                9/'67 - 10/'68

                Comment


                • #9
                  With the moderator's permission, I offer this reply. I fully understand if you choose not to post it. I understand your point of valid discussion and do not want to cause disruption or conflict. What follows is fact and not conjecture.

                  If a round pierces the Hyd 3 system on the V-22, there is no effect on the flying qualities. All actuators are fully functional and still powered by systems one and two. If we shut down the EAPS, or even the entire utility hydraulics system, there is no degredation of flying qualities. The aircraft does not need to auto because it will still be flying; all control surfaces are still functional.

                  Comment

                  Working...
                  X