failure analysis report v3 - final

7/31/2019 Failure Analysis Report v3 - Final

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Failure Analysis Assignment

The failure of Boeing Vertol 234 LR, Identification G-BWFC, 4kilometres east of Sumburg, Shetland Isles on the 6th November 1986.

Warren Arthur

200913116

Due 17th March 2011


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Table of Contents

1) Note on Bibliographic Citations...............................................................1

2) Introduction........................................................................................1

2.a) Time, Date and Place of Accident......................................................1

2.b) The aircraft model and manufacturer.................................................1

3) Operating Environment Conditions..........................................................2

3.a) History prior to failure.....................................................................23.b) Meteorological conditions.................................................................2

4) Assembly Affected by Failure..................................................................34.a) Description of the assembly affected.................................................3

4.b) The part that failed.........................................................................44.c) The final result of the failure.............................................................5

5) Contributing Factors Toward Failure.........................................................5

5.a) The maintenance factors..................................................................55.b) Service Update Bulletins from the Manufacturer..................................5

5.c) Implementation of service bulletins...................................................6

6) Failure Analysis....................................................................................6

6.a) The testimony of witnesses..............................................................66.b) Acquisition of specimens..................................................................6

6.c) Testing methods and results.............................................................7

6.d) The parts tested for confirmation......................................................96.e) The material factors........................................................................9

7) Damage not part of the aircraft.............................................................10

7.a) Damage to surrounding area..........................................................107.b) Statistics as a result of the failure...................................................10

8) The consequences from an engineering view..........................................10

8.a) Commentary and Conclusion..........................................................108.b) The lessons learned......................................................................11

8.c) The new measures put in place as a result........................................12

9) Bibliography.......................................................................................12

1) Note on Bibliographic CitationsThe entirety of this report has information from the original accident report, any

other sources found referred back to this authoritative document. It is for thisreason that citations were not used; they would be completely unnecessary.

Henceforth all facts, findings and analysis can be attributed to the [1] document,figures will be referenced as per usual.

Sumburg Boeing Vertol Failure Index Page I


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Illustration Index

Illustration 1: Boeing Vertol 234 LR showing outline of power transmission. [2]...1

Illustration 2: The transmission shaft common to both rotor assemblies.[2]........3Illustration 3: Assembly pictorial of partial shaft assembly. [2]..........................3

Illustration 4: Picture of the actual spiral bevel ring gear subject to failure.[2].....4Illustration 5: Separation due to circumferential cracking. [2]...........................8

Illustration 6: Groove present on mating flange of the failed gear. [2]................8

Illustration 7: Section through the flange groove showing the build up of corrosionproducts and the fatigue originating from V-notches.[2]...................................8

Illustration 8: Detail of another gear showing the same wear pattern groove. [2] 8Illustration 9: Picture showing rapid tearing and fatigue sections. [2]...............11

Sumburg Boeing Vertol Failure Index Page II


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2) Introduction

2.a) Time, Date and Place of Accident

The failure occurred at 11H31 GMT on Thursday 6 th November 1986, approximately 4

kilometers east of Sumburg Airport in the Shetland Isles: 0595330N 0011200W.

2.b) The aircraft model and manufacturer

Boeing Vertol 234 LR of a registration G-BWFC, henceforth referred to as WFC. This

aircraft manufactured in 1981 by Boeing Vertol Company, Philadelphia. WFC hadaccumulated 7690 flight hours, and was last checked 10 hours before failure.

The BV234 is a development of the CH47C Chinook helicopter. The main rotors are intandem configuration, being 18.3 metres in diameter, having 3 blades and aninterference rotation of about 80% when taken along the fuselage axis. The blades

are separated vertically, but due to droop and pitch considerations will mesh at lowspeeds and when given certain input at high speed.

Sumburg Boeing Vertol Failure Page 1

Illustration 1: Boeing Vertol 234 LR showing outline of powertransmission. [2]


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The blades are driven via a combining transmission fed by two right angle gearboxes,

supplied by two engines sitting either side of the tail pylon. The head and tail rotorsare fed by synchronous shafts that maintain the relationship between the rotors,

running fore and aft of the combining transmission. Engines are both the AvcoLyncoming AL5512.

3) Operating Environment Conditions

3.a) History prior to failure

WFC was prepared for flight on the morning of the 6 th, the crew having run theengines and confirmed the fixing of an oil leak from the left engine gearbox that had

been discovered the evening before, a breather pipe needed replacement.

WFC took off with 40 passengers from the Sumburg base at 08H58, the destinationwas the Brent Crude Oil Field in the Sumburg basin. No problems were experienced

on the outward flight and no re-fuelling took place, since the complement wassufficient for the return trip.

The return trip began at 10H22 with a full load of 44 passengers, WFC climbed to 760

metres and cruised along a main helicopter lane before descending to 310 metreswhen allowed by air traffic control, some 64 kilometres from Sumburg Airport. At

11H30 a coastguard Sikorsky S61N helicopter departed Sumburg Airport eastward ona training flight, equipped for Search and Rescue (SAR) operations. The pilot of WFC

radioed to control that WFC was four and a half miles west of Sumburg Airport. No

subsequent communication was received from WFC by control.

All the relevant certifications, training and medical assurances were in order for both

the pilot and co-pilot. The crew had been transferred to the Sumburg Airport on the

3rd of November, and had amassed 15000 hours of helicopter experience betweenthem. The take-off weight was well within the limits specified and the certificate of

airworthyness was valid for another year.

3.b) Meteorological conditions

The surface wind at Sumburg Airport was 290 at a nominal 46 kph, gusting to 65

kph. The Shetland basin was covered by a cold north-westerly airstream, while awarm front lay 800 kilometres to the west and was moving at 64 kph eastward,

toward the airport. Visibility was good at 20km and the lowest cloud was an isolatedcumulus at 460 metres. The WFC encountered precipitation in the form of rain and

solids while traversing the Brent field. The sea was in heavy swell, waves ranging

from 0.6 to 1.2 metres.



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4) Assembly Affected by Failure

4.a) Description of the assembly affected

The forward transmission, powering the forward rotor, is fed by the combiningtransmission via a synchronising shaft. Here a speed reduction from a nominal

6912rpm to 225rpm rotor speed is performed in three gear stages. The first stage is

a spiral bevel pair that reduces speed and angles the drive. The second and thirdstages are planetary gears for power conversion only.

The spiral bevel ring gear, the larger in the set, is bolted via 24 bolts to the sun gear

of the planetary system. The flanges between the two are separated by a steel shimwith anti-fretting coating on each side, used to aid alignment calibrations. Friction

through the shim is the primary torque transfer mechanism, the bolts providing thenecessary normal force component.


Illustration 2: The transmission shaft common

to both rotor assemblies.[2]Illustration 3: Assembly pictorial ofpartial shaft assembly. [2]


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4.c) The final result of the failure

The aircraft was completely destroyed, no economic recovery of any kind was

possible. The cause was de-synchronisation of the fore and aft rotors resulting inblade collision and subsequent reactionary removal of the aft rotor pylon. The loss

rendered the aircraft uncontrollable and the pilot could do nothing to prevent thecrash.

5) Contributing Factors Toward Failure

5.a) The maintenance factors

The forward transmission was supplied as installed on acquisition of the aircraft in

1982, the spiral bevel ring gear conformed to the -5 revision that was then current,

this original part remained until the failure incident. At the scheduled overhaul of2150 hours the transmission was modified to the more current -6 standard according

to manufacturer recommendations. A further re-installation due to service lifereplacement of the forward rotor was needed, only 10 hours after that a further

removal was needed due to metal particles being detected in the oil filter. The cause

was found to be the main oil pump failing, this was repaired according tomanufacturer recommendations and the bolts attaching the bevel ring gear were also

checked pro-actively. Another check of the bolts was done the required timeafterwards, with no loss of torque noted. The failure occurred 668 hours after

modification from the -5 to -6 standards. The monitoring systems were adhered toand gave no cause for alarm, however the oil particulate metrics of total quantity and

rate of increase were close to the limits.

5.b) Service Update Bulletins from the Manufacturer

Bulletin: SB 234-63-1009 (initial issue 29 June 1984). New issue for service, forward

rotor shafts to be inspected at 35 hour intervals or retirement after 1600 hours.

Bulletin: A34-63-1010 (Alert issue initial issue 20 August 1884). Inspection of anaft transmission in the model revealed a nut and washer from the bevel to sun

interface was found in the sump, the other bolts were found to be loose. The servicemanual was amended to include tightening procedures in three cases; the general

acceptability of the joint and torque conditions, decreased service interval in the case

of two loose bolt, and the case where the transmission was to be replaced.

Bulletin: SB 234-63-1014/1015 (initial issue 1 August 1985). Redesign of the

bevel/sun to solve the bolt looseness problem; justified by calculation, service data

extrapolation and a 150 hour test. This was passed by the Federal Aviation Authority(FAA) being the certification body while the Civil Aviation Authority was fully involved

in the process. The FAA imposed service interval regulations on the checking and fullydisassembly conditions of maintenance. This point of revision was defined as standard

-6.



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5.c) Implementation of service bulletins

The conversion of the bevel gear from the -5 to -6 standards was subcontracted, the

result was as required by the manufacturer. The order of manufacture was changedby the contractor to reduce duplication of operations, and the protection required of

the reworked surface was done with a product known as Presto Black rather thanthe specified Instant Gun 44/40. This change of product was approved by the

aircraft operator, the product was later approved by the manufacturer via interaction

with the manufacturer's representative, on condition that the correct applicationprocess was strictly adhered to.

The part was assembled, the worksheets indicated the part was of the -6 standardbut was torqued to the old -5 specifications. The operator however stated that the

correct torque had been applied and the transmission was run-tested and released for

service.

6) Failure Analysis

6.a) The testimony of witnesses

The primary witness in an aircraft crash is the cockpit voice recorder CVR, known asthe black box, even though it is usually a highly visible orange colour. The CVR

recorded the crew noticing an increase in noise, described as a roaring sound,

coming from the flight deck. An explosive retort was captured before theinstrumentation log captured the tail dropping to a near-vertical elevation, the pilot

immediately applied pitch control via the controls but the aircraft did not respond.Later the pilot recalls seeing the sea some 50 metres away as the aircraft dropped,

indicating that the tail-down condition did not remain.

A witness on the ground saw WFC drop from what he estimated was 90-120 metres,the aircraft weaved from side to side and the lower rotor blades flew off and hit the

sea. The splash produced was confirmed by fishing boats and observers at theairport.

The coastguard S61N reported to control of seeing two inflatable life rafts in the sea,

later the wreckage of WFC was seen floating on the sea and a survivor was seenclinging to a large piece of the wreckage. The coastguard immediately came to help,

and observed another survivor clinging to a raft while winching up the first. Many

lifeless bodies were floating around the wreckage, the coastguard did not detect anyother signs of life and thus flew the two rescued to Lerwick Hospital, some 29

kilometres away.

Later searches did not find further survivors, but the dead bodies still floating werebrought to Sumburg Airport.

6.b) Acquisition of specimens

The underwater search for the wreckage began the day after the incident, the WFC

was fitted with beacons that aided recovery efforts. High-tech equipment such as

navigation, positioning, sonar, submersibles and a saturation diving system were



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employed. Despite rough sea conditions the main fuselage was lifted by the 10 th,

including most of the transmission. The aft pylon and transmission were away fromthe wreckage, along the flight path that consisted mainly of small pieces of broken

rotor. Approximately 85%-90% recovery figures were estimated, despite theextensive structural disintegration suffered by the upper parts of the fuselage.

6.c) Testing methods and results

Audio from the flight recorder was analysed in the frequency domain, the whole30min recording had frequencies not normally found in the design. The different

signal was characterised by harmonics of the normal base frequency, the 12th

harmonic was present intermittently up until 62 seconds before the recording ended,

when it became a permanent feature. Just before the recording ended a loud increase

of noise like a bang was present.

Due to the damage sustained by the fuselage it was impossible to rule out high

velocity rotor impact with the fuselage, however a low velocity impact was confirmed

with the cockpit. The passenger seats were heavily damaged, having been overloadedin the rearward direction.

Nearly all the rotors and transmission were recovered. The markings were foundconsistent with a glancing blade impact between the yellow blades and

subsequently a collision between the red blades. Further minor collisions were

present, during abnormal relative positions of the rotors. The synchronising shaftswere found to have been damaged on impact with the sea, as well as the aft shaft

having been displaced from its' normal running position.

Pre-impact failure of the forward transmission was determined by the large metalfragment content of the oil, along with slight damage to the input pinion oil seal.

Severe damage was found on the teeth of the input pinion and bevel gears, theauxiliary oil pump and sun gear thrust plate bolts had been broken. The pinion

bearing retaining studs were sheared, the pinion having moved towards the sun gearshaft.

The spiral bevel ring gear had suffered a radial fracture of its' rim and a

circumferential fracture on the outer radius of its' flange. There was evidence of the

bevel ring gear impacting the side of the transmission case and sun gear bearingsupport. The ring gear circumferential crack followed the entirety of the flange

circumference in contact with the mating shim, it was found to be formed by acoalescence of smaller cracks of an alternating angular and radial nature. A groove

with a rough surface was found on the flange face showing corrosion not due to sea

water, through this the main circumferential crack ran while it wasn't snaking around

bolt holes. Both the radial and circumferential cracks showed evidence of fatiguealong most of their lengths.



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The fatigue cracks that formed the circumferential crack were examinedmicroscopically, they were earlier formed between the bolt holes and not from them.

Penetration of the flange by the fatigue cracking varied from 25%-100% in differentsectors between bolts. In general no fatigue origins were found at the bolt holes, the

circumferential crack had diverted to them at a later stage. The radial fracture

originated from the circumferential fracture, also

being due mainly to metal fatigue. Corrosionactivity while the cracks were progressing was

indicated by staining of the crack surfaces,different to the general corrosion present. Other

indications of a corrosion fatigue mechanism ofthe cracks were V-notches in the base of thegroove from which ghost-cracks emanated.


Illustration 5: Separation due to

circumferential cracking. [2]Illustration 6: Groove present on

mating flange of the failed gear. [2]

Illustration 8: Detail of another gear showing thesame wear pattern groove. [2]

Illustration 7: Section through

the flange groove showing the

build up of corrosion productsand the fatigue originating from

V-notches.[2]


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The surface of the shim had suffered corrosion due to the coating having lost

adherence to the core. There was intermixing of the coating material with corrosiondebris originating from the groove.

6.d) The parts tested for confirmation

After the accident the forward transmission was tested by the manufacturer in three

phases. The first test was a total of 60 hours dry run at 100% with a re-assembly

before a 10 hour 110% torque period, a manufacturer-modified part was used. Thesecond was the same except no re-assembly was performed and used a part re-

manufactured in the same fashion as that used on the WFC. The third test simulatedthe repetitive flight conditions on a run in sea conditions, the nature of the loading

and the cyclic nature of salted moisture spray from the atmosphere. The overnight

rest period was also simulated, the tests ran until 300 hours of flight wereaccumulated.The first test showed the beginnings of a shallow groove wear pattern

on the flange, in the same pattern as from the WFC but of no significant depth. Thesecond showed the same wear pattern, with additional shallow sharp-sided

depressions. The shot-peened layer was not penetrated by the wear. At this point theuse of Presto Black in the WFC was proven inconsequential.

The third test showed wear producing a groove, shallower than the WFC but showing

signs of corrosive attack. Further inspection in the groove showed embryonic open-mouth cracks forming in the same fashion as the gear from the WFC. Inspection of

other -5 and -6 standard gears showed electro-chemical interaction between flange

and shim, a broad wear pattern on the -5 gears and a narrow annulus of wear in the-6 standard parts. Corrosion was common in the groove of all -6 gears, along with

mechanical wear due to debris adhering to the shim and acting abrasively upon slightshifts in the mechanism. A -6 gear was found having the same cracking patterns due

to the same fatigue method, it had been in service longer than the WFC.

6.e) The material factors

The technology at the time of design did not allow the machining of the spiral bevel

gear as one piece with the sun gear shaft, this is not the case in the currentenvironment. Thus the joining flange mechanism was necessitated. A number of

revisions of the joint were made as the requirements of the later models increasedand the manufacturer tried to correct problems found in service. This included

modifications to the shim used, as earlier models suffered stress cracking and fretting

around the bolt holes. The bolt specifications were increased and torquing to a higherforce was instituted to try reduce relative movement between flanges. Later the shim

coating was changed and the flange thickened.

Through a series of changes the specifications of the bolts were changed, firstlocktite was specified and later removed, recommendations on bolt torque were

changed. Finally the bolt diameters were increased and torqued higher to reducemaintenance, this necessitated the original shim design due to increased pressures.



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The oil used in the transmission was hygroscopic to a certain extent, analysis showed

that the possible wet areas in conjunction with chloride ion contamination from themaritime environment would create a corrosive environment under stress conditions.

The corrosion from the moisture alone would increase wear, and with deposition onthe shim this wear would increase steadily as the interface changed character. A wet

environment has been shown to increase fatigue crack growth rates and chance of

formation due surface stress raisers.

The method used to monitor the health of the transmission was based around

monitoring the metal content of the oil and comparing it to previously gained data tocheck for patterns that would indicate wear. One of the limitations of the system isthe inability to detect fatigue and cracking, due mainly to the negligible impact on the

oil this has.

7) Damage not part of the aircraft

7.a) Damage to surrounding area

No notable damage to the environment occurred, a small amount of fuel was lost into

the sea. All wreckage landed in the sea, no other property damage occurred.

7.b) Statistics as a result of the failure

Deaths amounted to 43 passengers and 2 crew. There were serious injuries to 1 crewmember and 1 passenger.

8) The consequences from an engineering view

8.a) Commentary and Conclusion

The aircraft in question was lost as a result of metal fatigue on a vital power

transmission bevel gear. Failure was due to circumferential corrosion fatigue crackingas a result of corrosion and wear on the outer rim of the associated mounting flange.

Corrosion was found to be substantially increased in a marine environment due to the

nature of the oil and chloride ions in the atmosphere. The wear patterns producedamplify the effect of normally acceptable tensile stresses as a result of meshing

forces, corrosion enabling this by removal of the shot-peened layer created inmanufacturing to reduce cracking probabilities. Introduction of residual compressive

stresses is the aim of the shot-peening process. The fatigue resulted in a series of

cracks in a radial and circumferential pattern, the radial crack having severed the

gear rim and been secondary to the primary circumferential crack formed as a resultof tensile cyclic meshing forces. The fatigue eventually overloaded the less affectedareas of the gear by cantilever action resulting in progressive tearing and rapid

failure, this was corroborated by the flight recorder log in which the twelfth harmonic

increased shortly before failure.



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The primary collision of the rotors was due to an opening of the gear along the radial

crack, this increased the effective number of teeth in the gear. The addition of onetooth would have resulted in blade collision in only 1.5 seconds due to different ratios

between the fore and aft rotors. The wear on the case supports this, since the gapwould have increased with the load imposed by the driving pinion and effectively

expanded the gear until it impacted nearby geometry.

8.b) The lessons learned

The -6 gear joint as specified by the manufacturer was consistently unsatisfactory,

this was due to modifications made to the original -5 standard part. The acceptanceof the aircraft was based on the long military service history of the preceding model

and the work done in eliminating the issues with the previous gear specification. The

change of design was made to ease the maintenance burden on the civil operators,and a default technique was applied to the problem without much consideration. No

environment tests were done on the new transmission since the previous model did


Illustration 9: Picture showing rapid tearing and fatigue sections. [2]


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not suffer from issues in this regard. This culminated in a short 150 hour test, of an

aft transmission, which traditionally given more problems under the -5 design. TheFAA accepted this rationale and certified the design in collaboration with the CAA.

The expectation of the previous history of the original transmissions indicated even

wear between fore and aft transmissions. The work done before and experience withno fatigue issues on the flange lead the designers to believe that fatigue in the joint

was only as a result of stresses from the bolt holes. As a result the effect of flangesurface damage was not considered.

This chain of events indicates a failure on the part of the rigour of the certification

process.

8.c) The new measures put in place as a result

The flight recorder frequency analysis increased interest in the field of failure

forecasting based on the vibrational characteristic of a machine. Recommendationswere made to encourage the development of these systems, a realistic set of

requirement were laid out and the great need for development in this field wasrecognised.

9) Bibliography[1] K.V. Kellaway, F.D. King, P.G. McNeill, and P. Hancock, Aircraft Accident Report

2/88, London: 1988.

[2] K.V. Kellaway, F.D. King, P.G. McNeill, and P. Hancock, Aircraft Accident Report2/88 (Appendices), London: 1988.


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