Propeller/rotor malfunction

What happened

Following a number of accidents and serious incidents involving Robinson R22 helicopters where a failure of either one or both rotor drive v-belts has led to the occurrence event, the Australian Transport Safety Bureau (ATSB) initiated a Safety Issues investigation into the broader question of Robinson R22 v-belt operational reliability.

What the ATSB found

There were no systemic safety issues identified as a result of the ATSB investigation. However, drive belt reliability was found to be negatively influenced by a broad range of operational and maintenance-related factors, including:

• high gross or overweight operations
• high or excessive engine power settings (manifold pressures)
• sheave misalignment and/or poor drive system condition
• inadequate or infrequent inspections of the rotor drive system.

What's been done as a result

In July 2011, the ATSB issued safety advisory notice AO-2011-060-SAN-001, reinforcing the need for continued vigilance by operators and maintenance organisations regarding the routine inspection of the R22 drive system.

During the course of this investigation, the Robinson Helicopter Company released an updated ‘Revision-Z’ v-belt. Since that change, R22 industry feedback has indicated an overall improvement in the stability of the drive system and a reduction in failure rates.

Safety message

The Robinson R22 helicopter is the most popular light utility helicopter used in Australia and has a reputation for being an extremely reliable machine. Owners and operators should fully appreciate the nature and effects of the operational stresses placed on the helicopter, particularly if the machine is utilised in a dynamic and demanding manner such as required for cattle mustering operations.

Pilots, operators and maintainers should pay particular attention to the installation and condition of R22 drive belts and other components of the drive system, and should ensure that the manufacturer’s requirements for inspection and maintenance of the drive system are adhered to at all times.

The continued safe flight of an R22 helicopter that has sustained a v-belt failure can be assisted by the pilot’s awareness of the indications of a drive system malfunction, and the appropriate management of the emergency autorotation in accordance with published procedures.

On 4 February 2007, the owner-pilot of a Piper Aircraft Co. PA-30 Twin Comanche aircraft, registered VH-DIC, was conducting a private flight from Gold Coast Airport, Qld. The pilot was the sole occupant. Approximately 11 minutes into the flight, at 1622 Eastern Standard Time, the pilot declared an emergency reporting an engine failure and some 15 seconds later that he was also experiencing problems with the 'left engine'. Approximately 13 minutes after departure, the aircraft impacted the water about 100 m off Kingscliff beach, adjacent to the suburb of Casuarina, NSW. The pilot received fatal injuries.

Two days following the accident, the aircraft wreckage including most of the lower centre fuselage, wings, and both engines and propellers, was recovered and examined. The right propeller was recovered with the blades in the feathered position. The left propeller was recovered with the blades in the normal operating range with bending consistent with power being applied at the time of the impact with the sea.

The investigation determined that most probably the right engine stopped operating, followed by an unexplained power loss of the left engine. The aircraft airspeed then decreased below the minimum controllable airspeed during the emergency landing before power suddenly returned to the left engine causing the aircraft to pitch nose up and bank sharply to the right and impacting the water.

ANALYSIS

The circumstances of the accident were consistent with a loss of tail rotor thrust following the failure of the tail rotor drive shaft as the helicopter entered the hover.

The time at which the damage to the drive shaft occurred was not able to be determined. However, given the absence of rotational damage to the tail rotor blades, it is unlikely that it occurred during the accident flight.

The action of the pilot in increasing engine power when faced with the loss of tail rotor thrust was inappropriate and exacerbated the situation.

FACTUAL INFORMATION

At 1215 Eastern Standard Time¹ on 27 September 2004, a Kawasaki Heavy Industries, 47G3B-KH4² helicopter, registered VH-MTF, was being operated on a tourist flight with one adult and a young boy as passengers. The flight included landing on a 1 m high wooden platform in the Teepookana Forest in north-west Tasmania.

The pilot reported that as he brought the helicopter to a 1 m hover above the platform, the helicopter began to rotate slowly to the right. He unsuccessfully attempted to counter the rotation by applying left tail rotor control input. The pilot then increased engine power in an attempt to regain tail rotor control and to move the helicopter clear of the landing platform. That action had the effect of rapidly increasing the rotation of the helicopter to the right and it began to ascend, reaching about 5 m above ground level. The pilot then lowered the collective control and the helicopter impacted the ground heavily on its right side, several metres from the landing platform. The pilot and adult passenger released their seatbelts and then both assisted the young boy to exit the wreckage. The pilot and passengers received minor injuries.

The pilot described the wind conditions at the time of the accident as a headwind with an approximate strength of 8 kts. That assessment was consistent with the wind data for the Strahan area provided by the Bureau of Meteorology³. The pilot also reported that the main rotor RPM indications were normal and that the helicopter had sufficient power to complete the approach⁴. At the time of the accident, the weight and balance of the helicopter were within prescribed limits. There was no evidence that the helicopter had collided with anything during the approach.

The pilot was appropriately qualified and endorsed to operate the helicopter type and held a valid medical certificate. He was a very experienced agricultural aeroplane pilot and had obtained a commercial pilot (helicopter) licence 14 months before the accident. He had accrued at total of 292 hours in helicopters since that time; 286.4 hours of which had been in the Bell 47 helicopter type. The pilot was experienced with operations into and out of the Teepookana Forest landing platform.

The landing platform was located within a dense forest in an area that was cleared of trees but covered by 1 m high scrub. The trees closest to the clearing had been trimmed to a height of about 5 m to allow a 'fly-in, fly-out' approach. There was no requirement to conduct a vertical approach to the platform.

The helicopter's fuselage structure was deformed by the impact and the tail boom was bent in a downward direction at approximately station 100⁵. There was corresponding bending damage to the tail rotor drive shaft assembly long shaft at the same point. The operator reported that examination of the damaged tail rotor pitch control system revealed that the controls were intact and would have been capable of normal operation. All parts of the helicopter were accounted for by the operator at the accident site.

The two-blade tail rotor assembly, mounted on the right side of the tail boom, was intact and correctly attached to the helicopter. There was no evidence of rotational damage to the leading edges or tips of either blade (Figure 1). During the ground impact one blade had been bent outward at the tip and the other was bent in toward the tail rotor gearbox.

Figure 1:     Tail rotor blade damage

The helicopter's tail rotor drive shaft assembly consisted of a series of two short shafts and one long shaft that were situated on the top of the tail boom assembly. The long shaft was supported in eight hanger bearing assemblies and was secured at its front and rear by drive coupling assemblies. The operator inspected the tail rotor drive system and found that the long shaft assembly tubing was fractured and the pin situated through the front drive coupling assembly was sheared. There was also significant distortion of the corresponding pin in the shaft's rear coupling.

Inspection of the tail rotor drive system, including the drive shaft bearings, tail rotor extension housing and tail rotor gearbox, with the exception of the long drive shaft, revealed nothing that would have prevented normal operation.

ATSB specialist examination of the failed components (Appendix A) attributed the tail rotor drive shaft failure to a significant torsional overload event, leading to a loss of coupling security and the subsequent slippage, frictional heating and shear fracture of the shaft. That examination was unable to determine when the torsional overload occurred or what specific events may have contributed to it.

At the time of the accident, the helicopter had logged 72 flight hours since the issue of the current maintenance release. The last recorded maintenance carried out on the helicopter was a spark plug change on 21 September 2004, 1.0 flight hour prior to the accident. On 31 August 2004, 5.7 flight hours prior to the accident, one tail rotor blade was replaced because of delamination of the leading edge wear strip.

Information received from the operator and from the maintenance organisation indicated that there had been no known tail rotor strike or sudden rotor stoppage since the helicopter was placed on the Australian aircraft register in 1992. The helicopter's prior history was not examined.

The company operations manual contained the published normal and emergency procedures affecting aircraft operations. An appendix to the manual contained the flight check systems and operating procedures specific to each aircraft type operated by the company, with the exception of the Kawasaki-Bell 47G3B-KH4 helicopter. The company did however, make available to pilots a copy of the Civil Aviation Safety Authority approved Kawasaki-Bell 47G3B-KH4 helicopter flight manual.

With reference to tail rotor failures, that flight manual stipulated:

1. Immediately execute an autorotative descent and maintain an airspeed of 34 KIAS at least.
2. Execute a normal autorotative descent and landing.

The flight manual did not contain any specific advice for pilots in response to a tail rotor drive failure when hovering.

Information in the company operations manual regarding pilot response to a tail rotor drive failure in another piston-engine helicopter (Robinson R44) included:

LOSS OF TAIL ROTOR THRUST DURING HOVER

1. Failure is usually indicated by right yaw which cannot be stopped by applying left pedal.
2. Immediately roll throttle off into detent spring and allow aircraft to settle.
3. Raise collective just before touchdown to cushion landing

The generally accepted procedure for pilot actions in the event of a tail rotor failure is to quickly roll off the throttle or snap close the throttle and perform a hovering autorotation⁶,⁷,⁸,⁹ For example:

The likely worst place for loss of tail rotor thrust to happen is in the hover, and the reaction is quite simple - get rid of the engine power and land the helicopter from a hovering engine failure condition. Easy to do on those machines that have throttle(s) on the collective¹⁰.

1. The 24-hour clock is used in this report to describe the local time of day, Eastern Standard Time (EST), as particular events occurred. Eastern Standard Time was Coordinated Universal Time (UTC) + 10 hours.
2. The Kawasaki Heavy Industries, 47G3B-KH4 helicopter is a single pilot/single flight control helicopter manufactured under licence from Bell Helicopters. It is commonly known as the KH4 helicopter and is a derivative of the Bell 47.
3. Given that the pilot positioned the helicopter into wind during the approach and landing, the risk of loss of tail rotor effectiveness (LTE) was negligible.
4. There were no external conditions that would have placed the pilot at risk of overpitching or drooping the main rotor.
5. Positioned 100 inches aft of the datum. The datum was located 2 inches forward of the rotor mast centre-line.
6. Coyle, S. (2003). Cyclic & collective - More art and science of flying helicopters. Mojave, CA: Helobooks, pages 341and 342.
7. Federal Aviation Administration. (2000). Rotorcraft flying handbook (FAA-H-8083-21).  Washington, DC: FAA.
8. Newman, R. (1999). Helicopters will take you anywhere: A manual for helicopter pilots. Mentone, Vic: The Helicopter Book Company.
9. Becker, M. (1997). Mike Becker's helicopter handbook. Noosaville, QLD: Becker Helicopters Australia.
10. The Kawasaki-Bell 47G3B-KH4 helicopter had a throttle of this design.

At 1215 Eastern Standard Time on 27 September 2004, the pilot of a Kawasaki Heavy Industries, 47G3B-KH4 helicopter, registered VH-MTF, was being operated on a tourist flight with two passengers in north-west Tasmania. The pilot reported that as he brought the helicopter to a 1 m hover above the raised landing platform, the helicopter began to rotate slowly to the right. The pilot unsuccessfully attempted to counter the rotation by applying left tail rotor control input. The pilot then increased engine power, however, that action had the effect of rapidly increasing the rotation of the helicopter to the right and the helicopter climbed to about 5 m above the ground. After the pilot lowered the collective control, the helicopter impacted the ground heavily on its right side. The pilot and passengers received minor injuries.

The helicopter's tail rotor drive shaft had failed during the occurrence. ATSB specialist examination of the failed drive shaft, attributed the failure to damage from a significant torsional overload event, leading to the shear fracture of the shaft. The examination was unable to determine when the torsional overload occurred, however, examination of the wreckage indicated that it was likely that it had occurred prior to this accident

Information received from the operator and from the maintenance organisation indicated that there had been no known tail rotor strike or sudden rotor stoppage since the helicopter was placed on the Australian aircraft register in 1992. The helicopter's history prior to that time was not examined.

The action of the pilot in increasing engine power when faced with the loss of tail rotor thrust was also examined.

Analysis

The investigation determined that the bull gear failed as a result of a previously known high cycle fatigue cracking mechanism.

The Civil Aviation Safety Authority Airworthiness Directive (AD) requiring compliance with the manufacturer's Service Bulletin (SB) A72-2087 had been completed on the engine. However, the bull gear failed at less than half of the manufacturer's projected component life.

The diaphragm housing had been extensively damaged following the release of the section of the bull gear rim. That damage had prevented the investigation determining the housing's pre-failure condition and whether its condition had contributed to the failure.

At the time of the failure, the spectrometric oil and filter analysis program (SOAP) analysis had been carried out and assessed in accordance with the manufacturer's procedures. While the filter weight increases noted in the engine on 3 February 2004 and 1 April 2004 were within the manufacturer's 'normal sample' range, the above average filter weight, coupled with the traces of carbon steel may have been an indicator of the impending bull gear failure.

The crew handled the engine failure appropriately in accordance with the operator's procedures. The report of smoke in the cabin of the aircraft during the failure was consistent with the ingestion of engine oil into the compressor assembly immediately following the uncontained failure.

Factual Information

At 1100 Central Standard Time on 16 April 2004, a British Aerospace Plc, J32, Jetstream aircraft registered VH-OAE, with 2 pilots and 19 passengers on board, was on descent, during a scheduled passenger flight from Melbourne, Victoria to Mount Gambier, South Australia. As the aircraft passed through flight level (FL) 140, approximately 37 NM from Mt Gambier, the crew reported hearing a bang from the right engine. Simultaneously, the aircraft yawed to the right and they heard something impact the right side of the fuselage. Some smoke was evident in the cockpit.

A check of the aircraft's engine instruments confirmed a problem with the right engine and the crew shut down the engine and feathered the right propeller in accordance with the operator's quick reference handbook drills. The crew then advised air traffic control of the situation and continued for a landing at Mount Gambier Airport.

An inspection of the aircraft by the operator revealed that there had been an uncontained failure of the propeller reduction gearbox on the right TPE 331-12UHR-702H turboprop engine, serial number P66338C. There was also evidence of impact damage on the right side of the fuselage, below the co-pilot's side window area. That impact had not breached the aircraft's pressure hull.

An examination of the engine, supervised by the Australian Transport Safety Bureau (ATSB), found that a section of the spur gear teeth from the outer rim of the reduction gearbox bull gear, had detached during engine operation (See Figure 1). Spur gear teeth are radial, uniformly spaced around the gear's outer periphery and parallel to the shaft axis¹. The detached section of gear had penetrated the diaphragm housing (intermediate gearbox housing) and the gearbox accessory case, before exiting the engine through the compressor air intake.

Figure 1: Cutaway diagram of TPE 331 reduction gearbox

The hole in the accessory case had allowed engine oil to escape and flow over the engine cowling, with metallic debris and oil entering the engine's compressor intake. Once inside the compressor, the oil was able to enter the aircraft's compressor bleed air system that supplied the aircraft's air conditioning and pressurisation air. There was also significant associated damage to the diaphragm housing, the high speed pinion and compressor/turbine main shaft, with metallisation² observed on the turbine and exhaust sections.

An ATSB Technical Analysis report (Appendix A), on the mode of failure of the bull gear, part number 3108295-1, found that the gear had failed as a result of a mechanism known to the manufacturer. The report indicated that the progressive propagation of high cycle fatigue cracking within the gear web and rim transition region, had caused a section of the gear rim to separate from the gear.

The engine manufacturer had introduced several changes to the bull gear design to address 'reliability and reparability issues' that had occurred in the TPE 331 engine type. Among those were changes in gear relief, gear tooth roots had been ground and shot peened to improve fatigue life, the gear rim inside diameter was shot peened to increase fatigue resistance and a coating was added to the gear web to dampen gear vibrations. The engine manufacturer reported that despite those actions some of the re-worked and coated gears had a higher failure rate than non-reworked gears.

The engine manufacturer also investigated TPE 331 engine diaphragm housings, in which gears had failed, to ascertain if distortion of the housing could cause bull gear to pinion gear misalignment. Several problems were identified with those housings that may have contributed to the gear failures. These included bull gear to pinion gear centreline growth and misalignment, growth between diaphragm to gearbox alignment pins and out-of-round bearing bores.

In October 2001, the engine manufacturer issued Service Bulletin (SB) A72-2087³ in response to 16 in-service bull gear rim separations and 13 high speed pinion torque shaft failures. Four of those failures resulted in gearbox debris being ejected from the engine. One failure resulted in the penetration of the right side of an aircraft's pressure hull by a gear fragment. The bulletin indicated that high tooth loading on the bull gear to high speed pinion mesh, bull gear tooth profile, and distortion of the intermediate gearbox housings, had resulted in abnormal wear and subsequent failure of the assemblies.

Service Bulletin A72-2087 required replacement of the bull gear and high speed pinion with new, zero-time components, at intervals not to exceed 3,600 hours in service. It also required the inspection and the rework/overhaul of some gearbox components such as the diaphragm housing, plus a more stringent periodic inspection of specified gearbox components to ensure an optimum operating environment for the bull gear. At the time of failure, the bull gear assembly in this engine had accrued 1,199.55 hours and 1,523 cycles since installation. The engine had accrued a total of 10,755.7 hours and 12,295 cycles since new.

In Australia, the Civil Aviation Safety Authority (CASA) issued Airworthiness Directive (AD) AD/TPE 331/57 to require compliance with SB A72-2087. That AD became effective on 31 October 2001. Amendment 1 to that AD was issued in January 2002. The AD actions had been incorporated into the occurrence engine at the manufacturer's German maintenance facility on 20 December 2002.

Information received from the engine manufacturer following this occurrence, indicated that there had been three bull gear failures in post SB A72-2087 engines. One of those failures was the subject of a UK Air Accidents Investigation Branch (AAIB) investigation, published in AAIB Bulletin number 7/2005. Information from the AAIB on that failure indicated that the bull gear had failed in a similar manner to the gear in this occurrence.

The engine manufacturer reported that a spectrometric oil and filter analysis program (SOAP)⁴ was used to monitor an engine's in-service condition and to reduce the possibility of a premature mechanical failure. That program monitored the type and quantity of the deposits in the engine oil and oil filters over a specified period. A trending feature within that program could highlight an engine with a rapidly increasing filter 'weight' and indicate that further maintenance action was required. A high filter weight quantity of Carbon Steel in a sample could indicate a problem with the bull gear assembly. In November 2000, the engine manufacturer issued Alert Service Bulletin TPE 331-A79-0034⁵ that changed the SOAP interval periodicity to a fixed 100+/- 20 engine hours to minimise variability. On 25 January 2001, CASA issued AD/TPE 331/55 that required Australian compliance with that Alert SB.

The operator had complied with the engine manufacturer's and CASA's SOAP requirements, forwarding samples to the manufacturer's approved venue for testing. The operator reported that they had become concerned about a SOAP report for the occurrence engine that had been received on 3 February 2004. That report, although still within the manufacturer's 'normal sample' guidelines, had a significantly higher filter weight result than had been previously noted for the engine. When queried, the manufacturer confirmed the results of the sample and indicated that a higher reading may be seen following an engine oil change. The engine oil had been changed 101 engine hours prior to that sample being taken. In the subsequent SOAP sample taken 60 engine hours later, on 24 February 2004, the filter weight had returned to a similar level to that of the pre-3 February 2004 samples. The final sample taken prior to the occurrence, on 1 April 2004, was higher than usual and all of the samples had traces of carbon steel.

1. ASM International, Materials Information Society Handbook, Volume II.

2. Metal pulverised by the compressor becomes molten or burned in the combustion chamber and flows rearward, attaching to the turbine and exhaust assemblies (US Department of the Air Force.(1987). Safety Investigative Techniques (AF Pamphlet 127-1, Volume II). Washington DC: Author).

3. Honeywell Alert Service Bulletin - TPE 331-A72-2087, ENGINE - REDUCTION GEAR AND SHAFT SECTION - Replace Gearshaft (Sun and Bull Gear) Assembly, Part No. 3107037-9/10, 3107122-1, 3107162-1, or 3108222-1, or 3108294-1 with Part No. 3108294-1, issued October 2001 and revised 16 November 2001.

4.Service Information Letter - P331-97 - THE HONEYWELL SPECTROMETRIC OIL AND FILTER ANALYSIS PROGRAM FOR ALL TPE 331 ENGINES EXCEPT -14GR/HR ENGINES; Revision 10, Apr 5/02.

5. Honeywell Alert Service Bulletin - TPE 331-A79-0034 - OIL DISTRIBUTION - DECREASED TIME INTERVAL BETWEEN SPECTROMETRIC OIL (AND FILTER) ANALYSIS PROGRAM (SOAP) SAMPLING; Revision 4, Apr 5/02.

At 1100 Central Standard Time on 16 April 2004, a British Aerospace Plc, J32, Jetstream aircraft registered VH-OAE, with 2 pilots and 19 passengers on board, was on descent, during a scheduled passenger flight from Melbourne, Victoria to Mount Gambier, South Australia. As the aircraft passed through flight level (FL) 140, approximately 37 NM from Mt Gambier, the crew reported hearing a bang from the right engine. Simultaneously, the aircraft yawed to the right and they heard something impact the right side of the fuselage. Some smoke was evident in the cockpit.

The following factors were identified as significant to the development of the incident.

1. The tail rotor control servo unit developed a hydraulic fluid leak, with some of the lost fluid entering the pitch change shaft bearing space.
2. Migration of hydraulic fluid into the bearing diluted the grease, affecting the lubricant efficacy and producing accelerated wear and break-up of the bearing cage.
3. The bearing grease was soluble in the hydraulic fluid.
4. The bearing was allowed to remain in service following the discovery and rectification of the hydraulic leak.
5. The pitch change shaft inboard (servo end) nut was a conventional ( right-hand) thread, allowing it to be loosened and unscrewed by torque from the rotating tail rotor drive shaft.
6. Disconnection of the pitch change shaft from the servo actuator caused control of the tail rotor to be lost.

The investigation found that the loss of tail rotor control reported by the flight crew of VH-BHY occurred as a result of the disconnection of the tail rotor pitch change servo from the control rod. That disconnection was a direct result of a breakdown in the anti-friction properties of the tail rotor pitch change shaft bearing, allowing the rotational torque along the pitch change shaft to overcome the assembly torque and locking assembly of the servo end shaft nut. The rotating shaft subsequently unscrewed the nut, allowing it to drop into the tail structure from where it was recovered.

The tail rotor pitch change shaft bearing failure occurred as a result of the contamination and dilution of the grease lubricant, leading to the internal mechanical breakdown of the bearing cage and the partial seizure of the assembly. Testing showed that the bearing grease was contaminated by hydraulic fluid, which likely released from a leaking tail rotor servo-actuator unit identified and replaced 22 days before the incident. The bearing was inspected at the time of the leak discovery and was found to be satisfactory for further service. At that time, there was no requirement to change the bearing in the event of the leakage of hydraulic fluid into the bearing space.

History of the flight

At approximately 1715 on 29 August 2003, the crew of a Eurocopter AS332L 'Super Puma' helicopter, registration VH-BHY, being operated on an offshore commuter flight from Karratha, Western Australia, reported feeling a sudden airframe jolt, followed by a pitch up, roll, and a left yawing motion. Finding they had lost tail rotor control, the crew stabilised the aircraft using pitch and roll control inputs, before declaring a MAYDAY to air traffic services. After assessing the helicopter's condition and vibration levels, the crew elected to return to Karratha where a run-on landing could be performed. The MAYDAY condition was downgraded to a PAN and, after assessing the helicopter's performance during a precautionary approach, a safe run-on landing was conducted.

The aircraft was carrying a flight crew of two and six passengers who were uninjured.

Damage to the aircraft

Damage to the helicopter was limited to the tail rotor pitch change assembly and the tail boom lower keel fairing, which had pulled out several attachment screws. During the initial post-incident inspection, the operator's ground maintenance personnel found the nut and lock washer disconnected from the servo end of the pitch change rod, allowing the rod to move freely within the servo body. The nut and washer were subsequently found in the tail structure beneath the tail rotor drive shaft. The rod (P/N 332A33-0043-00) had sustained circumferential gouging and scoring around the surfaces adjacent to the inboard side of the pitch change spider bearing (P/N 330A33-9903-20). The bearing itself showed evidence of gross mechanical failure, with break-up of the ball cage and dislodgement of the outboard and inboard seals. The outboard bearing retention nut and lock washer remained in-place and secure (figure 2).

Aircraft information

Manufacturer Aerospatiale (Eurocopter)
Model AS332L 'Super Puma'
Serial Number 2129
Registration VH-BHY
Year of manufacture 1984
Total airframe hours 13,525 (approx, at time of incident)

Tail rotor assembly information

The Super Puma helicopter tail rotor control was effected by a hydraulic servo-actuator that applies control force to the tail rotor blades via a central shaft and spider assembly. A locking nut and lock washer secured the actuator to the shaft, assembled to a nominal 266 - 443 pound-inches ( 30 - 50 Newton-metres) dry torque. At the spider end, the connection was similar, with a nominal dry torque of 115 - 266 pound-inches (13 - 30 Newton-metres). The Super Puma tail rotor turns in a counter-clockwise direction when viewed from the right side of the aircraft. The securing nut on the servo end of the pitch change shaft had a conventional thread, while the nut on the spider end of the rod had a left-hand thread. Figures 3 and 4 illustrate the tail rotor assembly and pitch change shaft location.

Maintenance history

The failed tail rotor bearing was first fitted to VH-BHY in June 2000, as part of a complete replacement tail rotor gearbox (TRG) assembly. The gearbox, including bearing, had 199 hours time since overhaul (TSO) when installed. Replacement of the pitch change bearing is normally carried out during gearbox overhaul, however documentation to confirm that action was not available to the investigation.

In June 2003, maintenance action was carried out on the gearbox in response to elevated lateral vibration levels recorded by the helicopter's integrated health and usage monitoring system (IHUMS). Subsequently, on 7 August 2003, the tail servo was replaced after the discovery of leaked hydraulic fluid inside the boot between the tail rotor hub and the pitch change spider. It was evident that the fluid had travelled from the tail servo, through the tail rotor drive shaft and into the boot, bringing the fluid into close proximity with the inboard end of the tail rotor pitch change shaft bearing. The gearbox and assembly had accrued 1,888 hours TSO at that time. During the weeks following the hydraulic leak, the pitch change shaft bearing was inspected as required by service bulletin SB05-00-29 Rev. 3 and accepted for further service. At the time of failure on 29 August 2003, the TRG and pitch change shaft bearing had operated for 1,959 hours since overhaul.

Bearing failure

The bearing fitted to the tail rotor pitch change assembly on VH-BHY was a single race, fully sealed ball bearing, manufactured by SNFA, France. The bearing carried the following identifying marks:

330A33990320 8020141 SNFA FRANCE V80I24K14

ATSB laboratory examination of the bearing confirmed the mechanical failure and break-up of the bearing cage, allowing the circumferential movement of the balls relative to each other and the resultant development of abnormal race loading and frictional conditions (figure 5). The bearing internal surfaces were dry and in most places covered with an adherent black compound (figure 6) that was sampled for later analysis. There was no evidence of any viscous bearing grease remaining within the bearing confines. All rolling contact surfaces of the bearing showed bruising and particle indentation damage (figure 7), however there was no indication of spalling or other rolling contact fatigue type breakdown. None of the bearing components showed evidence of gross overheating or frictional seizure. The bearing cage showed gross levels of wear and metal loss in areas exposed to contact with the rolling elements (figure 8 ). Several fracture surfaces showed evidence of fatigue cracking. The external surfaces of the bearing outer race showed fretting corrosion and wear to the extent of seating within the pitch change spider assembly ( figure 9). There was no evidence of circumferential scoring or other indications of race rotation within the housing or on the bearing seat. Traces of light oil were found on the bearing seat. The odour and appearance of the oil were typical of hydraulic fluid.

Bearing construction

The ATSB examined the tail rotor pitch change bearings from two other AS332L helicopters maintained by the same operator. Both of those bearings and their integral seals were found to be in serviceable condition and showed none of the characteristic indications of failure presented by the bearing from VH-BHY. The service lives of both examined bearings were comparable to the failed unit from VH-BHY. A sample of grease from one of the serviceable bearings was subject to a solubility test with a small quantity of hydraulic fluid recovered from the tail rotor servo fitted to VH-BHY at the time of the incident. With a small amount of manual agitation, the grease proved miscible within the hydraulic fluid, producing a liquid with a characteristic viscosity not appreciably greater than the original hydraulic fluid. Weighing the bearing before and after cleaning found the unit carrying 1.88 grams of grease, which the aircraft manufacturer indicated was a nominal quantity.

Bearing contaminant analysis

Samples of the remnant lubricant from inside the failed bearing, the uncontaminated grease from a serviceable bearing and the hydraulic fluid from VH-BHY were forwarded to an analytical laboratory to determine whether any trace of the hydraulic fluid could be detected within the material from the failed bearing.

Results from that analysis confirmed the presence of characteristic spectral peaks from the hydraulic fluid to exist within the remnants of the grease from the failed bearing. These peaks did not exist within the sample of uncontaminated grease from the serviceable bearing.