ATSB response to WA Coroner

On 29 September 2010, the Western Australia Deputy State Coroner, Ms Evelyn Vicker, handed down her findings in the inquest into two deaths arising from a Robinson R22 helicopter accident that occurred on 28 September 2003 near Derby in Western Australia.  The ATSB had previously investigated this accident and published its finding on the ATSB website: ATSB investigation 200304074.  The Coroner fully agreed with the ATSB's findings.

The ATSB's key findings in the helicopter accident were:
1.    The failure of the A166 clutch shaft, involving torsional fatigue cracking, which was due to the inappropriate assembly of the shaft to the A907 yoke.
a.    a non approved jointing compound was used during the assembly; and
b.    the bearing blocks were installed over the painted yoke surface

2.    A loss of main rotor drive which was most likely to have occurred at a combination of height and speed that was insufficient to enable the pilot to conduct a successful auto rotation.

Two recommendations from the Coroner affect the ATSB:

Recommendation 3

CASA [Civil Aviation Safety Authority] seek input from the ATSB as to the reasonableness of mandatory inspection of both yoke and clutch shaft attachments in helicopters operating at low height for evidence of fretting in view of the fact this seems to have been a factor in failure of the A166 component in an R22 in 1992, 2003 and 2005.

ATSB Response:

The ATSB wishes to draw attention to the safety actions on page 10 of the ATSB's accident investigation report.  As a result of an ATSB recommendation on 6 November 2003, CASA issued Airworthiness Directive AD/R22/51 which mandated inspections of the A166 shaft to A907 yoke on all R22 helicopters operating in Australia.

CASA also issued AD/R44/019 on 28 November 2003, mandating the same inspection on those R44 helicopters that had the C166 shaft to C907 yoke disassembled since installation at the factory.

On 7 May 2009 the Airworthiness Directives were cancelled by CASA because the instructions contained in them with respect to mandatory inspections were introduced by Robinson Helicopters into the maintenance manuals for the R22 and R44 models.

The ATSB has been advised by CASA:

"In light of the fact that the relevant maintenance manuals were updated to adequately reflect the maintenance practices required by the ADs, it was considered that the ADs were no longer required and they were subsequently withdrawn.

The ADs were no longer considered necessary because person's performing maintenance on Australian Aircraft are required to do so in accordance with the instructions contained in the applicable approved maintenance data (which includes the manufacturer's maintenance manual) - see r.42V of the Civil Aviation Regulations 1988."

The review of the maintenance manuals for the R22 and R44 helicopters in light of this accident and the findings of the ATSB led CASA to the conclusion that there was a heightened risk of improper maintenance practices being employed in the assembly of the clutch shafts in these types of helicopter. CASA addressed this risk by promulgating the ADs which were later adopted by the Robinson Helicopter Company. CASA considers that this risk is now adequately addressed via the amendments which have now been made to the manufacturer's maintenance manuals for both helicopter types."

The ATSB notes that it is normally CASA that would make an assessment as to the reasonableness of the implementation of a specific recommendation after a safety issue has been identified.  In this instance the ATSB issued the initial recommendation on 6 November 2003.  The ATSB is satisfied with CASA's response that the inspection requirement to address improper maintenance practices is contained in the Robinson Maintenance Manuals and mandated through the application of regulation 42V of the Civil Aviation Regulations 1988.

Recommendation 5

ATSB continue to circulate relevant investigation findings to the industry to remind operators and maintenance engineers manufactures recommendations are made for sound technical reasons.

ATSB Response:

The ATSB wishes to draw attention to section 12AA of the Transport Safety Investigation Act 2003 (TSI Act) which outlines that the ATSB's function is to improve transport safety through means that include:
-    Identifying factors that have contributed to transport safety matters;
-    Identifying factors that might affect transport safety;
-    Communicating those factors to relevant sectors of the transport industry and the public.

Through its investigation and research and analysis activities the ATSB is committed to fostering safety awareness, knowledge and action.  During the course of an ATSB investigation or research project the ATSB works with the relevant sectors of the industry to encourage safety action as safety issues are identified.  The final report is always published on the ATSB's website and hard copies made available as required.  Further, the ATSB regularly issues media releases and alerts to provide notification of the Bureau's activities.

Education materials are also supported and issued by the ATSB.

Cooperation with Coroners

ATSB investigations are conducted with the objective of providing findings that can be used to improve transport safety in the future.  Coronial Inquests are a separate process to the ATSB investigation and they are usually supported by their own investigation and brief of evidence.  However, as Inquests also have the objective of seeking to prevent a death occurring again, the ATSB provides cooperation through the explanation of the ATSB's findings in its report.  The ATSB appreciates the interest of Coroners in working with the ATSB in the interests of improving future safety.

Questions concerning the inquest findings should be directed to the Coroner's Court in Western Australia:
Western Australian Coroner's Court
Level 10
Central Law Courts
501 Hay Street
PERTH WA 6000



 

FINDINGS

The investigation identified the following contributing safety factors.

• The part number A44700-009 roll spoiler actuator, as fitted to TNG, was susceptible to fatigue cracking as a result of deficiencies within the design of the actuator cylinder base
• The actuator ruptured as a result of fatigue cracking initiated due to the design deficiency and the accumulation of a sufficient number of flight cycles
• Rupture of the actuator allowed the loss of fluid contents from the number 2 hydraulic system, and the subsequent failure of that system.

SAFETY ACTION

Aircraft manufacturer

On 18 August 2004, the manufacturer issued In-Service Activity Report (ISAR) 2004-07-2760 advising that it considered the overall risks associated with the failure of the actuator as low, and to raise operator awareness of the actuator failure mechanism.

Transport Canada

Transport Canada, being the certification authority for the Dash 8 series of aircraft, conducted its own investigation of the issues associated with actuator failure. In communication with the ATSB, Transport Canada concurred with the aircraft manufacturer's risk management position and agreed that on the basis of current failure rates, no additional action was warranted.

Australian Transport Safety Bureau

The ATSB reviewed the responses and positions of the aircraft manufacturer, Transport Canada and the aircraft operator in relation to the potential for roll spoiler actuator failure to affect operational safety of the Dash 8. While it supports the risk management approach adopted by the manufacturer and Transport Canada, the ATSB believes that the issue of landing performance limitations imposed by a hydraulic system failure is not adequately addressed by this approach. Therefore, the ATSB issues the following safety advisory notice.

Safety Advisory Notice SAN20050012

The Australian Transport Safety Bureau notifies operators of Dash 8 aircraft fitted with part number A44700-007 and/or A44700-009 roll spoiler actuator components, of the increased risk of the cracking and/or rupture failure of the actuator cylinder section as the actuators accumulate service cycles. Failures have been sustained by actuators with service lives ranging from 27,909 to 41,716 cycles. Cracking or rupture of an actuator cylinder results in the loss of contents and subsequent failure of the associated aircraft hydraulic system. Given that the loss of an aircraft hydraulic system will result in extended landing distance requirements, operators are advised to consider the safety implications of operating affected aircraft to regions where the increased landing distance requirements may not be easily accommodated by available primary or alternate aerodromes.

ANALYSIS

Actuator failure

The investigation determined that the roll spoiler actuator from TNG had failed as a result of the growth of fatigue cracking from within the actuator cylinder. Originating from the internal threads at the end of the cylinder, the cracking propagated under service pressure cycles to a point where it compromised the cylinder wall and subsequently caused the rupture of the housing.

ATSB research found records of 16 actuator failures within the world fleet, of the type and nature sustained by TNG. All such failures had occurred in part number A44700-007 or -009 actuators and all had occurred in units that had accumulated over 27,000 flight cycles. There was no record of failures of the newer part number 1556000-1 actuators.

On the basis of the investigation findings, it was apparent that the design of the A44700-007 and -009 actuators renders the units susceptible to fatigue cracking and failure under service conditions. The nature of design influenced fatigue failures is such that they typically present a range of cycles-to-failure that is characterised by a lower threshold value and an increasing number of failures as cumulative cycles increase. The distribution of known actuator failures reflects these features and as such, it is a reasonable expectation that the number of future failures will increase with the age of the Dash 8 fleet equipped with the A44700 series actuators.

Operational implications

Following the TNG actuator failure and increasing awareness of the actuator deficiencies, the aircraft manufacturer reviewed the issues from a risk management perspective and elected not to pursue any direct course of safety action, on the basis that the probability and risks presented by the events were low. From the information available, it was not apparent whether the aircraft manufacturer, in its decision not to implement any direct safety action, had considered the likelihood of an increasing frequency of actuator failure as the number of component flight cycles build across the fleet. Nor is it known if the manufacturer had considered the performance implications of actuator failure and hydraulic system loss where the aircraft was being operated into areas with limited runway distance availability.

Shortly after reaching a cruising altitude of 12,000 ft (FL120), the flight crew of a de Havilland Canada DHC-8-102 (Dash 8), registered VH-TNG, that was operating a scheduled passenger service from Brisbane to Roma (Queensland), were alerted to the failure and rapid loss of contents from the number-2 hydraulic system. Electing to return the aircraft to Brisbane, the flight crew carried out the relevant quick-reference handbook (QRH) checks, and after some initial difficulty, manually extended the landing gear.

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.