The Boeing 767 aircraft was conducting an international passenger flight from Auckland to Melbourne. During cruise the pilot in command (PIC) felt increasingly fatigued, and while outside the flight deck his condition deteriorated. He felt shaky and nauseous and had pain in the back of his head and neck. He was administered oxygen by a member of the cabin crew. The PIC was relieved of duty and the flight continued to the destination with the copilot at the controls. An alert phase was declared. After landing the PIC was taken to hospital for observation. Subsequent tests proved inconclusive, but no evidence was found of a heart-related problem.
The pilot reported a history of stress-related difficulties over several years. He had received treatment for anxiety through a combination of stress management and medication, in the form of a selective serotonin reuptake inhibitor (SSRI). The pilot was also being treated for hypertension. It is possible that the incapacitation of the PIC was related to an anxiety reaction precipitated by a combination of factors including low blood pressure due to hypertension medication, fatigue and a head cold.
The CASA policy of granting medical certification to some private and commercial pilots and air traffic controllers who are taking medication such as SSRIs differs from that of most other Civil Aviation Authorities. However, the approach taken by CASA is in line with that recommended by the Aerospace Medical Association. In 2005, CASA published a safety evaluation of the policy. The report concluded that the policy was appropriate and that there were no safety concerns relating to the practice.
At approximately 1703 Western Standard Time, on 1 August 2005, a Boeing Company 777-200 aircraft, (B777) registered 9M-MRG, was being operated on a scheduled international passenger service from Perth to Kuala Lumpur, Malaysia. The crew reported that, during climb out, they observed a LOW AIRSPEED advisory on the aircraft’s Engine Indication and Crew Alerting System (EICAS), when climbing through flight level (FL) 380. At the same time, the aircraft’s slip/skid indication deflected to the full right position on the Primary Flight Display (PFD). The PFD airspeed display then indicated that the aircraft was approaching the overspeed limit and the stall speed limit simultaneously. The aircraft pitched up and climbed to approximately FL410 and the indicated airspeed decreased from 270 kts to 158 kts. The stall warning and stick shaker devices also activated. The aircraft returned to Perth where an uneventful landing was completed.
The aircraft’s flight data recorder (FDR), cockpit voice recorder and the air data inertial reference unit (ADIRU) were removed for examination. The FDR data indicated that, at the time of the occurrence, unusual acceleration values were recorded in all three planes of movement. The acceleration values were provided by the aircraft’s ADIRU to the aircraft’s primary flight computer, autopilot and other aircraft systems during manual and automatic flight.
Subsequent examination of the ADIRU revealed that one of several accelerometers had failed at the time of the occurrence, and that another accelerometer had failed in June 2001.
Graphical and animated representation of flight data
Various representations of key parameters were prepared from the 9M-MRG downloaded flight data to assist in the analysis.
Graphical representation of relevant recorded data
General parameters over a 60-minute period containing the entire incident flight are displayed, see figure 6. Other relevant parameters are displayed over a 5-minute period incorporating the upset event, see figures 7-10.
Animated representation of relevant recorded data
An animation of the incident was prepared using Insight Animation™ software and is part of this report. A file containing the animation in Insight View™ format (.isv) is available for download from the ATSB website. This file requires the installation of an Insight Viewer that can be downloaded from www.flightscape.com at no charge. A still screen capture of the animation is shown at figure 11.
Download animated representation of flight data [4.4Mb.zip] please see the information above regarding the playing of this file.
On 8 July 2005, the pilot was conducting a charter flight, with two passengers on board, in a Piper PA31-350 Navajo Chieftain. The flight was initially planned to proceed from Essendon Airport to Mount Hotham, Victoria. However, because of adverse weather, the pilot revised his destination to Wangaratta.
While en route, he diverted the aircraft to his originally intended destination, Mount Hotham. The pilot subsequently reported to air traffic control that he was overhead Mount Hotham. He changed the flight category from visual flight rules to instrument flight rules and advised his intention to conduct an instrument approach to runway 29. At about 1725, the pilot told the Mount Hotham Airport Manager by radio that he was on final approach for runway 29 and asked him to switch on the runway lights. After doing so, the manager attempted to tell the pilot that the lights had been switched on but received no response. Subsequent attempts by air traffic control and the crews of other aircraft to contact the pilot were also unsuccessful. Because of hazardous weather conditions over the following two days, the search for the aircraft was primarily conducted on foot and horseback.
The aircraft was located on a tree covered ridge, partially covered by snow. It had flown into trees in a level attitude, slightly banked to the right. Initial impact with the ridge was at about 200 ft below the elevation of the Mount Hotham aerodrome. The Chieftain had broken into several large sections and an intense fire had consumed most of the cabin. The occupants were fatally injured.
The investigation determined that the aircraft systems had been operating normally. The weather conditions were ideal for a 'flat light' phenomenon that was likely to have denied the pilot adequate visual reference. The pilot may have experienced disorientation and loss of situational awareness. The aircraft was not equipped for flight in icing conditions, nor had the pilot complied with the requirements for flight under the instrument flight rules or in accord with the visual flight rules.
The operator issued a Maintenance Alert requiring each 50-hour inspection to include lubrication of the main and nose landing gear down-lock actuating mechanisms in accordance with Part I of Piper Service Letter 755C.
The Australian Transport Safety Bureau (ATSB) provided information about this and the other related occurrences to the Civil Aviation Safety Authority (CASA). As a result, on 6 September 2005 CASA mailed an All Operators Letter to operators of PA-31 and PA-42 series aircraft. That letter informed operators of the occurrence and advised them to lubricate the down-lock assembly in accordance with Part I of Piper Service Letter 755.
On 27 September 2005, CASA issued Airworthiness Bulletin (AWB) 32-005. The AWB, applicable to PA-31 and PA-42 series aircraft, recommended inspection and lubrication of the down-lock latch and pivot bolt in accordance with Part 1 of Piper Service Letter 755.
CASA advised the ATSB that 'As part of its surveillance program, CASA will monitor operators to ensure systems of maintenance include the requirements as stated in the All Operators Letter and the Airworthiness Bulletin'.
Analysis
The right main landing gear down-light did not illuminate because the down-lock hook had not completely engaged and actuated the microswitch. Post-accident extension of the landing gear indicated that stiffness of the down-lock assembly prevented complete engagement. The initial inspection indicated that failure of the right landing gear down-lock to properly engage was the result of inadequate lubrication. This was consistent with the circumstances involved in four Chieftain occurrences in the preceding 12 months that each involved failure of a main landing gear down-lock assembly to properly engage.
Had the main landing gear down-lock assemblies been lubricated in accordance with Part I of Service Letter 755C, it is likely that the right main gear down-lock assembly would have fully engaged. Although the service letter referenced in the operator's 50-hour maintenance schedule indicated that the down-lock assembly should have been lubricated, that task was not specifically stated in the schedule. Incorporation of a specific requirement for down-lock assembly lubrication at 50-hour intervals in Chieftain maintenance schedules will reduce the risk of incomplete down-lock engagement and gear-up landings.
Factual Information
On 1 August 2005, at about 1000 Eastern Standard Time, a Piper Aircraft Corporation PA-31-350 (Chieftain), registered VH-LMB, departed Adelaide on a scheduled passenger flight to Port Augusta. The aircraft was being operated under the instrument flight rules with a pilot and seven passengers, including an aircraft maintenance engineer. On arrival in the circuit area at Port Augusta the pilot selected the landing gear down. The landing gear appeared to operate normally, but the right main landing gear down-light did not illuminate, the gear selector did not return to the neutral position and the transit light remained on.
A number of landing gear retractions and extensions produced the same result. With the assistance of the engineer in the copilot seat, the right gear-down indicator light was changed and the manual extension procedure carried out. However, the landing gear unsafe condition remained and the pilot conducted a low pass to allow an aircraft maintenance engineer on the ground to observe the gear. The landing gear appeared to be down and locked, but the inboard gear doors (flipper doors) remained extended, indicating that the extension cycle was not complete. The pilot attempted to engage the right main landing gear down-lock by manoeuvring the aircraft and repeating the normal and manual gear extensions, but was unsuccessful.
The pilot reported that, consistent with the operator's procedures, he elected to land with the landing gear retracted. The passengers were briefed and the engineer moved to the seat adjacent to the emergency exit. After a total of about 1 hour 40 minutes in the Port Augusta area, the pilot landed on runway 33. The propellers, underbelly skin and flaps were damaged. The occupants were not injured.
Aircraft maintenance engineers who inspected the aircraft after it was lifted, found that when the landing gear was manually extended, the right down-lock assembly was stiff and did not completely engage. A force applied to the down-lock assembly completed its engagement and actuated the right down-light. The operator advised that an engineering report would be completed once the aircraft was recovered and full retraction tests carried out.
The aircraft manufacturer issued Service Letter 755C in November 1985. Part I of the service letter addressed inspection and lubrication of the landing gear lock actuator rod and rod end bearing assemblies. This was recommended at each 50 hours of operation and whenever landing gear and wheel areas were washed. Part I included the following statement:
It has been determined that the location of the exhaust outlets on the … PA-31-350 … aircraft are such that an increased frequency of inspection and lubrication is recommended. It is also recommended that the inspection and lubrication includes downlock latch and pivot bolts.
Although operators were not required to comply with service letters, the operator's Chieftain System of Maintenance included a 50-hour maintenance schedule that specified inspection of the main landing gear down-lock rod and cable assemblies and referred to Piper Service Letter 755C. The 50-hour maintenance schedule did not specifically require lubrication of the landing gear, and the aircraft's main landing gear down-lock had not been lubricated at the recent 50-hour inspection.
A search of the Australian Transport Safety Bureau occurrence database showed that in each year from 1997 to mid 2004, there were up to two occurrences that involved Chieftain main landing gear down-locks. In the 12 months prior to the occurrence, there were four Chieftain main landing gear problems involving stiff or sticky down-lock assemblies. The operators or maintainers of those Chieftains indicated that inadequate lubrication was the main reason for the down-lock stiffness or stickiness.
Summary
On 1 August 2005, at about 1000 Eastern Standard Time, a Piper Aircraft Corporation PA-31-350 (Chieftain), registered VH-LMB, departed Adelaide on a scheduled passenger flight to Port Augusta. The aircraft was being operated under the instrument flight rules with a pilot and seven passengers, including an aircraft maintenance engineer. On arrival in the circuit area at Port Augusta the pilot selected the landing gear down. The landing gear appeared to operate normally, but the right main landing gear down-light did not illuminate, the gear selector did not return to the neutral position and the transit light remained on.
The investigation considered a number of possible factors that could have resulted in the elevator control system difficulties reported by the pilot. Those factors included: mechanical interference in the system, control lock insertion, an excessive nose-down trim setting or nose-down autopilot inputs.
The investigation found no pre-existing defects or mechanical interference within the elevator control system that could have contributed to the pilot encountering resistance in the system while rotating the aircraft during the take-off roll.
The control lock was found in the cockpit stowage pocket after the accident, and it did not display indications of having been inserted in the locking collar during the impact sequence. However, given the nature of the damage to the aircraft, the elevator control system may not have sustained damage and stresses sufficient to mark or deform an inserted control lock.
The elevator trim setting was found to be close to the normal take-off setting of about three rotations of the trim wheel back from the full nose-down position. It is unlikely that the trim setting would have significantly changed during the impact sequence and the investigation concluded that the elevator trim was probably set within the range suitable for take-off.
The pilot could not recall whether the autopilot engagement light was illuminated during the take-off roll. However, if the autopilot had been engaged, the pilot should have been able to override any system inputs and move the elevator controls during rotation.
Accordingly, on the evidence available to the investigation, it was not possible to establish the reason for the significant resistance to the pilot's nose-up inputs.
The pilot's decision to reject the take-off was an appropriate response to the difficulties experienced in rotating the aircraft. The normal accelerate-stop distance chart indicated that if certain conditions were met, the aircraft could have been stopped by the end of the runway. However, the aircraft over ran the runway by 162 m resulting in occupant injuries and aircraft damage.
The rolling take-off with slow application of power extended the actual accelerate-stop distance required, as the chart figure was based on take-off power being set before brakes release. The rejection of the take-off at a speed between 90 and 100 kts would have also extended that accelerate-stop distance, which was predicated on 88 kts.
While the investigation was unable to determine the actual time interval between the attempted rotation and the initiation of the rejected take-off, the rate of power reduction and the amount of braking during the rejected take-off, these factors would have had a significant influence on the accelerate-stop distance.
Pilots operating this type of aircraft can be conditioned in normal operations to gradually reduce power to decrease engine wear and applying minimum braking to reduce tyre and brake wear. While this may be appropriate during normal operations, the pilot of the accident aircraft may have adopted a similar technique during the rejected take-off. The dark conditions may also have contributed to the accident by making it more difficult for the pilot to determine the amount of remaining runway available for stopping the aircraft.
Australian Transport Safety Bureau comment
The risk of runway overruns resulting from rejected take-offs can be reduced by setting maximum power before brakes release, regular practice in performing rejected take-offs and conducting pre-take-off safety briefings that include recalling the actions of the rejected take-off procedure and considering the specific operational requirements of the runway to be used.
This occurrence also highlights the critical importance of pilots checking that the flight controls are capable of full and free operation prior to commencing the take-off roll.
Factual Information
Sequence of events
On 25 July 2005, at about 1835 Eastern Standard Time, a Piper Aircraft Corporation PA-31-350 (Chieftain) aircraft, registered VH-PRJ, overran runway 27 at Nhill aerodrome following a rejected night take-off. The aircraft was being operated on an instrument flight rules charter flight to Charlton, Vic, with the pilot and three passengers on board. The pilot and passengers sustained injuries during the overrun and the aircraft was substantially damaged (Figure 1).
Figure 1: Wreckage of aircraft with runway in the background
Earlier that day, the pilot had flown the aircraft from Essendon to Ouyen and then to Nhill, arriving at about 1500. During the stopover at Nhill, the pilot secured the flight controls by inserting a lock pin in the control column. The pilot stated that, when he started and warmed the engines at about 1700, he removed the lock from the control column and that it was reinserted after the engines were shut down. The passengers arrived at the aerodrome at about 1820 and boarded soon afterwards.
The pilot reported that, during the pre-flight preparation at Nhill, he set the elevator trim to the take-off position by winding the trim wheel to about three rotations back from the full nose-down position. The pilot recalled removing the lock from the control column prior to starting the engines for departure.
The pilot stated that he usually checked the flight controls for full and free movement while backtracking on a runway but he could not recall whether he performed that check when backtracking on runway 27 at Nhill. Part B of the operator's Operations Manual included a series of checklists to be used during operation of the aircraft. The 'Start Up' and 'Pre takeoff' checklists included checking full and free movement of the flight controls.
A witness, located in a house adjacent to the western perimeter of the aerodrome, reported that he observed the aircraft taxi to the eastern end of the runway and commence the take-off roll but lost sight of the aircraft as it moved along the runway.
The pilot stated that the take-off roll was from a rolling start with power being slowly applied until engine turbo-charger output stabilised. At about 90 kts indicated air speed, the pilot attempted to rotate the aircraft but encountered resistance to rearward movement of the control column. He decided to reject the takeoff because the aircraft speed at the time was below his nominated decision speed of 100 kts. The pilot reported that he then reduced the engine power to idle and applied maximum braking.
The aircraft overran the runway, passed through the airport boundary fence, continued across a public road, and passed through another fence before coming to a stop in a paddock about 162 m beyond the end of the runway. An inspection of the runway revealed a skid mark from the aircraft's right tyre, which commenced 65 m before the end of the runway.
After the aircraft came to a stop, the pilot exited the aircraft via the crew door, assisted the passengers to evacuate the aircraft through the rear cabin door and marshalled them to an area away from the aircraft. A passenger then used a mobile phone to contact emergency services, who attended the scene soon after.
The aircraft
A subsequent inspection of the aircraft revealed that there were no pre-existing defects in the elevator control system and elevator trim system or evidence of interference with the elevator surfaces. The elevator trim setting was found to be three and a half rotations of the trim wheel from the full nose-down position.
Between flights, the aircraft elevators and ailerons were locked with a removable control lock (Figure 2). The lock was a 10 cm long pin with a red plastic warning tag which was inserted through the control column and a locking collar that was attached to the instrument panel. There were no external control locks fitted.
The Australian Transport Safety Bureau was advised that the control lock was found in a cockpit stowage pocket after the accident. An inspection of the lock shaft did not reveal any witness marks or deformation caused by the lock being left in the locking collar during the impact sequence.
Figure 2: View of control lock partially removed from control column
The aircraft was fitted with a Bendix Altimatic V FD-1 autopilot system. The autopilot controller panel was located on the centre cockpit pedestal behind the engine controls. Engagement of the autopilot system was accomplished by pressing the AP ENGAGE BUTTON, located on the left side of the controller panel. The button would illuminate when the autopilot was engaged.
The pilot could not recall whether any buttons were illuminated on the autopilot controller prior to, or during, the take-off roll. The system was designed to allow a pilot to momentarily override an autopilot input to the flight controls.
The main landing gear wheel brakes were serviceable and did not exhibit any indications of overheating.
Aircraft performance
The take-off weight and centre of gravity of the aircraft were within limits for the flight.
The approved Aircraft Flight Manual (AFM) contained take-off distance and accelerate-stop distance performance charts for both normal and short field takeoffs. The investigation calculated the take-off performance of the aircraft using the following aerodrome and meteorological information:
runway length 1,000 m with a bitumen surface
runway slope 0.8 per cent down to the west
wind 300 degrees true, 10 kts gusting to 13 kts
no rainfall recorded during the previous three hours.
The normal take-off distance chart indicated that with take-off power being set before brakes release, the prevailing weather conditions and a take-off weight estimated by the investigation to be 2,941 kg, the aircraft would have a take-off distance of about 644 m from brakes release to 50 ft. With application of the 1.24 factor specified in Civil Aviation Order 20.7.4 for charter operations, the take-off distance required was 799 m, which was within the length of runway 27.
The normal accelerate-stop distance charts were based on take-off power being set before brakes release, wing flaps retracted, a paved, level and dry runway surface, and an 'abort' (reject) speed of 88 kts indicated air speed. The chart allowed for a failure recognition time of 3 seconds. If the pilot technique in applying take-off power differed from that stated on the accelerate-stop chart, the distance to perform the accelerate-stop manoeuvre would be more than the chart derived figure. Provided the takeoff was rejected at 88 kts, the chart indicated that, under the conditions prevailing at the time of the accident, the accelerate-stop distance was about 845 m, which was within the length of runway 27.
The aircraft manufacturer's procedure for a rejected takeoff was included in the emergency procedure for an engine failure during a normal takeoff at or below 85 kts. That procedure specified that the engine throttles were to be immediately closed and brakes applied as required to stop straight ahead.
Summary
On 25 July 2005, at about 1835 Eastern Standard Time, a Piper Aircraft Corporation PA-31-350 (Chieftain) aircraft, registered VH-PRJ, overran runway 27 at Nhill aerodrome following a rejected night take-off. The aircraft was being operated on an instrument flight rules charter flight to Charlton, Vic, with the pilot and three passengers on board. The pilot and passengers sustained injuries during the overrun and the aircraft was substantially damaged.
The pilot stated that the take-off roll was from a rolling start with power being slowly applied until engine turbo-charger output stabilised. At about 90 kts indicated air speed, the pilot attempted to rotate the aircraft but encountered resistance to rearward movement of the control column. He decided to reject the takeoff because the aircraft speed at the time was below his nominated decision speed of 100 kts. The pilot reported that he then reduced the engine power to idle and applied maximum braking.
A subsequent inspection of the aircraft revealed that there were no pre-existing defects in the elevator control system and elevator trim system or evidence of interference with the elevator surfaces.
The investigation considered a number of possible factors that could have resulted in the elevator control system difficulties reported by the pilot. Those factors included: mechanical interference in the system, control lock insertion, an excessive nose-down trim setting or nose-down autopilot inputs.
The 747 was being operated on a scheduled passenger flight from Singapore to Frankfurt with four pilots, 14 cabin crew and 378 passengers. At the time of flight planning in Singapore, the terminal aerodrome forecast for Frankfurt indicated CAVOK2 conditions at the estimated time of arrival (ETA), and there was no requirement to plan for an alternate aerodrome. The fuel uplift at Singapore for the flight to Frankfurt was in accordance with the operators Civil Aviation Safety Authority (CASA) of Australia-approved fuel policy. The 747 departed Singapore on 27 July at 1521. The ETA for Frankfurt was 0319 on 28 July.
The 0220 Frankfurt routine aviation weather report obtained by the crew included information that CAVOK conditions existed, with no significant changes expected. The crew of the 747 commenced descent into Frankfurt at about 0300, and the Frankfurt automatic terminal information service provided the crew with information that CAVOK conditions existed. As the 747 approached position GED, which was 35 NM from Frankfurt, the air traffic controller instructed the crew to enter a holding pattern at GED. As the 747 was in the holding pattern, radio transmissions from the crews of other aircraft alerted the crew of the 747 that the weather conditions at Frankfurt had suddenly deteriorated, and that there were thunderstorms and heavy rain showers at the aerodrome. The crew elected to divert the 747 to Munich, where it landed without further incident. A number of other international flights were similarly affected. Because of the unexpected and unscheduled diversion from Frankfurt, the reserve fuel remaining on board the 747 when it arrived at Munich was less than that required by the operators CASA approved fuel policy for the planned flight.
The BFU is the independent German government entity responsible for the investigation of accidents and incidents in Germany. The Australian Accredited Representatives role in the investigation has been to provide the BFU with information about the aircraft and its operation, crew details, and a record of discussion taken during an interview conducted with the pilot in command of the aircraft. The BFU will publish the Final Report related to the investigation of this occurrence.
Footnote: 1 Clause 5.18 - The State of Registry, the State of the Operator, the State of Design and the State of Manufacture shall be entitled to appoint an accredited representative to participate in the investigation. Clause 5.23 Any State which on request provides information, facilities or experts to the State conducting the investigation shall be entitled to appoint an accredited representative to participate in the investigation.
2 CAVOK Visibility 10 km or more; no cloud below 5,000 ft or below the highest minimum sector altitude, whichever is the greater, and no cumulonimbus; no precipitation, thunderstorm, shallow fog, low drifting snow or dust devils.
Summary
On 5 August 2005, the Federal Bureau of Aircraft Accidents Investigation of Germany (BFU) notified the Australian Transport Safety Bureau (ATSB) that it was investigating an air safety occurrence involving an Australian-registered and operated Boeing 747-438 (747) aircraft, registered VH-OJE, which occurred at Frankfurt, Germany on 28 July, 2005. The ATSB appointed an accredited representative to participate in the investigation into the occurrence, in accordance with clause 5.18 and clause 5.231 of Annex 13 to the Convention on International Civil Aviation. To protect the information supplied to the BFU, the ATSB initiated an investigation under the Transport Safety Investigation Act 2003. The report presented below was prepared principally from information supplied to the ATSB. Reported times are referenced to Coordinated Universal Time.
Preliminary investigation was undertaken into a category 4
occurrence involving a Cessna 310R aircraft that had been issued a
clearance to cross runway 06 at Jandakot Airport while a Cessna
Citation 550 aircraft was on final approach to land. The Citation
had been issued a landing clearance. The ATSB has terminated the
investigation based on information from Airservices Australia that
there had been an inadvertent omission by the controller in
following an established check procedure. There was relatively
little safety benefit to be gained from continuing the
investigation compared with other priorities.
Status: Downgraded the occurrence to category 5
and investigation discontinued.
The aircraft manufacturer has informed the Australian Transport Safety Bureau that they have instigated an internal investigation and intend to publish a Service Bulletin to address the problem. The Service Bulletin will contain "on-wing" inspections and inspection and refinish actions during heavy maintenance.
Aircraft operator
The aircraft operator has informed the Australian Transport Safety Bureau that they have introduced a series of additional one-time and repetitive inspections of the landing gear trunnions on their 747 fleet. These additional inspections are:
A one-time detailed visual inspection of the trunnion at the first maintenance opportunity following the failure of the component on this aircraft.
Repetitive detailed visual inspections of the trunnion either before every international flight, or during the daily check for domestic flights.
A one-time inspection at the first appropriate maintenance opportunity of the following:
Detailed visual inspection of the internal surface of the trunnion bore by borescope.
Eddy current inspection of the external surface of the trunnion.
Eddy current inspection of the internal surface of the trunnion bore.
Ultrasonic inspection from the external surface of the trunnion, around the area in which the fracture originated, to measure the wall thickness and determine if it is less than the 0.180 inches allowable minimum.
Completion of these inspections without detection of a fault cancelled the requirement for the repetitive inspection detailed in item 2.
Repetitive inspections at every 1C Check of the following:
Detailed visual inspection of the external surface of the trunnion.
Detailed visual inspection of the internal surface of the trunnion bore by borescope.
Eddy current inspection of the external surface of the trunnion.
Eddy current inspection of the internal surface of the trunnion bore.
Significant factors
The wing landing gear trunnion did not conform to the design specifications. The component's wall thickness in the region of the failure was less than the allowable minimum and the internal bore had not been adequately shot peened.
Several fatigue cracks developed in the inner bore at the bore transition region.
The fatigue cracks were likely present during the last overhaul, but were not detected during the magnetic particle inspection.
The fatigue cracks developed until the loads during the pushback operation exceeded the residual strength of the component, leading to failure of the trunnion.
Analysis
The left-wing landing gear forward trunnion sustained a complete through-section fracture during the pushback at Sydney International Airport as a result of fatigue cracks in the bore of the trunnion. The fatigue cracks originated at an internal bore diameter transition and developed until they intersected to form a single crack.
The development and growth of the fatigue cracks was attributed to three principal factors:
the wall thickness of the trunnion was below the minimum required by the manufacture specifications
the surface had machining marks in the surface at the radius
the inner surface of the bore had been inadequately shot peened.
The effect of the reduction in wall thickness was to increase the working stress in the component. This increase in working stress reduced the number of cycles required to produce and develop fatigue damage.
The radius at the transition in the trunnion bore diameter is a natural stress concentration point when the item is smooth, but the presence of the machining marks on the surface of this radius provided further stress concentration. This stress concentration further reduced the number of cycles required to produce and develop fatigue damage.
The lack of adequate shot peening likely had a two-fold detrimental effect on the trunnion fatigue life. Firstly, as the smooth regions in the bore showed, effective shot peening obliterated the machining marks. Those marks remained in the unpeened areas and thus presented an additional stress concentration. Secondly, the absence of adequate shot peening denied the component the fatigue life improving qualities that shot peening brings.
Because there were no entries in the maintenance documents regarding repairs in the internal bore and the blending of the shot peened and non-shot peened areas, it is likely that the trunnion wall thickness was below the minimum design limit and was inadequately shot peened during original manufacture.
The presence of multiple secondary fatigue cracks in the component, also emanating from the root of machining marks, further verified that the failure was not due to a single material defect. As such, it would be likely to occur in other trunnions, which do not have the machining marks obliterated by the shot peening process.
The varying nature of the corrosion within the fatigue cracks and the demarcation between the various regions suggested that the cracks had existed during several overhaul cycles of the component. During overhaul, the component was subjected to chemicals that had a corrosive effect on the material, but would not be readily flushed away from a tight crack. Therefore, it is likely that the crack was present in the component at the last overhaul. The fracture surface indicates that the crack was approximately 4mm long and 1.6mm deep at the last overhaul in 2001.
The component had undergone the manufacturer required inspections at overhaul and no cracks were detected. The Magnetic Particle Inspection (MPI) method used to check the item for defects such as cracks is sensitive enough to detect a crack much smaller than the one suspected to have existed at the last overhaul. Possible masking of crack indications by the machining marks or a lack of expectation by the operator to find cracks in the region may have contributed to any cracks not being detected by the MPI operator.
The machining marks in the surface of the part can give non-relevant indications4 of cracks. Those spurious indications may mask true indications of cracks. If the operator was not aware that the machining marks should not be present, they would be likely to discount them and pass the component.
The aircraft manufacturer provided standard practices in relation to the inspection method used. These practices were general and were to be used by maintainers in developing their component specific procedures. Neither the overhaul procedure for the trunnion nor the general MPI process specification directed the MPI operator's attention to the radius in the bore diameter transition. Therefore, the expectation for an operator's repair shop to find cracks in that region would be low.
Because the manufacture documentation for the particular component was destroyed in 1994, the investigation could not determine how the trunnion was manufactured and released in a state that did not conform to the manufacture drawings. The overhaul and service instructions for the trunnion did not provide a mechanism by which the non-conformances could be detected.
Non-relevant indications are indications that are defect-like in appearance, but are due to the local geometry and features of the component. Heavy machining marks are one such
Factual Information
Sequence of events
At about 1200 Eastern Standard Time on 30 May 2005, a Boeing Co 747-300 (747), registered JA8184, was being pushed back from its gate at Sydney International Airport for a scheduled passenger flight to Osaka, Japan. During pushback, the ground staff heard a loud cracking noise. The pushback was stopped and an inspection by the ground crew identified a structural failure in the left wing landing gear forward trunnion fork (trunnion), as shown in Figure 1 and 2. After an on-site inspection by the Australian Transport Safety Bureau (ATSB), the aircraft was moved to a hangar for maintenance and the fractured component was removed from the aircraft and sent to the ATSB for a detailed examination.
Figure 1 : Left wing landing gear
Figure 2: Looking up and outboard into wing landing gear well
Examination of fractured trunnion
The trunnion had sustained a complete through-section fracture, located approximately mid-way between the ball-end and the fork-end (Figure 3).
Figure 3 : Fracture location on trunnion
A general inspection of the fractured trunnion revealed a discoloured (orange/brown) region on the fracture surface (Figure 4) in the upper outboard region. The corroded nature of that region, compared with the adjacent bright fracture surfaces, indicated the presence of a pre-existing defect, and that the trunnion had been cracked for a period of time prior to the final failure during the pushback.
Figure 4 : Fracture surface
The fractured component was examined in a metallurgical laboratory under the supervision of the ATSB. Chemical analysis of a sample taken from the trunnion near the fracture showed that the material met the specification for AISI/SAE 4340M alloy steel. Metallographic examination confirmed a fine-grained lightly tempered martensitic microstructure, typical of the 4340M alloy in the hardened and tempered condition. The inner and outer surfaces were also observed to have been finished with a metallic type plating and painted with a surface primer and topcoat.
Hardness measurements taken indicated that the material had an ultimate tensile strength of approximately 275,600 psi (1900 MPa).
Wall thickness measurements taken around the fractured trunnion circumference showed that the minimum local wall thickness of 0.137 inches (3.48 mm) corresponded with the corroded and discoloured area of cracking.
A detailed technical examination of the corroded region identified two transverse fatigue cracks originating at the inner surface of the trunnion bore (noted as C1 and C2 in Figure 5). The cracks had initiated approximately 11mm apart and had joined to form a single crack. This crack continued to grow until the final fracture occurred during the pushback.
Figure 5 : Fatigue crack development
Crack C2 presented well defined fatigue progression marks, including several distinct regions of fatigue and corrosion (Figure 6). The region bounded by the red dotted line was a distinct region of heavily corroded fatigue cracking and was about 4 mm long and 1.6 mm deep.
Figure 6 : Fatigue progression and corrosion marks
Close examination of the origins in the inner wall of the trunnion found that crack C1 had initiated from multiple closely spaced origins at the root of a machining groove, giving the appearance of a longer crack following the machining groove (or mark). Crack C2 had also originated at the root of a machining groove from multiple closely spaced origins, but over a much smaller distance before aligning with the principal stress plan1, resulting in the apparent difference in the planes of the cracks as shown in Figure 7. The plane of the initial cracks in both C1 and C2 were approximately parallel and were aligned with the machining grooves in the inner surface.
Figure 7 : Plane of crack origins
The inner and outer surfaces did not have a consistent surface roughness. Well defined machining marks were observed in the large diameter bore and to a lesser extent on the small bore. However, the outer surface and the taper region in the bore were relatively smooth without defined machining marks. The well-defined machining marks on the large and small diameter bore blended into the smoother surface of the taper region (that is, there was no abrupt change in surface roughness).
Examination of the surface microstructure in the region of the cracks revealed that the smooth regions (outer surface and taper section) had a thin layer of deformed material typical of a cold working process such as shot peening2. The area of surface deformation in the taper region ran out just before the radius (a few millimetres from the cracks). Figure 8 shows the differences in the surface roughness at a microscopic level (scale is 25µm, or 0.025 mm).
Figure 8 : Smooth surface (upper); surface with distinct machining marks (lower)
The sections taken from adjacent to the primary cracks for micrographic examination contained multiple independent fatigue cracks of various sizes. One example is shown in Figure 9. Each of these cracks originated in the root of the machining groove and were associated with shallow intergranular penetrations, which also existed in the roots of the machining grooves (Figure 8).
Figure 9 : Secondary fatigue crack indicated by arrow
Component manufacture
The landing gear trunnion was manufactured to the aircraft manufacturer's specifications by an approved external supplier. Both the supplier and the aircraft manufacturer informed the ATSB that the trunnion was manufactured at some time prior to 1975; however, the original manufacture documentation (including the manufacture plan and conformance records) was destroyed in 1994. The trunnion specifications were supplied by the aircraft manufacturer. Those documents included construction drawings and process specifications.
In the trunnion specifications, the aircraft manufacturer specified the use of 4340M steel, heat treated to an ultimate tensile strength of 275,000 to 300,000 psi. Therefore, the material used in the manufacture of the failed component met the steel alloy and strength requirements of the design.
The minimum allowable wall thickness specified3 for the trunnion at the location of the failure was 0.180 inches (4.57 mm). Therefore, the minimum wall thickness measured at the crack of 0.137 inches (3.48 mm) was 0.043 inches (1.09 mm) thinner than the design allowed.
Component maintenance
The maintenance documents supplied by the aircraft operator indicated that the trunnion had been fitted to five aircraft and had amassed a total of 25,095 landing cycles during its service life. The records also showed that the trunnion had been overhauled by the operator's component repair workshop on four occasions (Table 1). The landing gear assembly had an overhaul interval of 8 years or 12,000 cycles, whichever occurred first. Therefore, the landing gear was not due for overhaul for another 4 years, or 9,509 cycles.
Table 1 : Overhaul history
Overhaul
Date
Total cycles at overhaul
1
October 1979
3,517
2
November 1987
14,069
3
March 1992
16,508
4
October 2001
22,604
The overhaul records showed that on each occasion, the trunnion had undergone repair work. The documents indicated that the repairs were limited to the lugs and were within the repair limits permissible by the aircraft manufacturer. There was no record of any repair work carried out in the bore of the trunnion or on the outer surface in the region of the bore diameter transition.
Comparison of the overhaul instructions maintained by the operator with those supplied by the aircraft manufacturer confirmed that the operator's workshop maintained the correct instructions for the overhaul. The overhaul instructions for the component did not require a dimensional check for wall thickness or a specific check for surface finish (roughness) in the bore.
As one of the first processes in the overhaul, the protective finishes (including metallic plating) were removed from the surface of the trunnion. These finishes are removed using a chemical process, some of which, including water for cleaning, can have a corrosive effect on the trunnion material.
The overhaul procedure for the landing gear components required that the component undergo a magnetic particle inspection (MPI) to detect any defects, including cracks, that may have developed during service. The manufacturer did not provide specific instructions on how to perform the MPI on this particular component but provided a general process that the operator was to use as the basis for a component specific process. Neither the overhaul procedure nor the MPI process directed the MPI technician's attention to the radius in the bore diameter transition as a possible location for cracking. The MPI process used was capable of highlighting cracks of less than 0.5mm.
The overhaul records provided by the operator indicated that an MPI of the component was carried out at each overhaul. The MPI procedure defined by the operator was in accordance with the aircraft manufacturer's recommended procedure. It used the equipment and materials recommended by the manufacturer and specified sufficient examinations to highlight any cracks in the component. The overhaul records indicated that the item had been found satisfactory on each occasion, suggesting that no cracks had been detected.
The component maintenance manual for the repair of high-strength steel landing gear parts directed the operator to obtain advice from the manufacturer if cracks were detected during the MPI. The manufacturer did not have a record of any request for advice relating to the failed component.
The crack was in a location that was not readily viewable during a normal visual ground check. There was no requirement to perform a detailed inspection for cracks in the body of the trunnion between overhauls.
The principal stress plane is a plane that the stress in the part acts perpendicular to. In this case, the principal stress plane was not aligned with the machining marks.
Shot peening is a process where small 'shot' beads are fired against the surface of a component producing a residual compressive surface stress and a thin layer of deformed material. This process has been demonstrated to increase the fatigue life of high-strength steel components.
In the manufacture drawings for the component.
Summary
At about 1200 Eastern Standard Time on 30 May 2005, a Boeing Co 747-300, registered JA8184, was being pushed back from its gate at Sydney International Airport for a scheduled passenger flight to Osaka, Japan. During pushback, the ground staff heard a loud cracking noise. The pushback was stopped and an inspection by the ground crew identified a structural failure in the left-wing landing gear forward trunnion fork.
Examination of the trunnion fork revealed that it had failed due to fatigue cracking that had originated on the inner surface of the trunnion fork bore. It was found that the wall thickness at the crack origin was below the minimum allowed by the design and that the inner surface of the bore did not meet the specifications of the design. These factors contributed to the formation and development of the fatigue crack, which lead to the final failure on pushback.
The trunnion fork had amassed a total of 25,095 landing cycles and had been overhauled by the operator on four occasions. During the overhaul the item was inspected for cracks and on each occasion the item was passed. The inspection procedure was general for the item and did not specifically indicate that the area where the cracking originated required particular attention. The surface finish of the inner surface of the bore may have masked indications of any cracks that may have been present.
As a result of this occurrence, the aircraft manufacturer and the aircraft operator have commenced actions to determine the extent of the problem in the remaining fleet and improvements in the inspection of items during maintenance.
On 25 June 2005, a Bombardier Aerospace Dash 8-315 (Dash 8) aircraft was being operated on a scheduled passenger service from Weipa to Cairns, Qld, while three Aero Commander Div Shrike Commander aircraft (Aero commander) were tracking to Cairns from Cooktown, Qld. The four aircraft were in an arrival sequence of seven aircraft tracking to runway 15. Visual meteorological conditions existed during the period of the aircrafts' arrivals.
Based on the aircrafts' estimated arrival time, the Dash 8 was 3 minutes behind the three Aero commanders. The Aero commanders were maintaining 1,000 ft above mean sea level, and the Dash 8 was on descent from 5,000 ft. The Cairns approach controller estimated that low level headwinds would delay the Aero commanders more than the Dash 8. Consequently, the controller re-ordered the landing sequence to place the Dash 8 ahead of the Aero commanders for landing. The approach controller notified the aerodrome controller of the change in the landing sequence.
At 1705 Eastern Standard Time, when the Dash 8 was on final approach at about 6 nautical miles1 (NM) from the airport, the crew received a traffic alert and collision avoidance system (TCAS) traffic advisory. The crew saw an Aero commander to the left at about 3 NM, and observed on the TCAS display another aircraft in close proximity. Very shortly after, they received a TCAS resolution advisory to climb their aircraft, which they followed. The crew reported that as they were climbing, they saw an Aero commander pass to the left of the Dash 8.
The four aircraft were being operated on instrument flight rules category flights that required separation by air traffic control. Controllers could use a 3 NM radar separation standard, a vertical separation standard of 1,000 ft, visual separation by the aerodrome controller, or require a pilot to sight and follow another aircraft. The approach controller's intention was to separate the aircraft using sight and follow procedures.
As the aircraft approached the airport the aerodrome controller used non-standard coordination phraseology that was misunderstood by the approach controller. The approach controller subsequently instructed the crew of the Dash 8 and the three Aero commanders to transfer to the aerodrome controller's radio frequency. The resultant traffic situation presented to the aerodrome controller was difficult to resolve using visual separation or sight and follow procedures.
The TCAS events occurred as the first Aero commander turned right to track behind the Dash 8 on the final approach to runway 15. The Aero commanders were from the same operator and often used sight and follow procedures between company aircraft on arrival.
An Airservices Australia (Airservices) investigation found that the occurrence was due to the following factors:
the approach controller made an error of judgement in assessing the new traffic sequence
the aerodrome controller used non-standard coordination phraseology that was misinterpreted by the approach controller
the use of sight and follow procedures for the number of aircraft involved increased controller workload
the approach controller's lack of recent familiarisation with tower visual separation procedures.
The Airservices report made six recommendations, and the following action has been taken in regard to the recommendations:
tower and approach/departure controllers were reminded to use standard coordination phraseology to reduce the possibility of ambiguity
a tower and approach/departure area familiarisation program has been implemented that requires participating controllers to complete a project paper
letters of agreement regarding sight and follow procedures have been finalised with two local operators.
A nautical mile is equal to 1852 metres.
Summary
On 25 June 2005, a Bombardier Aerospace Dash 8-315 (Dash 8) aircraft and three Aero Commander Div Shrike Commander aircraft (Aero commander) were in an arrival sequence of seven aircraft to runway 15 at Cairns Airport. At 1705 Eastern Standard Time, when the Dash 8 was on final approach at about 6 nautical miles (NM) from the airport, the crew received a traffic alert and collision avoidance system (TCAS) traffic advisory. The crew saw an Aero commander to the left at about 3 NM, and observed on the TCAS display an aircraft in close proximity. Very shortly after, the crew of the Dash 8 received a TCAS resolution advisory to climb their aircraft, which they followed. The crew reported as they were climbing, they saw a second Aero commander pass to the left of the Dash 8.
An Airservices Australia (Airservices) investigation of the occurrence found that the occurrence was due to the following factors:
the approach controller made an error of judgement in assessing the new traffic sequence
the aerodrome controller used non-standard coordination phraseology that was misinterpreted by the approach controller
the use of sight and follow procedures for the number of aircraft involved increased controller workload
the approach controller's lack of recent familiarisation with tower visual separation procedures.
The Airservices report made four recommendations to Cairns Terminal and two recommendations to Cairns Tower, for action.
On 2 July 2005 at about 0920 Eastern Standard Time, the pilot of a Piper PA-32RT-300 aircraft registered VH-SEQ conducted pre-flight checks for a local scenic flight from Townsville Airport, Qld. The pilot then boarded the five passengers and conducted a safety briefing. The passengers were provided with intercom headsets.
The pilot obtained a clearance from the aerodrome controller (ADC) and commenced the take-off. During the take-off roll the pilot noticed an object go past the windscreen. He rejected the take-off and returned to the apron where he shut down the engine and disembarked the passengers.
The pilot found that the engine cowling bungs1 had not been removed during the pre-flight inspection. He removed pieces of the bung material from the cowl openings. The pilot and passengers again boarded the aircraft and departed for the scenic flight.
While over the city the pilot twice attempted unsuccessfully to contact the ADC by radio for approach instructions. The pilot reported that he then noticed that the ammeter did not indicate a positive battery charge rate. Realising that the electrical system had failed, he contacted the ADC by mobile telephone and obtained approach and landing instructions. A passenger later stated that the headsets had operated intermittently during the flight and stopped functioning prior to the landing.
The pilot selected the aircraft's landing gear down while on the downwind leg of the circuit for runway 07. The gear position lights did not illuminate, and he was unable to confirm that the electro-hydraulic powered landing gear had locked in the down position. The pilot stated that he twice applied the emergency landing gear extension checklist procedure. The landing gear emergency extension checklist required a number of actions to be completed. These included a fish-tailing or yawing manoeuvre to ensure that the main landing gear had locked in the down position.
The pilot then attempted to contact the ADC, but the mobile telephone had failed due to low battery charge. While on final approach the pilot received a green light from the ADC confirming the landing clearance. The pilot stated that he did not advise the passengers of the potential problem with the landing gear.
The pilot reported that he was aware the landing gear could collapse and said he had taken extra care to ease the aircraft gently onto the runway. The right main landing gear collapsed during the landing roll and the aircraft then veered to the right and departed the runway surface. The left main and nose landing gears then collapsed. The pilot and five passengers vacated the aircraft without assistance. There were no reported injuries to either the passengers or the pilot.
The subsequent inspection found that a large part of the leather casing of one engine cowling bung had jammed between the starter ring gear and starter motor and dislodged the alternator drive belt. The pilot reported that the cowling bungs were partially obscured by the 2-bladed propeller and were not fitted with streamers. The colour of the bungs was similar to that of the engine cylinder heads. The passengers reported that the pilot said he had overlooked removing the bungs during the pre-flight inspection as he was distracted at the time.
Due to the damage to the landing gear and the aircraft structure, the investigation was unable to establish detailed reasons for the right main landing gear collapse.
Figure 1: The dislodged alternator belt and fluff from the bungs
ATSB COMMENT
The circumstances of this accident demonstrate the need for thorough pre-flight inspections. Once the pilot became aware that the cowling bungs were damaged, the need for a comprehensive inspection prior to further flight should have been apparent.
The loss of power to the electrical system occurred when the battery discharged after the alternator drive belt was dislodged.
The lack of colour contrast probably contributed to the pilot not seeing the engine cowling bungs during the pre-flight inspection. The common practice of using brightly coloured bung covers with attached streamers would have helped to ensure that the pilot recognised their presence.
Although the pilot assessed that the risk of landing gear collapse was remote, he should have provided a specific emergency briefing to the passengers.
Used to prevent bird entry to the engine bay when the aircraft is not in use.
Summary
On 2 July 2005 at about 0920 Eastern Standard Time, the pilot of a Piper PA-32RT-300 aircraft registered VH-SEQ conducted pre-flight checks for a local scenic flight from Townsville Airport, Qld. The pilot then boarded the five passengers and conducted a safety briefing. The passengers were provided with intercom headsets.
