ATSB

British Aerospace Plc BAe 146-200-11, VH-JJT

Summary

While VH-JJT was backtracking on runway 13 at Gove after conducting an instrument approach, VH-NJN became visual on final approach, also for runway 13. The airport safety officer reported that NJN continued the approach until very late final before executing a missed approach.

The crews of both aircraft were in radio contact with each other on the Gove MBZ frequency. After landing, the crew of JJT advised the crew of NJN that JJT was backtracking on runway 13 and would shortly be clear of the runway. The pilot in command of NJN elected to continue the approach, anticipating that JJT would clear the runway in sufficient time to permit NJN to land. When it became apparent that JJT would not be clear of the runway, the pilot in command of NJN discontinued the approach.

Examination of the Flight Recorder data from NJN revealed that the vertical separation between the aircraft during the missed approach was not less than 300 ft.

Occurrence summary

Investigation number	199901622
Occurrence date	16/04/1999
Location	Gove, Aero.
State	Northern Territory
Report release date	05/05/1999
Report status	Final
Investigation type	Occurrence Investigation
Investigation status	Completed
Mode of transport	Aviation
Aviation occurrence category	Missed approach
Occurrence class	Incident
Highest injury level	None

Aircraft details

Manufacturer	British Aerospace
Model	BAe 146
Registration	VH-JJT
Sector	Jet
Operation type	Air Transport High Capacity
Departure point	Darwin, NT
Destination	Gove, NT
Damage	Nil

Aircraft details

Manufacturer	British Aerospace
Model	BAe 146
Registration	VH-NJN
Serial number	Air Transport, Domestic, High Capacity, Passenger, Scheduled
Sector	Jet
Operation type	Air Transport High Capacity
Departure point	Darwin, NT
Destination	Gove, NT
Damage	Nil

Boeing 737-376, VH-TAK

Safety Action

As a result of this occurrence the Bureau of Air Safety Investigation, now the ATSB, issued interim recommendations (IRs) to the manufacturer, regulators and operators on 27 May 1999 and identified the following safety deficiency:

Cracking and failure of main landing gear (MLG) trunnion pin actuator attach lugs of Boeing 737 aircraft prevents retraction of the MLG and may result in damage to the aircraft structure. While the manufacturer's service information regarding MLG corrosion requires a general one-time inspection of the trunnion pin, the accomplishment instructions do not specifically direct inspection to the attach lugs for cracking. The inspection of the trunnion pin is not mandated by regulatory authorities.

IR19990046

The Bureau of Air Safety Investigation recommends that Boeing Commercial Airplane Group alert Boeing 737 operators to this safety deficiency and implement an appropriate inspection program.

Boeing response
Boeing responded that it did not agree that failure of the 737 MLG actuator trunnion pin would result in damage to aircraft structure and, as a result, did not consider this failure mode was a safety issue. A metallurgical examination of the failed trunnion pin by Boeing determined that improper restoration of finishes during trunnion pin overhaul was the likely cause of corrosion that resulted in the trunnion pin failure. Boeing released Maintenance Tip, 737-MT-32-009 R1, in January 2001, to highlight the need to properly restore corrosion prevention compounds when landing gear components are removed and replaced during maintenance.

Boeing advised that investigation of other in-service occurrences of trunnion pin lug fractures concluded that some of these fractures were due to a high preload on the clevis due to bolt clamp up. As a result, Boeing revised the Airplane Maintenance Manual (AMM) 32-32-11 in 1997 to reduce the bolt nut torque when installing the actuator rod end bolt at the clevis in question. This was to reduce the clamp up loads and accompanying stresses that may contribute to the initiation of lug stress corrosion cracking.

Boeing advised that given their evaluation of the consequences of the trunnion pin failure, they considered that the maintenance tip accomplished the action to notify operators to the need for proper maintenance to avoid corrosion in these parts.

Response classification: CLOSED-ACCEPTED

IR19990047

The Bureau of Air Safety Investigation recommends that Boeing Commercial Airplane Group review the effectiveness of Service Bulletin 737-32-1198 revision 2.

Boeing response Boeing responded that it had completed a review of the entire 737 MLG for corrosion problems and did not anticipate releasing further service bulletins or similar type documents relative to the trunnion pin failure. Additionally, Boeing considered that the aforementioned service bulletin was properly categorised and did not need to be upgraded to an "Alert" level bulletin.

Response classification: CLOSED-ACCEPTED

IR19990048

The Bureau of Air Safety Investigation recommends that the Federal Aviation Administration (FAA) note the safety deficiency and recommendations and take appropriate action as considered necessary to ensure the integrity of Boeing 737 main landing gear trunnion pin assemblies.

FAA response
The FAA responded that an inability to retract the landing gear was not considered a safety deficiency since the landing gear was in the down and locked position. The FAA considered that an airworthiness directive to mandate the inspections under the service bulletin was not warranted. The FAA after considering there had been four cases of lug failures, advised that they would recommend that the manufacturer consider adding instructions to inspect for corrosion and cracks in the lugs to the subject service bulletin.

Response classification: CLOSED-ACCEPTED

IR19990049

The Bureau of Air Safety Investigation recommends that the Civil Aviation Safety Authority (CASA) note the safety deficiency and interim recommendations and initiate appropriate action as considered necessary to ensure the integrity of Australian Boeing 737 main landing gear trunnion pin assemblies.

CASA response
CASA responded that Boeing was currently reviewing the entire Boeing 737 landing gear assembly for corrosion problems and was expected to address this, and other faults, via service bulletin or similar document. CASA advised it would continue to monitor the manufacturers response to this problem.

CASA also advised that pending issue of such data by the manufacturer, both Australian operators had developed inspections to check the trunnion pin lugs for corrosion. CASA believed that this action was appropriate under its existing regulatory framework, and the inspections carried out were suitable to prevent further problems. CASA considered that issue of an Airworthiness Directive was unnecessary because of the operators' actions and the impending issue of improved manufacturer's data.

Note: Qantas and Ansett were the only Australian Boeing 737 operators at the time the recommendation was issued

Response classification: CLOSED-ACCEPTED

IR19990050

The Bureau of Air Safety Investigation recommends that Australian operators of Boeing 737 note the above safety deficiency and interim recommendations and take appropriate action as considered necessary to ensure the integrity of Boeing 737 main landing gear trunnion pin assemblies.

The Australian operators of Boeing 737 aircraft conducted a visual inspection of the MLG trunnion pins of their fleet of B737 aircraft shortly after this event. Qantas issued an Engineering Instruction (EI 737-032-0105) on 06 May 1999 to perform ultrasonic inspections for cracks in trunnion pins that had approximately 10,000 cycles since new or overhaul. Ansett raised an Alert Engineering Release (B73-32-10-19) on 08 April 1999 to perform a lug ultrasonic inspection on Boeing 737 MLG trunnion pins at 10,000 flight cycles or 4 years since new or overhauled. Repeat inspection was to be at 600 cycle intervals.

Note: Qantas and Ansett were the only Australian Boeing 737 operators at the time the recommendation was issued.

Response classification: CLOSED-ACCEPTED

Analysis

Trunnion pin examination

Inspection of the trunnion pin revealed that the fracture of the lugs was a result of stress corrosion cracking. This had initiated at the surface of the lug bore, where extensive corrosion pitting was present. The hard chromium plating on the actuator rod end attach bolt had been lost from the bolt surface in areas that were in contact with the bushes.

Stress corrosion cracking of high strength steel components of aircraft MLG occurs when the items are exposed to moisture. Stress corrosion cracking in the trunnion pin lugs occurred because of movement of the bushes installed in the lugs and penetration of moisture into the gap created between the bushes and lugs.

Failure of the trunnion pin occurred during extension of the landing gear; however, the same failure could have occurred during landing gear retraction. Fracture of the MLG trunnion pin attach lugs causes the MLG to fall back to the extended position and is therefore fail-safe. A MLG trunnion pin fracture prevents retraction of the landing gear since it results in disconnection of the MLG actuator. Damage to the surrounding aircraft structure and landing gear assembly may result from a subsequent attempt to recycle the landing gear.

Trunnion pin service and maintenance

The manufacturer issued a Service Bulletin (Boeing SB 737-32-1198 revision 2) in January 1993 for inspection, preventive modification, rework and replacement to improve the corrosion resistance of the MLG trunnion pin and actuator rod end bolt. The manufacturer recommended inspection of the trunnion pin at the next convenient maintenance check with modifications or rework at the next major landing gear overhaul.

The accomplishment instructions for the initial inspection of the MLG trunnion pin vary slightly between SB 737-32-1198 and SB 737-32A1224. Neither bulletin specifically directs inspection to the actuator attach lugs.

The operator completed a rework and overhaul of the right MLG fitted to the aircraft in March 1993. As part of this rework, the operator performed a magnetic-particle inspection of the trunnion pin under SB 737-32A1224 revision 2 (now incorporated in SB 737-32-1198 revision 2), with nil defects found. Some thread corrosion damage was repaired by blending before reinstallation on the aircraft.

The MLG trunnion pin had accumulated 34,316 flight hours, 23,159 flight cycles and 12.8 years in service. The period since last overhaul was 19,316 flight hours, 12,211 flight cycles and 6.1 years. The time remaining on the trunnion pin before overhaul was 3,084 flight hours with 51841 cycles remaining life based on the manufacturer's life limits.

The manufacturer advised that it had received three previous reports of trunnion pin fractures at the actuator attach lugs, none of which resulted in MLG collapse. These trunnion pins had previously been inspected and reworked as necessary under either SB 737-32A1224 revision 2 or SB 737-32-1198 revision 2.

The manufacturer's service bulletins did not appear to be effective in preventing recurrence of corrosion and cracking in these parts. The bulletins did not direct inspection for cracking to the MLG trunnion pin attach lugs unless the trunnion pin was subject to a rework.

Summary

During approach, the crew of the Boeing 737 selected "gear down". A loud "bang" was heard and the "gear safe" green light for the right main landing gear illuminated immediately. The left main and nose landing gear lights illuminated in a time consistent with a normal extension sequence. As the cockpit indications showed that the landing gear was locked down, the approach was continued, and a normal landing was completed.

History of the flight

Inspection of the aircraft revealed that the right main landing gear (MLG) trunnion pin attach lugs had failed. As a result, the right main gear hydraulic actuator disconnected from the MLG assembly, resulting in free-fall of the right MLG during extension to the "down and locked" position. The trunnion pin failure resulted in little damage to the actuator and surrounding structure. No damage to control cables or hydraulic lines occurred.

Diagrams of the B737 landing gear assembly and photographs of the damaged trunnion pin are available on the ATSB website linked to this occurrence number.

Occurrence summary

Investigation number	199901455
Occurrence date	04/04/1999
Location	Melbourne, Aero.
State	Victoria
Report release date	01/05/2001
Report status	Final
Investigation type	Occurrence Investigation
Investigation status	Completed
Mode of transport	Aviation
Occurrence class	Incident
Highest injury level	None

Aircraft details

Manufacturer	The Boeing Company
Model	737
Registration	VH-TAK
Serial number	23485
Sector	Jet
Operation type	Air Transport High Capacity
Departure point	Perth, WA
Destination	Melbourne, VIC
Damage	Minor

Boeing 747-438, VH-OJC

Safety Action

Local Safety Action

Airservices Australia Sydney District is investigating whether there is a need to roster staff for the flow position prior to 0600, to ensure that aircraft arriving at the end of the curfew are sequenced appropriately. This task is scheduled for completion by October 1999. In the interim, the evening shift centre coordinator controller will monitor the forecast weather and pending traffic numbers for the following morning, and provide staff for flow duties if warranted

Summary

The Sydney approach radar controller was operating a combined departures/approach service during the early morning shift when staffing was minimal, and had been on duty since 0245.

For noise abatement reasons, runway 34 was the preferred runway for arrivals prior to 0600, but was not utilised due to an excessive downwind component. As a result, a number of inbound aircraft were required to hold, in order to land on runway 16 after 0600. The approach controller was required to nominate to the adjacent sector controllers the minimum longitudinal spacing required between successive arriving aircraft. Local procedures recommended a 15 NM spacing. The approach controller requested and was provided with 10 NM longitudinal spacing between aircraft, including a Boeing 747 (B747) approaching from the south-west, which was sequenced to land ahead of a Boeing 767 (B767) arriving from the north. The controller was also managing a number of other arriving aircraft.

Independent visual approaches (IVAs) to runways 16L and 16R were in use. The approach controller subsequently amended the initial arrival sequence when it became apparent that the B767 would arrive earlier than the B747. This placed those aircraft as number two (runway 16L) and three (runway 16R) respectively in the arrival sequence.

The crew of the B767 were vectored to intercept the runway 16L localiser at approximately 30 NM, and instructed to report when they had that runway in sight. The B747 crew had been instructed to turn right onto a heading of 120 degrees in order to intercept the runway 16R localiser. They were subsequently cleared to make a visual approach after reporting that they had runway 16R in sight. However, as the B747 turned onto final, the aircraft drifted to the left through the centreline of the 16R approach path, triggering a resolution advisory, from its traffic alert and collision avoidance system (TCAS), for the crew to descend. The approach controller observed the close proximity of the aircraft and issued instructions to both crews to turn their respective aircraft from final using a "breakout" procedure. The B767 was sighted by the B747 crew as their aircraft passed through the final approach path. The lateral distance between the aircraft was reduced to 0.3 NM at a time when the vertical separation was 500 ft.

Normally, IVAs are conducted by a director controller using specific procedures, which included the use of a 20 NM scale on the radar display, and a map for intercept guidance. In this instance the approach controller had his display set to a scale greater than 20 NM, and did not use the IVA map. Controllers are also required to provide a radar vector not exceeding 30 degrees for intercept of the localiser. The vector issued to the B747 crew provided a 36-degree intercept of the localiser. Moreover, crews subject to IVA procedures are responsible for a number of actions detailed in the Aeronautical Information Publication (ENR 1.1 - 48, paragraph 36.3.1) including, "ensuring that the runway centreline is not crossed during intercept".

Just prior to the occurrence, two controllers arrived to commence the morning shift from 0600. One of these would have normally replaced the overnight controller; however, due to the number of arriving aircraft, these controllers were instructed to staff the director and flow control positions. A third controller arrived and was waiting to take over from the approach controller when the incident occurred. However, the approach controller's workload prevented him from handing over responsibility for the position at that time.

The approach controller limited his options by coordinating the provision of a 10 NM spacing between successive arriving aircraft. As a result, when the crew of the B747 allowed their aircraft to pass through the extended centreline there was limited margin for error, resulting in an immediate traffic confliction. The performance of the controller was probably degraded by the effects of fatigue and workload. The performance of the B747 crew was likely to have been affected by similar factors.

The provision of additional staff prior to 0600 to provide either a flow or director controller to assist the approach controller, or to relieve the approach controller earlier, would have reduced the approach controller's workload during a critical period.

Occurrence summary

Investigation number	199901401
Occurrence date	07/04/1999
Location	22 km NNW Sydney, (VOR)
State	New South Wales
Report release date	26/07/1999
Report status	Final
Investigation type	Occurrence Investigation
Investigation status	Completed
Mode of transport	Aviation
Aviation occurrence category	Loss of separation
Occurrence class	Incident
Highest injury level	None

Aircraft details

Manufacturer	The Boeing Company
Model	747
Registration	VH-OJC
Serial number	24406
Sector	Jet
Operation type	Air Transport High Capacity
Departure point	SINGAPORE
Destination	Sydney, NSW
Damage	Nil

Aircraft details

Manufacturer	The Boeing Company
Model	767
Registration	VH-RML
Serial number	22980
Sector	Jet
Operation type	Air Transport High Capacity
Departure point	Brisbane, QLD
Destination	Sydney, NSW
Damage	Nil

Amateur Built Aircraft Lancair 235, VH-LWA

Summary

The pilot of the Lancair had built the aircraft as an owner-builder. He and his passenger had intended to fly from Perth, Western Australia to an airshow at Mangalore, Victoria. On the day prior to the accident, the aircraft arrived early in the afternoon at Aldinga, South Australia, where it was refuelled in preparation for the next leg of the flight. The pilot and passenger then stayed overnight with a friend. The following morning, the aircraft departed for Murray Bridge, South Australia in order to meet up with another aircraft for the remaining flight to Mangalore.

When the aircraft was 3 km to the north-east of Aldinga aerodrome, witnesses heard the engine surge and lose power. The aircraft was then seen to enter a spin and crash into a dry creek bed. Both occupants were fatally injured. The accident was not considered to be survivable.

Traces of aviation fuel were found on the ground at the accident site under the wing fuel tanks. There was no evidence that a significant quantity of fuel had been released during the impact. There had been no post-impact fire.

The investigation established that the aircraft had been refuelled the previous day at Aldinga to a capacity of approximately 80L. It was operating within weight and balance limitations, close to its maximum weight, and close to the aft limit of its centre of gravity. There was no evidence to suggest that the aircraft was not airworthy prior to the accident, nor was there any indication that either the pilot or the passenger had been incapacitated immediately before the accident.

The aircraft utilised three fuel tanks: one was located in each wing below the level of the engine, while the third fuel tank was located in the fuselage, above the level of the engine. The engine could be fed with fuel directly from any one of these.

It was the pilot's normal policy to use fuel from the fuselage tank when priming the carburettor prior to starting the engine, and then to select a wing fuel tank once the engine had been started. During the flight from Western Australia, only the wing fuel tanks had been filled at all refuelling ports except for Ceduna, where it was not possible to ascertain how the aircraft had been refuelled. Investigation revealed that at the time of the accident, the low-fuel warning light for the fuselage tank was illuminated, indicating that only a small quantity of fuel remained in that tank.

The investigation found that aviation fuel had been spilled on the ground, chemically burning the grass, at the aircraft's overnight parking location at Aldinga aerodrome. The shape of the burnt grass area was consistent with fuel having been spilled over the sides of a 20 L fuel drum. Police reported that fuel had been stolen on other occasions from aircraft at Aldinga aerodrome. A road near the aerodrome was regularly used for car racing, and it is possible that fuel was siphoned overnight from the aircraft's fuel tanks.

Damage to the aircraft systems precluded a determination of the fuel tank that was selected at the time of the accident. A lack of both fire damage and evidence of fuel spillage at the accident site indicated that there was only a small quantity of fuel in the aircraft at the time of the accident, despite the aircraft having been refuelled the previous day. It was not possible to ascertain if the engine lost power due to fuel exhaustion from a wing fuel tank, or if the fuselage tank had been selected prior to take-off and that that tank had become exhausted.

The reason for the aircraft entering a spin after the engine lost power could not be determined.

Occurrence summary

Investigation number	199901340
Occurrence date	02/04/1999
Location	3 km NE Aldinga Aero.
State	South Australia
Report release date	21/12/1999
Report status	Final
Investigation type	Occurrence Investigation
Investigation status	Completed
Mode of transport	Aviation
Aviation occurrence category	Fuel exhaustion
Occurrence class	Accident
Highest injury level	Fatal

Aircraft details

Manufacturer	Amateur Built Aircraft
Model	Lancair
Registration	VH-LWA
Serial number	W139
Sector	Piston
Operation type	Private
Departure point	Aldinga, SA
Destination	Murray Bridge, SA
Damage	Destroyed

Piper PA-36-375, VH-HSQ

Safety Action

The pilot did not ensure that the road was clear prior to commencing the take-off.
The pilot did not observe a vehicle travelling along the road.

Summary

The pilot of a Piper Pawnee Brave had undertaken to spray an extensive area of land with a herbicide. He commenced operations at around daybreak, departing from his base with a load of chemicals. After that trip, and for the rest of the day, he operated from an airstrip on a property closer to the spray area.

The strip was a private access road that had been upgraded for a length of 800 m to serve as an airstrip. It ran perpendicular to a north-south sealed public road. Adjacent to the sealed road, and running parallel to it, was a powerline about 5 m high. The western end of the airstrip was about 50 m from the powerline. There were no signs on the road to warn motorists that low flying aircraft may be operating from the strip and crossing the road at low level. On each flight the pilot was landing the aircraft towards the east and taking off towards the west.

Because of its weight, the aircraft was flown under the powerline during most take-offs. Prior to commencing each take-off, the pilot checked for traffic on the road. From the cockpit he had a view of the road for about one kilometre in each direction. The pilot commented that he had to wait for traffic on a number of occasions during the day. He reported that because of fatigue and the low sun angle during the accident flight, he did not notice a vehicle travelling south. Just after becoming airborne the pilot saw the car but was unable to take avoiding action.

The left main wheel of the aircraft collided with the front left corner of the car's cabin and ran across its roof before breaking off and coming to rest in an adjacent paddock. The pilot considered that the aircraft was operating normally so he continued with the intended spraying operation and then returned to his home base for a landing. He provided no explanation as to why he did not report the accident immediately.

The driver of the car stated that she saw the aircraft just prior to the impact but was unable to take avoiding action. A passenger in the front seat received lacerations from the broken windscreen.

Occurrence summary

Investigation number	199901299
Occurrence date	28/03/1999
Location	28 km W Pittsworth
State	Queensland
Report release date	19/01/2000
Report status	Final
Investigation type	Occurrence Investigation
Investigation status	Completed
Mode of transport	Aviation
Aviation occurrence category	Collision with terrain
Occurrence class	Accident
Highest injury level	Serious

Aircraft details

Manufacturer	Piper Aircraft Corp
Model	PA-36
Registration	VH-HSQ
Serial number	36-8202019
Sector	Piston
Operation type	Aerial Work
Departure point	28 km W Pittsworth, QLD
Destination	28 km W Pittsworth, QLD
Damage	Substantial

Boeing 767-238, VH-EAN

Safety Action

As a result of the fan blade fracture, the engine manufacturer and operator took safety action to address shroud lockup, leading edge erosion and blade foreign object damage.

Shroud lubrication

The manufacturer issued service information on fan blade shroud lubrication and recommended that lubrication of the fan blade shroud hardface surfaces be performed at every "A" check. The manufacturer revised the JT9D-7R4 Engine Maintenance Manuals to incorporate the lubrication procedure. No on-wing shroud lubrication had previously been required.
The operator incorporated the fan blade shroud lubrication procedure into their Maintenance Manual. Lubrication of the fan blade shrouds is to be carried out after a compressor wash, both on-wing and in the test cell, and during workshop assembly.

Fan blade leading edge erosion

The manufacturer recommended that the operator maintain proper leading edge contours on all fan blades as per the Engine Manual, and that leading edge restoration be performed between 2,000-3,000 cycles, with a 5,000 cycle maximum time between refurbishment.
The operator commenced a program to recontour the fan blade leading edges as per the Engine Manual every 2,500 cycles. Procedures were put in place to monitor and track the time and removal of fan blade sets reaching the 2,500 cycle threshold. A set of float fan blades and tooling was ordered to support the program.

Foreign object damage inspection

The manufacturer recommended that the operator inspect fan blades for foreign object damage in accordance with the maintenance manual.
The operator advised that a routine visual inspection for foreign object damage as per the maintenance manual was already in place and was conducted every 200 cycles. An eddy current inspection of the fan blade leading edge is performed every 350 airframe hours.
Immediately following the fan blade failure, the operator conducted a close visual and eddy current inspection of all Boeing 767 engine fan blades. A number of damaged fan blades were found and corrective action taken.

Analysis

The fractured fan blade and several liberated portions of the blade were examined by the ATSB and by the engine manufacturer. The other 39 fan blades were returned to the engine manufacturer for review.

The blade had fractured about 470 mm above the blade platform, just inboard of the mid-span shroud. About one-quarter of the blade had been liberated. The fan blade had fractured as a result of fatigue crack growth. The failure had originated at a foreign object damage impact site 2.54 mm aft of the blade leading edge on the rear (concave) face of the blade. Traces of mineral debris were detected at the crack origin, indicating that the foreign object damage was the result of stone ingestion. Fatigue crack growth, from a crack depth of 1.5 mm, probably occurred over about 35 flight cycles. The blade had no material abnormalities at the fracture site. No evidence of a birdstrike was found.

The fractured blade had an eroded leading edge in the mid-span shroud region of the fan blade. The engine manufacturer reported that all of the fan blades had eroded leading edges in front of the mid-span shroud (about 152 mm above and below the shroud), and that the shroud hardfaces showed signs of possible lockup. It could not be determined whether the shroud lockup had occurred before or after the fan blade fracture.

The cracking and fracture of the fan blade appeared to be the result of high stress at a foreign object damage impact location in combination with vibratory stress and possible locked shrouds. The manufacturer reported that an adverse combination of these items could produce stresses high enough to fracture a fan blade. The eroded leading edge evident on all fan blades would have contributed to the blade fracture by affecting the vibratory characteristics of the fan blades.

The engine manufacturer said there had been only one other reported fracture of this type of fan blade. This was also due to fatigue crack growth starting at a foreign object damage impact site at a similar radial location to this fracture.

Findings

A fan blade of the right engine fractured as a result of fatigue crack growth during aircraft climb, resulting in shutdown of the engine.
The crack originated at a foreign object damage impact site on the rear face of the blade near the mid-span shroud.
All fan blades of the right engine had significant leading edge erosion/blunting in the mid-span shroud region.
The fan blade shroud hardfaces showed signs of shroud lockup.
High stress at the foreign object damage impact site, vibratory stresses due to blade leading edge erosion and possible locked shrouds caused the crack to grow.
Liberation of one-quarter of the fan blade and the resulting fan imbalance damaged the fan case, nose cowl and other fan blades.

Factual Information

History of the flight

Approximately 21 minutes after departure while the Boeing 767 was passing through flight level (FL) 285 on climb to FL310, a loud bang sound was heard from the right engine. Cabin crew and passengers reported a flash and sparks from the rear of the engine. The pilot reported an immediate drop in the right engine pressure ratio (EPR) and a rise in the exhaust gas temperature (EGT). A moderate vibration was felt through the airframe.

The crew actioned the engine surge/stall checklist but as the EGT continued to increase, the engine severe damage/separation checklist was actioned and the right engine was shut down. The engine fire bottles were not discharged.

The crew made a PAN call stating that the right engine had failed and requested a descent to FL240 and a return to Sydney. The pilot reduced speed to 240 kts in an attempt to reduce the vibration. However, the vibration reduced only during the final approach.

Engine inspection

An inspection of the right engine revealed that about one-quarter of the no. 28 fan blade had broken away, resulting in substantial damage to the inside of the nose cowl and to the majority of the fan blades. Abnormal displacement of fan blades (shingling) was evident on the mid-span shrouds of a number of blades surrounding the fractured blade.

The fan blade attrition lining was damaged around its entire circumference as a result of heavy fan blade rub. The forward fan case was distorted with five nose cowl attachment points damaged.

The nose cowl internal acoustic liners were damaged through to the outer skin in several locations with two punctures of the nose cowl outer skin. The nose cowl had shifted forward, creating a gap between the nose cowl and the forward fan case flange. Fan blade debris and the mid-span shroud root section of the fractured blade were found embedded in the nose cowl following removal of cowl access panels.

Fan blade maintenance

Operator maintenance records indicated that the right engine fan blades were inspected for leading edge cracks on 13 March 1999 (93 hours and 54 cycles prior to the incident). No indications of cracks were detected at that time. Routine inspections of the fan blade leading edges for cracks had been carried out on a monthly basis. A vibration survey test was carried out on 2 February 1999 after removal and inspection of two fan blades.

In the 9 months before the incident, 13 fan blades on the right engine were found to have incurred foreign object damage. These blades had been blend repaired and inspected for cracks. The fractured fan blade was not one of those blades.

The operator reported that fan blade leading edge restoration was carried out at a 5,000 cycle interval or at 3,000 cycles should an overhaul opportunity occur. Lubrication of the mid-span shroud to help prevent shroud lockup was not required while the engine was fitted to an aircraft (on-wing). The fan blades were overhauled 6,689 hours / 3,889 cycles before this incident.

Occurrence summary

Investigation number	199901215
Occurrence date	22/03/1999
Location	56 km E Cowra, (NDB)
State	New South Wales
Report release date	22/11/2000
Report status	Final
Investigation type	Occurrence Investigation
Investigation status	Completed
Mode of transport	Aviation
Aviation occurrence category	Abnormal engine indications
Occurrence class	Incident
Highest injury level	None

Aircraft details

Manufacturer	The Boeing Company
Model	767
Registration	VH-EAN
Serial number	23402
Sector	Jet
Operation type	Air Transport High Capacity
Departure point	Sydney, NSW
Destination	Perth, WA
Damage	Minor

Boeing 747-200, DQ-FJI

Factual Information

Following flap retraction shortly after the Boeing 747 departed Sydney, the crew reported a flap disagreement indication. The flaps could not be extended and were stuck in the up position. The aircraft returned to Sydney where the crew carried out a flapless landing.

An inspection revealed that the right inboard fore-flap carriage stop had failed. This resulted in extensive damage to the fore flap, mid flap, and overload failure of the right flap drive torque tubes.

Specialist examination found that the fore-flap carriage stop failed due to stress corrosion cracking. The cracking initiated at the bolt hole surface. The manufacturer's instructions require that the bolt should be installed with wet sealant, however inspection indicated that the sealant did not cover the centre portion of the bolt, potentially allowing moisture to enter.

Occurrence summary

Investigation number	199901111
Occurrence date	16/03/1999
Location	Sydney, Aero.
State	New South Wales
Report release date	02/01/2001
Report status	Final
Investigation type	Occurrence Investigation
Investigation status	Completed
Mode of transport	Aviation
Aviation occurrence category	Flight control systems
Occurrence class	Accident
Highest injury level	None

Aircraft details

Manufacturer	The Boeing Company
Model	747
Registration	DQ-FJI
Serial number	22145
Sector	Jet
Operation type	Air Transport High Capacity
Departure point	Sydney, NSW
Destination	Nadi, FIJI
Damage	Substantial

Boeing 737-377, VH-CZL

Factual Information

History of the flight

At the appropriate time during the approach sequence, the co-pilot of the Boeing 737, who was the handling pilot for the sector, called for the landing gear to be extended. When the pilot in command placed the landing gear lever to the "down" position, a loud thump was heard and the "gear safe" green light for the right main gear illuminated immediately. This was followed by the illumination of the left main and nose landing gear lights, consistent with a normal extension sequence. The aircraft rolled approximately 4 degrees to the right while the gear was extending. This was counteracted by a left roll control input.

As the crew had received indications that the landing gear was safely locked down, they continued the approach and completed a normal landing.

Ramp maintenance staff briefly inspected the aircraft but did not find any immediate cause for the reported thump. The aircraft was then placed on jacks for a retraction test. When the landing gear lever was selected to the "up" position, the right main landing gear moved inboard approximately 15 cm before a grinding noise was heard. The test was immediately suspended and the landing gear was extended.

When access panels were removed, it was found that the actuator beam arm inboard lugs and beam hanger had fractured. The rear wing spar, landing gear beam, aileron bus cable, pulley bracket, aileron and spoiler cables and hydraulic lines had been damaged extensively following the fracture of the lugs and hanger.

Actuator beam arm examination

Inspection of the actuator beam arm revealed that the fracture of both lugs was due to stress corrosion cracking.

Service information

The manufacturer addressed main landing gear (MLG) corrosion and cracking problems with the issue of Service Bulletin (SB) 737-32A1224 in July 1989 and with production changes to the beam arm assembly. Revision 2 of this SB was issued in April 1991 and Part A of this SB was mandated by the issue of Federal Aviation Administration (FAA) Airworthiness Directive (AD) 91-05-16, and by the issue of Australian Civil Aviation Safety Authority AD/B737/57. The AD action required in-situ inspection of the beam arm at 600 flight-cycle intervals. Alternatively, modification or replacement of the beam arm with a new production arm could be accomplished as a terminating action.

The modified beam arm assemblies incorporated the following changes to improve the corrosion resistance of the actuator beam arm and actuator beam attach bolt:

- improved bushings,
- an increased bushing interference fit,
- an improved actuator beam bolt,
- more extensive cadmium plating, and
- improved lubrication of components.

SB 737-32A1224 Rev 3, issued in October 1992, recommended replacing the actuator beam arm components with production components rather than part modification.

There have been two reports of lug failures on beam arm assemblies that had been modified by AD 91-05-16. These failures occurred overseas in March 1997 and October 1998 and were similar in nature to this incident.

Maintenance history

The operator advised that in July 1991, the right beam arm fitted to the aircraft involved in the March 1999 incident was reworked for incorporation of SB 737-32A1224 Rev 2 at the operator's maintenance facility. In June 1993, the right MLG was overhauled at an overseas maintenance facility. In June 1993, the beam arm was installed with the MLG onto the aircraft. At the time of failure, the beam arm had been in service for 33,681 flight hours, 25,025 flight cycles and 10.25 years. The period since last overhaul was 17,879 flight hours, 12,296 flight cycles and 5.75 years.

ANALYSIS

When the aircraft was on the line it showed no visible damage, however, the operator elected to carry out a more comprehensive inspection. Fracture of the lugs causes the main landing gear to fall back to the extended position and is therefore fail-safe in that respect. However, an actuator beam arm fracture allows the actuator, actuator beam, and support link to travel outboard, contacting and possibly damaging the main landing gear beam and the wing rear spar. It can also result in contact with nearby flight control cables (aileron and spoiler) and hydraulic lines. This contact has the potential of inadvertent flight control inputs that could reduce the controllability of the aircraft.

The aileron bus cable in this incident was found to be almost completely severed. This damage potentially could have resulted in serious control problems.

The aircraft right roll experienced during gear extension was considered to have been the result of the right landing gear extending more rapidly than the left landing gear, and not as a result of flight control inputs caused by the actuator beam arm failure.

The location of the fractures in both inboard lugs of the right actuator beam arm coincided with the positions for possible cracks identified in SB 737-32A1224. The fractures in the lugs were a result of stress corrosion cracking. Previous failures analysed by the manufacturer were also attributed to stress corrosion cracking.

Stress corrosion cracking of high-strength steel components of aircraft main landing gears occurs when the components are subjected to a sustained tensile stress and are exposed to an environment that allows stress corrosion cracking to initiate. The components are susceptible to stress corrosion cracking when exposed to the normal operating environment of landing gear (moisture, salt laden moisture). Stress corrosion cracking in the actuator beam arm lugs occurred as a result of the movement of the bushes installed in the lugs and the penetration of moisture into the gap created between the bushes and lugs.

The ATSB Technical Analysis Report 19/01 further analysed failures of B737 main landing gear pin/lug joints and is available on the ATSB website or on request from the ATSB.

SAFETY ACTION

Operator fleet inspection

As a result of this incident, the operator conducted visual and ultrasonic inspections of the actuator beam arms of its fleet of Boeing 737 aircraft. A cracked actuator beam arm was found during inspection of another aircraft.

As a result of this occurrence, the ATSB (which includes the former Bureau of Air Safety Investigation) issued interim recommendations (IRs) to the manufacturer, regulators and operators on 19 March 1999 and identified the following safety deficiency:

"Damage to flight control cables and primary aircraft structure resulting from the failure of main landing gear (MLG) actuator beam arm lugs of B737 aircraft has the potential to seriously affect the safety of flight."

IR19990019

The Bureau of Air Safety Investigation recommends that Boeing Commercial Airplane Group alert Boeing 737 operators to this safety deficiency and implement an appropriate inspection program.

The following response (in part) was received from the Boeing Commercial Airplane Group on 6 April 1999:

"We have released Boeing All- Operator Message M-7200-99-02383 dated 26 March 1999. In summary we plan to release an alert service bulletin (737-32A1314) by the end of April, 1999. This bulletin will call for the resumption of the ultrasonic inspections of all Main Landing Gear Beam Arms every 600 flight cycles upon the accumulation of 10,000 cycles or 4 years (whichever comes first). This will be considered interim action until we have analyzed and implemented the course of terminating action to preclude the onset of corrosion noted in the parts."

Boeing released alert service bulletin 737-32A1314 on 15 April 1999 requiring all 737-100, -200, -300, -400, -500 airplane operators to do a visual and ultrasonic inspection of the clevis on the actuator beam arm. If cracks or corrosion are found, the actuator beam arm is to be replaced. Boeing recommended this inspection be done before the actuator beam arm has 10,000 cycles or four years of service. Boeing recommended the inspection be repeated every 600 flight cycles or 90 days (whichever occurs first).

Boeing advised that they were evaluating design improvements that will further inhibit the initiation of corrosion of these high strength steel parts. Boeing expected to complete the evaluation and begin implementation of the new hardware by the fourth quarter of 2002.