The student pilot was operating the last sector of a solo navigation exercise, tracking from Murray Bridge to Adelaide where a touch-and-go landing was made, before returning to Parafield at 1,500ft.

Parafield Tower cleared the aircraft for a straight-in approach to runway 03 L, with a clearance to land. The pilot reported on final at 1,500ft, but the aircraft was not initially sighted by the controller. It was then observed north west of the field, inside the RAAF Edinburgh Control Zone.

An Orion aircraft, conducting right hand circuits for runway 36 at Edinburgh, was on right base heading west when the other aircraft entered the control zone.

Radar plots revealed that VH-YHH passed behind the Orion at about the same level, with a horizontal separation of about 1000 metres.

The student pilot then made a right turn to enter left downwind for runway 03 L, Parafield, and continued for a landing.

The pilot stated that on Monday 21 August 1995, he was the captain of VH-JOC rostered to fly from Cairns to Weam via Daru. The aircraft departed Daru for Weam as per flight plan and arrived overhead Weam at 1215 pm local time. The pilot said he observed a Piper PA-31 Navajo, and a BN-2 Islander aircraft parked in the parking area. The pilot of the Islander had previously provided the pilot of VH-JOC with a strip report and had advised that the strip was serviceable.

The pilot stated that a normal circuit entry, approach and landing were conducted. The aircraft touched down normally and at approximately 60 kts braking was applied. The right main gear then appeared to lock up or strike something. Braking was released but the aircraft yawed to the right. The pilot said he tried to correct the yaw with rudder, but it became worse. The nosewheel appeared to dig into the runway surface and he was unable to hold the weight off the nosewheel due to the yawing effect. The gear then collapsed, and the aircraft slewed further right before coming to rest.

Subsequent examination of the right main gear by an engineer revealed that the pivot bolt had pulled through the torque links after the retaining washer had been sheared. This failure then allowed the wheel to castor outwards and resulted in loss of directional control and collapse of the left main and nose gear.

Two pilots who witnessed the accident said that there was a strong crosswind from the left, backing to a tailwind component on the landing. They said the aircraft was going sideways and looked as if it was going to ground loop prior to the failure of the torque link.

The Boeing 767 was being processed by the Approach North controller (APPN) for a visual approach to runway 16R. A clearance was issued to descend to 2500 ft, together with radar vectoring to position the aircraft for left base. At the same time a helicopter was operating in the Middle Harbour area at altitudes between 1000 ft and 1500 ft.

The Boeing was equipped with a Traffic Collision Avoidance System, version two (TCAS II). TCAS II contributes to collision avoidance protection by issuing traffic alerting and conflict resolution advice to the flight crew.

As the Boeing descended on an extended base leg the crew informed APPN that they were responding to a TCAS Resolution Advisory (RA) and had initiated a climb. Although the controller reported that the Boeing was 1000 ft above the helicopter, the crew responded with, "Roger, have 500 ft indicated". Shortly after, the pilot of the helicopter reported sighting the Boeing and indicated there was no confliction. The Boeing was instructed to descend to 1500 ft and subsequently carried out an approach and landing without further incident.

Later analysis of radar data indicated that the Boeing had been descending at approximately 1000 fpm but discontinued its descent at 2800 ft pressure altitude and commenced to climb, consistent with receiving the RA. The helicopter had been maintaining level flight at approximately 1000-1200 ft pressure altitude. At the time the crew of the Boeing initiated the climb the vertical separation with the helicopter was 1700 ft, with a horizontal separation of 1.27 NM. The minimum required separation standard was 1000 ft vertical separation or 3 NM horizontal.

Because TCAS II software cannot anticipate that an aircraft may level off, the predicted flight paths of the Boeing and the helicopter, as determined by the TCAS II software, was most likely sufficient to activate the RA. It could not be determined why the crew of the Boeing reported the helicopter being 500 ft below their aircraft, as indicated on their TCAS II cockpit display.

The aircraft was cruising at 2,000 ft above ground level when the pilot heard an unusual noise coming from the engine. He immediately placed the aircraft into an autorotation descent. The chip detector light illuminated, the airframe began to vibrate, and the engine-out warning sounded prior to landing. The pilot was able to complete a safe but heavy emergency landing in a heavily timbered area.

Inspection indicates the engine failure was caused by the decoupling of the power module. Once separated, the fourth stage turbine wheel breached the turbine casing.

The turbine assembly and the fourth stage wheel had completed 1612 hours time in service since they were last overhauled. The turbine had 1887 and the wheel had 2032 hours to run to their next overhaul.

Examination of the power turbine components indicates that the power turbine coupling fractured in the reduced section of the coupling, aft of the region that forms the number six bearing inner race. The number six bearing components were also extensively damaged. Most of the damage was consistent with eccentric rotation of the power turbine assembly.

A blade had been lost from the fourth stage turbine wheel. The shroud ring and blades are an integral part of the wheel. Blade separation requires fractures in both the shroud ring and the blade aerofoil section. The shroud had fractures on both sides of the blade. The blade had fractured near its base. Fatigue cracking was associated with all three fractures. The failures in the shroud ring were consistent with slow fatigue crack growth under a uniform loading environment over many engine cycles (low cycle fatigue). The failure in the blade was consistent with crack growth in a variable amplitude loading environment over a lesser number of engine cycles (high cycle fatigue). No abnormalities were found which may have contributed to fatigue crack initiation.

The different nature of the fatigue cracks in the shroud and blade indicate that the shroud fractures preceded the crack in the blade. The fracture of an integrally cast shroud ring on both sides of a turbine blade would result in a changed loading condition that could lead to premature blade failure. The loss of a blade from the fourth stage wheel would create an out of balance condition. This probably led to failure of the power turbine rotor assembly.

The circumstances which caused fatigue cracking and final failure in the shroud could not be established.

As part of an endorsement training exercise the captain operated the emergency electrical Bus to simulate an electrical fire. When the Bus was selected all electrical power was removed from the aircraft. Later examination found a broken wire in a relay box. It was found that the fault was not detected during previous maintenance test/checks because the test circuit goes to the earth side of the circuit, and the breakage occurred on the power side.

The operator is currently corresponding with the aircraft manufacturer to recommended a change to the maintenance manual procedural checks that will identify any similar faults during ground tests.

FACTUAL INFORMATION

The helicopter had undergone a maintenance check which included changing both forward and aft coupling assemblies of the main driveshaft and the forward boot assembly. Prior to release for service, vibration monitoring tests were carried out on the ground and during flight. On the final flight, whilst climbing to 4,000 ft in preparation for a descent test, the maintenance crew observed an increase in longitudinal vibration of the main driveshaft. On reaching 4,000 ft, there was a noticeable grinding noise and all power to the main rotor was lost. Following the driveshaft failure, the pilot carried out an autorotation into the only available open field. The field initially appeared flat but as the helicopter approached the ground level, the pilot saw that it was undulating and sloped upwards in the direction of his approach. The pilot was unable to correct for this in time to avoid a heavy landing.

A coupling is fitted to each end of the driveshaft. Each coupling is a splined unit packed with grease as a lubricant between its inner and outer components. A rubber boot assembly bolts to the outer coupling to contain the grease as the assembly rotates. An O-ring packing is retained in a groove in the boot as a seal between the boot and the outer coupling.

Investigation found that the main driveshaft forward coupling assembly had failed. The outer coupling remained intact, but the splines of the inner coupling had failed. A metallurgical examination determined that, though the coupling had probably met manufacturing specification, it had been subjected to high temperature which resulted in weakening of the core material underlying the nitrided zone. The splines were then unable to transmit the applied torque to the outer coupling and failed from overload. The associated rubber boot assembly had disintegrated, its attaching bolts were loose, and the locking wires were broken. There was some charred debris in the O-ring groove of the boot, which was probably the remains of the O-ring which formed the seal between the boot and the coupling. The grease lubricant, normally retained in the coupling by the boot assembly, had been ejected and was spattered over the driveshaft well.

The forward boot assembly had been requisitioned from stock under the existing part number. However, the boot had been modified by the manufacturer, given a new part number, and issued as a replacement part. The modification included securing the grease dam in the boot assembly by a process known as roll staking. This process consisted of mechanically removing material from the O-ring groove wall and rolling it to lock the dam in position. The operation modified the O-ring groove profile so that steps were formed in the inner circumference of the groove. There was no information provided by the manufacturer to advise maintenance personnel of the modification.

ANALYSIS

The failure of the driveshaft coupling was consistent with a failure of the O-ring. This provided an open path for the grease to be ejected by centrifugal action as the driveshaft rotated at its operating speed of 6,000 RPM. Failure of the O-ring would immediately reduce the torque on the attachment bolts. It is likely that the ensuing vibration resulted in failure of the locking wire. The loss of the grease resulted in overheating of the coupling gears which eventually failed and all power to the main rotor was lost.

Maintenance personnel, unaware of the modification to the O-ring groove, probably allowed the O-ring to be pinched between the boot and the coupling during assembly. Bolt torque would have appeared normal but during operation the O-ring would fret between the mating surfaces until it failed, leaving a path for the grease to escape.

SIGNIFICANT FACTORS

1. The manufacturer did not provide adequate details of the modification to maintenance personnel.
2. Maintenance personnel, unaware of the modification to the boot, probably allowed the O-ring to be pinched when the boot was mated to the coupling during assembly.

SAFETY ACTION

The Bureau of Air Safety Investigation made the following recommendations to the Civil Aviation Safety Authority and to Bell Helicopter Textron Inc on 23 January 1997:

"R960155

"The Bureau of Air Safety Investigation recommends the Civil Aviation Safety Authority advise all Bell Helicopter

412 operators in Australia of the potential, during installation, for the O-ring (PN 204-040-691-003) to sag in the

O-ring groove of the (PN 212-040--688-003) coupling and hence, become pinched between it and the (PN 212-040-176-101) boot assembly. This would allow a path for the grease lubricant to escape due to the centrifugal forces acting on the driveshaft during its normal operation, thereby causing the coupling to overheat and fail." "R960156

"The Bureau of Air Safety Investigation recommends that Bell Helicopter Textron Inc. include a cautionary note in the relevant section of the Bell 412 maintenance manual (BHT-412-CR&O, at 63-9 ASSEMBLY-MAIN DRIVE

SHAFT), warning of the possibility for the O-ring to sag during installation in the replacement PN

212-040-176-101 boot assemblies and hence, allow the potential for pinching to occur between the boot and coupling assemblies. This would allow a path for the grease lubricant to escape due to the centrifugal forces acting on the driveshaft during its normal operation, thereby causing the coupling to overheat and fail."

1.3 The rejected take-off was accomplished at a speed reported by the captain to be approaching V one (V1), however the Flight Data Recorder information could not be retrieved, and the actual speed is unknown. V1 is the maximum speed from which the aircraft can be safely stopped when take-off is discontinued. The aircraft was stopped at taxiway Bravo 4 (B4) which is approximately 2200 metres from the threshold of runway 15.

1.4 Examination of procedures in the Qantas Airport Services Manual revealed that there was no standard procedure for loading and unloading a container compartment. The manual states that many variations of loading and unloading can be accomplished with proper selection of control switches and restraint hardware. The design of the sill locks is such that although they are painted red, there is no visual indicator to warn that they are not locked down before the compartment door is closed. There is no reason to believe that had the loader not been interrupted in his task that he would not have put the sill locks down. The leading hand who asked the loader to man the tug for pushback, omitted to check that the locks were engaged before he closed the door. He stated that this was due to a variance of procedure, because he deploys the locks before lowering the bridge, and the loader deploys the locks after lowering the bridge. The interruption to the procedure, at the point where the locks would have otherwise been engaged was thus unfortunate and led to the omission.

1.5 The sill locks provide vertical and lateral restraint on one side of the container only. The container was restrained vertically and longitudinally by the remainder of the locking devices and could not have moved if the aircraft had become airborne.

3. CONCLUSIONS

Findings

1. A container in the rear cargo hold at position R44 was not fully restrained by the aircraft locking system.
2. The persons responsible were aware of their responsibility, but due to task interruption, and variance in procedures, a final check to ensure the sill locks were activated was omitted.
3. There was no warning device or mechanical means to prevent the door from being closed when the sill locks were not activated.
4. There was no specific procedure laid down for loading or unloading, however the Load Instruction Report certifies that the aircraft has been loaded in accordance with the Aircraft Load and Balance Manual, and that the load is secured by the aircraft locking system.

Significant factors

1. The aircraft departure was late.
2. The loading team changed positions to save time.
3. Loading procedures were not standardised.

During a ferry flight the aircraft's engine began to run rough, then stopped about 20 seconds later.

The pilot made a successful forced landing into scrub-covered terrain in an inaccessible area of a national park, with the aircraft suffering only minor damage.

An examination of the engine revealed the cam ring had fractured.

The pilot reported that, during the landing roll, as the aircraft slowed through about 40 knots, three cows ran on to the runway. The right propeller of the aircraft struck one of the cows and the pilot was unable to prevent the aircraft running off the end of the strip into a fence.

The pilot indicated that the runway was clear when he checked it on downwind, base, and finals. The strip was surrounded by an electric fence, but the fence was not activated at the time the aircraft landed.

During cruise the aircraft suffered an engine failure with oil covering the windscreen. The pilot carried out a successful landing onto a beach 40 km west of Milingimbi.

Later as the tide filled the aircraft, which was undamaged, became immersed in salt water.

An inspection of the engine indicated that the number six cylinder gudgeon pin had failed allowing the connecting rod to separate from the piston. As the engine continued to rotate the loose connecting rod struck and holed the crankcase.