The aircraft departed Whyalla for Kalgoorlie, tracking via Olympic Dam to pick up passengers, then Forrest to refuel.  Later in cruise, between Olympic Dam and Forrest, the pilot became aware of power fluctuations originating from the right engine which he attributed to a faulty density controller.

At Forrest the aircraft was refuelled to its maximum capacity.  After departure the pilot noted that the right engine manifold pressure and RPM indications lagged behind the left engine, and it also failed to obtain correct manifold pressure indications from 4,000 ft above mean sea level (AMSL), and higher.

As the aircraft climbed through 5,000 ft AMSL, approximately 55 km west of Forrest, a passenger attracted the pilot’s attention to the right engine which he believed was on fire.  The pilot noticed smoke and also thought he could see flames so immediately turned the aircraft back towards Forrest, advising Perth Flight Service of the situation. He set the right mixture control to "idle cut-off", then dived the aircraft to increase airspeed.  This being a flight manual recommendation as a means to extinguish engine fires.

After descending to about 3,000 ft AMSL the right propeller control was selected to the feather position, but the propeller failed to feather and continued to rotate.  The pilot then made several more unsuccessful attempts to feather the propeller, but without having increased power to the left engine the airspeed decayed.  The aircraft continued to descend to about 100 ft above ground level (AGL), but with the windmilling right propeller, and now using maximum power on the left engine, the pilot was unable to prevent the airspeed from further decreasing below the best single engine rate of climb speed while attempting to maintain level flight.  About 26 km from Forrest the pilot realised that the aircraft would not maintain height any longer and elected to carry out a forced landing.

The aircraft touched down on an uneven, rocky surface, then after a ground roll of about 390 metres the landing gear collapsed.  The aircraft slide for a further 20 metres, turning through 60 degrees to the right.

The uninjured occupants evacuated the aircraft and walked about a kilometre to the railway line where they hailed and boarded a freight train, which had been alerted to their problem and enabled them to continue their journey to Kalgoorlie.

There were no signs of a fire in the right engine compartment, but there was evidence that a considerable amount of oil had sprayed onto the exhaust pipes and surrounding structure etc near the turbocharger.  The engine oil levels had been checked prior to departure from Whyalla, where they both indicated about 11 quarts.  After the accident the right engine oil level was found to be 6 quarts.

A test run of the right engine found it capable of developing normal power.  Examination of the right turbocharger revealed it had a worn bearing seal which allowed oil to escape during engine operation.

The right propeller blade angles appeared to have been in the normal operating range at impact, and further inspection of the propeller, after disassembly, indicated there was no pre-impact damage which could have prevented it feathering.

Inspection of the aircraft maintenance logbooks showed that the constant speed unit (CSU) on the right engine had been installed for more than 1,000 hours beyond the approved overhaul period for this type of component.  A bench test revealed its feathering operation to be very slow, apparently caused by the speeder spring lift rod being out of adjustment.  The propeller pitch control lever was operated several times through its full range while the pilot attempted to feather the propeller, each occasion having the effect of re-starting the slow feathering cycle which possibly allowed the engine RPM to reduce to a speed where the latch mechanism engaged, preventing the propeller blades from moving to the feathered position.

The induction manifold upper deck air sense hose to the turbocharger differential pressure controller was found to be charred, in poor condition and separated from its end fitting at the differential pressure controller attachment.

A partially melted aluminium heat shield, located between the turbocharger and the hose, showed a definite pattern of a hot gas flow in a path from a leaking exhaust flange gasket at the turbocharger to the hose end fitting.  This released the clamping pressure of the hose to the end fitting nipple, causing it to separate.  The detached hose would have caused the power lagging and reduced manifold pressure as reported by the pilot.

The hose was also found to be covered with oil from the leaking turbocharger seal, which when heated by the exhaust gases may have produced a flame as seen by the pilot and passenger.  The smoke was caused by oil spraying onto the hot exhaust pipe surface.

An incorrect empty weight was used on the trim sheet when calculating the weight and balance for this flight, allowing the aircraft to be loaded in excess of its maximum take-off weight on departure from Forrest.  This may also have had an influence on the aircrafts inability to maintain height.

The pilot was conducting agricultural spraying operations on a 12 ha cotton crop. He said he had a good knowledge of the field and had sprayed it several times during the season. He was aware of two power lines running across the field and planned to fly under them. One powerline ran diagonally across the direction of the intended spray run.

After passing under this wire the pilot felt a vibration through the helicopter. Although it continued to fly in a straight direction, the helicopter would not respond to the pilot's control inputs. It then descended, entered the crop, and slid along the ground for 20 m before coming to rest.

The pilot assessed that he had maintained inadequate clearance from the powerline, and that the main rotor blades had brought the powerlines into contact with the tail rotor. The clearance below the wires was reduced by the fully grown cotton. The diagonal orientation of the wires and reduced ambient lighting may also have impaired the pilot's judgement of the available clearance.

A visiting Japanese glider pilot was undergoing conversion to the "Junior" glider. On her first flight she appeared to be aiming too close to the tugs and pie cart on the left of the strip, but she successfully corrected and continued for a normal landing.

The instructor judged that she understood the problem and briefed her for the next flight which was to be solo. On final approach she again aimed too close to the parked obstructions but this time she failed to take corrective action.

During the landing roll the left wing of the glider was substantially damaged when it collided with a parked tug aircraft. The collision swung the glider through 180 degrees. The tug was less severely damaged.

Significant Factors

The following factors were considered relevant to the development of the accident:

1. The pilot did not position the glider to give adequate obstacle clearance.

2. The glider collided with a parked tug.

A anonymous ground witness observed the Cessna 206 overfly the windsock before commencing an approach to land on the north westerly strip. Most of the time there appeared to be a 20-knot crosswind from the left, but occasionally the wind swung so that there was a tailwind component. The witness advised that after touching down about 100 metres into the 490-metre strip, the aircraft bounced once and floated until almost two thirds down the strip where, after touching down again, the brakes appeared to be applied. Shortly after the brakes were applied, the wheels appeared to lock causing loss of directional control. The aircraft departed the centreline to the right, struck bushes about 70 metres from the end of the strip and overturned.

According to the pilot, he performed a normal short field approach and landing uphill on the 310 degree strip with the wind 220 degrees gusting from 10 to 15 knots and the aircraft touched down near the threshold without bouncing or floating. During the landing roll the pilot believed that the wind changed to provide a significant tailwind component which made the aircraft difficult to slow down and control. The pilot did not attempt a go-around.

FACTUAL DATA

1. The flight

Passing through 15,000 ft. during climb out of Canberra the left engine N1 and N2 rpm indicators began to fluctuate to 10% of engine rpm. This was accompanied by fluctuating ITT and fuel flow. At this time the engine was being controlled by the onboard computer in the Auto mode, therefore, in accordance with flight manual procedures, the pilot selected Manual mode. The engine parameters then steadied, apart from N2 which continued to fluctuate by 10% rpm. It was then noted that the oil temperature had increased to the top of the green arc on the gauge. A reduction in power had little or no effect on the temperature, however the oil pressure remained normal.

Approximately five minutes later the engine flamed out with N1 observed to be windmilling and N2 stationary. The engine was secured and the aircraft continued for an asymmetric landing at Essendon where it is based.

2. The investigation

The starter motor was removed and the engine hand turned to check for freedom of rotation. Although a slight rumble could be heard from the vicinity of the transfer gearbox the N1 appeared to turn freely. However, on removal of the cover plate from the transfer gearbox it became apparent that the horizontal bevel gear was damaged. The freedom of rotation of the N2 assembly was then checked with no apparent problems noted.

The engine was removed from the airframe and, when re-fitting the cover plate to the transfer gearbox in preparation for dispatch to the overhaul facility, a screw and washer were found loose in the bottom of the transfer gearbox. A borescope inspection of the gearbox found that the top screw, P/No. NASllOIE3-8, was missing from the cylindrical roller bearing housing in the transfer gearbox assembly p/n 3070093-4.

The missing screw was one of three that holds the cylindrical roller bearing in place. During assembly these should be torqued to 40 in lbs. A torque check prior to disassembly of the other two screws showed that they were below that torque and both needed a further quarter turn before attaining the required 40 in lbs torque.

The horizontal bevel gear P/No. 3070212-3 had impact damage on the back of the gear and also damage to the gear teeth. The vertical bevel gear P/No. 3070219-2 had sustained severe damage to the gear teeth. The transfer gearbox housing assembly P/No. 3070429-6 was foreign object damaged (FOD) over its internal surfaces, and the transducer P/No. 3070722-3 was severely damaged. Metal contamination was found throughout the gearbox components including the bearings. The engine oil pump was found FOD contaminated and completely seized.

The gearbox had operated 4,253 hours and had not been disassembled since new.

ANALYSIS

The top cylindrical roller bearing screw P/No. NASllOlE3-8 became dislodged in service resulting in substantial secondary damage to the transfer gearbox components.  Considerable metal contamination of both the transfer and accessory gearboxes occurred. The failure of the transfer gearbox resulted in the break-down of the drive to the accessory gearbox which in turn resulted in the fuel pump failing to deliver adequate fuel to the engine resulting in flame out. The considerable metal contamination as a result of the failure also resulted in the oil pump seizing.

The fluctuations in N1 and N2 rpm coupled with the fluctuating ITT and fuel flow as reported by the pilot were genuine fluctuations as a result of the transfer gearbox failing, with consequent intermittent drive to the accessory gearbox and fuel pump. The fact that the engine settled down again when manual mode was selected is considered to be a coincidence. The continuing apparent fluctuations in N2 rpm would have been due to the transducer beginning to sustain the damage which eventually rendered it unserviceable. The reason for the N2 gauge failing to read for the remainder of the flight after engine flameout, although the LP shaft was windmilling, is because of the damage to the transducer rendering it incapable of sending a signal to the indicator.

It is considered that during initial installation the bolts may have been oil lubricated. This procedure is known to result in a lower torque than that achieved in a dry installation.

The gear box installation on all jet engines is subject to heat and vibration. If the securing screws were not properly installed, and were not properly tightened to the designated torque, the operating environment would be conducive to those screws loosening and becoming displaced.

Significant factors

The following factors was considered relevant to the development of the incident.

1. During initial assembly of the gearbox the three cylindrical bearing retaining screws were not tightened to the specified torque.
2. One cylindrical bearing retaining screw loosened and became displaced.
3. The displaced screw caused distress to the gearbox components sufficient to fail the gearbox drive to the engine fuel pump.
4. The engine flamed out due lack of fuel.

GLOSSARY OF TERMS

rpm     engine speed in revolutions per minute

ITT      engine operating temperature (Interstage Turbine Temperature)

N1       engine HP core speed

N2       engine LP fan speed

FOD    foreign object damage

in lbs   inch pounds - the measure of torque applied to a screw

LP       engine low pressure (N2) stage

The pilot of VH-OGN advised that whilst on descent approaching their cleared altitude of 5000 feet via a Hopla 1 STAR for runway 16, they received a TCAS proximity traffic warning with no altitude readout. The intruding traffic was sighted and was reported to be similar to either a Cessna 182/206, tracking from left to right and about one and one half nautical miles to their left. VH-OGN was levelled at 5400 feet and banked to the right.

The other aircraft maintained its track and passed approximately half a nautical mile behind VH-OGN. The captain of VH-OGN said that while this was occurring, the TCAS changed to a "yellow traffic with aural warning". He estimated that the Cessna was between 4800 and 5000 feet.

Although the Cessna was squawking code 2000, there was no altitude readout on air traffic control screens. Attempts to contact and identify the Cessna were not successful.

The student pilot was engaged in a solo cross training flight at the time of the accident. As he was approaching the aerodrome he reported that he checked the windsock and noticed that the wind appeared to be from the north-east at approximately 10 knots. The pilot entered the circuit area and completed an approach and landing on runway 07. The aircraft touched down, on the gravel centre section, approximately 300m in from the threshold of the 1221m long strip.

The pilot observed that the aircraft was not slowing as fast as expected and he applied full braking. When it became apparent that the aircraft would not stop by the boundary fence the pilot shut the engine down and attempted to steer the aircraft to the right, on to the grass and away from the fence.

The left wing collided with a fence post, swinging the aircraft through 140 degrees and the aircraft stopped on the gravel verge of a two-lane road with the right wing blocking one lane.

As the pilot exited the aircraft, he observed that the windsock indicated a tailwind of approximately 20 knots on runway 07.

The pilot advised that during the landing roll on bitumen runway 36, the aircraft swung sharply to the right. He applied left brake to correct the swing. The aircraft began to straighten but then it veered sharply to the right again like a ground loop. The left landing gear leg collapsed, and the left wing struck and propeller struck the ground. At the time, the wind was gusting 10 to 15 knots from 010 degrees.

After the accident, the pilot thought that the left brake might have failed causing him to lose directional control. An engineer inspected the brakes and advised that there was no evidence of a brake failure. However, the engineer rebuilding the aircraft discovered a pre-existing fatigue crack through about half the left landing gear leg. It is probable the fatigued landing gear leg failed during the landing roll.

According to the engineer, the landing gear legs in VH-ICA were of the thicker metal version. Therefore, they were not subject to mandatory magnetic particle method crack testing as were the earlier version of spring legs in accordance with Airworthiness Directive AD/Cessna 188/3.

Significant Factors

This accident was not the subject of an on-scene investigation and there was insufficient evidence available from other sources to determine the factors that led to the accident.

Safety Action

The maintenance organisation rebuilding VH-ICA advised that a defect report would be submitted to the Civil Aviation Authority concerning the pre-existing crack found in the landing gear leg.

The Bureau of Air Safety Investigation advised the Civil Aviation Authority of the fatigue crack by way of the Bureau's weekly summary of occurrences.

Whilst in the circuit, the pilot noticed that the landing gear down light was not illuminated. He checked the visual drive indicator in the cabin floor and the external cowl mirrors which indicated to him that the gear was down. During the initial landing roll, the gear collapsed, and the aircraft came to a halt on the runway.

Engineering advice indicated that the gear was not fully down when the aircraft landed. This had been caused by the seizure of the gear motor. It was possible for the pilot to have believed the gear was down from the view he had in the mirrors.

Prior to this occurrence, the pilot had experienced problems with the dimming function of the landing gear indicating light. He had previously landed during the day with the gear light dimmed and thought that this was again a light dimming problem.