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.

The instructor and student were practising engine out landings, to touchdown, on a grassed area of the airport. On the accident approach, ground speed had been reduced to below five knots before the aircraft touched down. As the aircraft slid along on its skids they dug into the soft ground and the helicopter pitched forward until the main rotor blades struck the ground. The helicopter then rolled onto its right side.

The instructor was aware of the soft ground and was monitoring the approach carefully. Because the ground speed was low, the instructor was not expecting any problems with the landing and, as a result, he had insufficient time to take corrective action before the main rotor blades struck the ground.

The aircraft stalled and touched down heavily following an approach for a landing.

The student pilot was practicing a glide approach under the supervision of the instructor. The student misjudged the flare for landing and the aircraft ballooned to a height of about 25ft above the runway. The instructor applied some power as he told the student to lower the nose. The student was slow to react, and the airspeed reduced rapidly. The instructor took the controls, lowered the nose and applied full power but he was unable to prevent the aircraft from stalling and landing heavily.

Crew

In later discussion, the pilot in command stated that he was attempting to comply with the request from Canberra Approach to maintain a maximum speed descent. He believed that the aircraft would pass the 20 NM range circle indicated by the aircrafts Distance Measuring Equipment (DME), above 6,000 ft which was the lower limit of controlled airspace at that distance. He thought that it had done so as the aircrafts DME indicated 17 NM from Canberra when GPWS warning was received. The first officer, who was the handling pilot during the incident, stated that he checked the DME reading when the warning was received. At 17 DME, the lower limit of controlled airspace was 5,000 ft.

Flight data recorder information

The flight data recorder (FDR) data confirmed that the aircraft was descending through 6,000 ft when the GPWS activated. FDR data also confirmed the GPWS warnings and the 300 ft climb.

Recorded values of latitude and longitude revealed that the GPWS warnings were activated as the aircraft overflew Bulls Head, spot height 4,482 ft, approximately 20 NM on the 240-degree radial from Canberra VOR. The data confirmed that a valid Mode 2A Excessive Terrain Closure GPWS warning had occurred, and the crew had taken the appropriate actions to remove the aircraft from the GPWS activation envelope.

The FDR data showed the aircraft on descent at an airspeed of approximately 315 Kt on a heading of 078 degrees magnetic when the GPWS warning occurred. Recorded radar altimeter data indicated showed rates of closure with terrain in excess of 9,000 feet per minute (fpm). The radar altitude reduced from 2,658 to 2,044 ft in a three-second period.

ANALYSIS

GPWS parameters

The GPWS will give a Mode 2A warning if the aircraft is flying with less than 25 degrees of flap and there is an excessive rate of closure with rising terrain. If the airspeed exceeds 250 kts, the upper boundary of the radio altitude envelope is 2,450 ft with a rate of closure of 5,105 fpm and the lower boundary is 50 ft with a rate of closure of 2,063 fpm for a "TERRAIN, TERRAIN" warning. Within this envelope is the envelope triggering the "WHOOP WHOOP PULL UP..." warning.

The combination of the aircrafts rate of descent and the rising terrain on the western side of the Bulls Head resulted in a rate of closure in excess of 9,000 fpm and a radar altitude of 2,044 ft which was within the GPWS inner operating envelope.

Pilot’s intentions and results

The pilot’s intentions were to maintain the maximum speed during descent and to pass the 20 DME range circle just above 6,000 ft, continuing descent to 5,000 ft when inside the 20 DME circle. Both radar and FDR data revealed that the aircraft descended below 6,000 ft beyond the 20 DME circle and thus operated outside controlled airspace (OCTA) for a short time.

An aircraft operating at an altitude which is the published lower limit of controlled airspace is operating OCTA. Another aircraft, about which the air traffic services held no information, could have been operating OCTA at the altitude reached by the B737.

As both pilots reported that the aircrafts DME indicated 17 NM from Canberra when the GPWS warning was received, the DME equipment may have been faulty.

SIGNIFICANT FACTORS

1. While the pilot was attempting to maintain a maximum speed descent for air traffic control sequencing purposes, the aircraft descended below the lower limit of controlled airspace.
2. The combination of the aircrafts rate of descent and rising terrain placed the aircraft within the GPWS warning envelope.