The pilot reported that the helicopter suffered an engine power loss and that he conducted an autorotation into trees. Damage was caused to the tail rotor assembly and tail boom as well as a main rotor blade. The engine continued to run after the aircraft touched down.

The cause of the engine power loss was not determined.

During the pre-flight inspection the pilot noticed that the left aileron was buckled. Further inspection revealed diagonal wrinkling of the lower surface of the left wing, the leading edge of the left wing outboard of the engine nacelle, and evidence of movement in the left-wing attachment area.

The pilot stated that when he flew the aircraft on the previous day it had handled normally. The weather conditions on that day were such that at cruise the flight was smooth with light turbulence during the descent.

An analysis of the damaged aileron concluded that it had failed in static overload. An engineering assessment of the damaged to the aircraft concluded that the most likely scenario to cause the damage was one involving high speeds and gusts. These conditions were not reported by the pilot.

VH-TJH was departing from runway 27. The clearance was for a radar one departure, initial heading 270 degrees and a requirement to maintain an altitude of 3000 feet. After take-off the crew called on Melbourne departure control frequency at 2029.51, stating they were turning right onto 270 degrees, climbing to 3000 feet and passing 2000 feet. In response they were told they were identified.

A short time later, at 2030.55 they were told to turn right onto a heading of 310 degrees to intercept the 285 radial from the Melbourne VOR. At 2032.25 the crew of VH-TJH told the departures controller they required an immediate climb. The controller immediately cleared them to climb from 3000 feet to flight level 350.

The aircraft track had taken the aircraft towards rapidly rising terrain west-north-west of Melbourne. The higher points along this track are between two and three thousand feet. The aircraft Ground Proximity Warning System operated to give a warning of 'Terrain, pull up.' The crew responded by quickly commencing a climb, and this caused the GPWS to cease operating.

At that time of night, due to decreased traffic conditions, the approach and departures air traffic control positions were combined so that they could be operated by one controller. This also included the radar advisory service, (RAS). For the departure of VH-TJH the controller put an altitude restriction on the aircraft due to an aircraft inbound from the west. Other traffic was handled during this period and a call on the RAS frequency also occurred.

A further call was received from Moorabbin Tower to arrange a clearance for an aircraft taxying at Moorabbin. The situation regarding the aircraft approaching from the west was resolved by the time VH-TJH first called after becoming airborne. The altitude restriction of 3000 feet could have been removed then but was overlooked. While handling other matters the controller omitted to remove the restriction as the aircraft tracked outbound towards rising terrain. The GPWS warning was genuine.

The crew of VH-TJH did not take any early action to request a climb despite the fact they were tracking towards rising terrain while maintaining 3000 feet.

Significant Factors

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

1. The Controller omitted to remove the altitude restriction placed on VH-TJH, in a timely manner. The precise reasons for this happening could not be determined.

2. While tracking towards rising terrain at an altitude of 3000 feet, the crew of VH-TJH took no action to request a climb prior to activation of a GPWS alert.

The accident occurred during glider towing operations. After the third landing for the morning, the aircraft was being turned to back track on runway 22. The pilot stated that when power was applied to assist the turn being made with rudder and right brake, the aircraft immediately pitched 45 degrees nose down and the propeller was embedded in the earth strip. The wind was estimated to be 18-20 knots gusting to 25 knots from 230-240 degrees magnetic. The pilot stated that he considered inadvertent relaxation of back pressure on the control column, the power application, and possible excessive brake application causing sudden locking of the right-hand wheel were contributing factors.

Two days prior to the accident the pilot checked the aircraft and found fuel leaking from the right wing near the drain valve. He defueled the right tank to 30 litres to minimise fuel loss if the leak continued. Suspecting that the leak was from the drain valve, he reseated the valve and cleaned the fuel stains from the aircraft. The aircraft was checked again the next day. There was no evidence of fuel leakage, so the right tank was filled to 60 litres. On the following day there was again no evidence of fuel leakage, so the pilot assumed the leak had been from the drain valve.

The purpose of the flight was to transport three passengers from Mittagong to Cudal and return. The aircraft left Mittagong with 120 litres of fuel in the left tank and 60 litres in the right tank. During taxy prior to departure there was no evidence of fuel leakage and a post flight inspection after arrival at Cudal again revealed no evidence of leakage, so the tanks were refilled to 90 litres per side for the return flight.

On the return flight the weather at the destination deteriorated necessitating a diversion. While planning the diversion the pilot noted that the right tank fuel quantity gauge had dropped significantly since he last checked fuel quantities a short time previously. A short time later the pilot noted a moderate smell of fuel in the cabin. He suspected a serious fuel leak and decided to make a precautionary landing.

The pilot informed his passengers of the situation and his intentions. He then selected what appeared to be a suitable area and made three inspection runs over it. On short final approach, at about 20 feet off the ground, the pilot noticed that the selected area sloped sharply upwards. He flared the aircraft but could not arrest the descent. The aircraft touched down heavily and bounced. The pilot relanded and brought the aircraft safely to a stop. An inspection revealed substantial firewall damage as a result of the heavy landing.

After the aircraft had been recovered to a maintenance base an examination revealed a fuel stain down the right side of the fuselage just behind the cabin door. There was also heavy fuel staining on the root rib around the cutout for the right fuel tank aft supply line spigot. Further investigation revealed the right tank aft outlet spigot had fractured through 360 degrees about five millimetres inboard of the root rib cutout. The spigot is about 60 millimetres long and accommodates the aluminium supply line which is secured by a clamp at the inboard end of the spigot. The inboard end of the spigot complete with its supply line and clamp remained with the fuselage. The heavy fuel stain in the area was evidence that the spigot had been cracked for some time.

Significant Factors

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

1. A broken fuel tank supply line spigot permitting fuel leakage from the right fuel tank.

2. The leakage resulted in fuel vapour entering the cabin which led to the pilot making a decision to make a precautionary landing.

3. The area chosen for the precautionary landing had a significant upward slope which was not perceptible from the air. This slope was not noted by the pilot until very late on final approach. The result was that the pilot was unable to flare the aircraft sufficiently to prevent a heavy touchdown.

The pilot, who had limited experience, stated that during the take-off run the aircraft started to drift to the left of the strip due to a gusting crosswind. The aircraft struck a runway marker and the take-off was aborted. The aircraft then ran off the runway and struck a pile of dirt.

The check and training pilot was demonstrating a single engine approach to a simulated helideck in crosswind conditions and towards the setting sun.

As the aircraft entered the flare for landing the pilot misjudged the approach and adopted a higher-than-normal nose attitude. During the manipulation of the controls to complete the landing the pilot lowered the collective control whilst holding aft cyclic control. This resulted in one main rotor blade striking the base of the vertical fin.

The pilot felt the contact but assumed it was a tail skid strike. The skid was inspected, and when no damage was found, the flight was continued.

The blade strike was discovered after the flight was completed.

The possibility of a blade strike on the tail during single engine landings is a known problem and is covered in Sikorsky Aircraft Customer Service Notice 76-62. The pilot was aware of the contents of this notice.

During a maximum power take-off, the engine exceeded Exhaust Gas Temperature (EGT) limits. The outside air temperature (OAT) was reported to be 31 degrees C.

The engine was shut down and the aircraft returned to land. The engine was removed and a subsequent test cell run showed that it had a low EGT margin. This would have resulted in the observed EGT exceedance under high OAT conditions.

The operator has carried out a maximum power assurance run on all engines and has instituted a continual EGT monitoring program to ensure that adequate EGT margins are available.

When passing through 1,000 feet after take-off from Essendon runway 26 a loud bang was heard, and the left engine was noted to have failed. Concurrent with the left engine failure the right engine driven hydraulic pump failed. The pilot initiated a return to Essendon. Due to the loss of both hydraulic systems the landing gears were extended using the manual system. The manual system does not have any provision for closing the landing gear doors therefore they remained in the down position and created considerable drag. The aircraft would not maintain height on one engine. Rather than approach and land on the shorter Essendon runways the pilot diverted to Melbourne Airport for a successful asymmetric landing.

Investigation

1. Overhaul agency investigation

The left engine was removed and transported to a manufacturer approved overhaul agency for disassembly and investigation. The engine internal rotating and stationary components were found to be severely damaged.

Assessment by the overhaul agency suggested that the primary failure was a first or second stage low pressure (LP) turbine blade. It was suggested that liberation of this blade resulted in extensive foreign object damage and imbalance of the rotating assemblies with consequent blade, shroud, disc and impellor distress.

It was determined that at the time of failure the engine had completed 6,858 hours and 4,114 cycles since new. Repairs, during which some turbine blades were replaced, were carried out in August 1989, 2,728 hours and 1604 cycles prior to this failure, and in October 1991 1,300 hours and 807 cycles prior to this failure.

2. Materials Evaluation Facility investigation

The damaged rotating and stationary components were forwarded to the Civil Aviation Safety Authority's Materials Evaluation Facility (MEF) for further investigation.

This metallurgical evaluation determined that, among other damage:

• a total of 84 1st stage low pressure turbine (LPT) blades had been fractured. The fractures were consistent with the application of excessive stress associated with interference from other turbine module components with the tips and trailing edges of the 1st stage blades while the turbine was operating.
• a total of 84 2nd stage LPT blades had fractured with the application of excessive stress. Unlike the damage to the tips and trailing edges as noted on the 1st stage blades, these 2nd stage blades were damaged on their leading edge and the blade platforms, similarly while the turbine was operating.
• the 2nd stage LPT stator assembly had been extensively damaged. The leading and trailing edges of the nozzle guide vanes had been damaged, and the shroud was fractured.

In analysing the sequence of the failure, the MEF considered:

1. the contact evidence of the aft face of the 2nd stage low pressure turbine stator shroud and the leading edges of the 2nd stage rotor blades,
2. the contact evidence of the 1st stage low pressure turbine blades with the forward faces of the 2nd stage low pressure turbine stator shroud fracture surfaces,
3. the deposits of metallic aluminium on the 2nd stage low pressure turbine stator shroud fracture surfaces,
4. the location and extent of the fracture in the 2nd stage low pressure turbine shroud, and(e) the mechanism of the fracture in the 2nd stage low pressure turbine stator shroud.

The MEF considered that the fracture of the 2nd stage LPT stator assembly was the first event in the engine failure sequence. Evidence of fatigue cracking was found at locations distributed around the circumference of the shroud. Fatigue cracking had initiated at the outer surface of the stator casing at a location that coincided with a marked change in the geometry of the shroud. Thermal expansion of the nozzle guide vanes with each engine thermal cycle would result in the imposition of alternate bending loads on the stator casing - the combination of vane expansion and restraint of the outer casing at its forward and aft edges would lead to the development of a tensile stress state at the outer surface.

3. Manufacturers investigation

The manufacturer disagreed with the MEF findings. The manufacturers analysis of the 2nd stage low pressure turbine stator shroud found no evidence of any fatigue propagation anywhere on the fracture surface. Their analysis of the secondary damage to the engine indicated it was consistent with a second stage low pressure turbine failure as evidenced by the associated damage compared to other recorded failures in the fleet. They also advise that there has not been any reported second stage low pressure stator assembly fatigue failures during more than 15 million hours of operation of this engine type.

The manufacturer considered that;

1. the heavy rub on the leading edges of the 2nd stage low pressure turbine blades was secondary to the impact damage on the leading edges of the blades,
2. the heavy rub on the aft side of the 2nd stage low pressure turbine stator had a fresh appearance and occurred after the impact damage on the vanes, and
3. the fatigue zone identified by the MEF was a shear lip with extensive smearing indicating an overload failure.

Based on their own material analysis the manufacturer stated that:

'the second stage low pressure turbine stator outer casing was not the primary failure... but rather (the primary failure was) an air foil separation of a second stage low pressure turbine blade'.

and, in relation to an air foil separation;

'Unfortunately, the exact cause of the second stage blade separations was not determined. The separation of these blades is believed to be related to the non-discernible grain areas detected in the air foils of one or more blades'.

The manufacturer used a micro etch technique to examine five random blades from this engine, finding one that had a non-discernible grain area.

The manufacturer has issued a service bulletin requiring the removal from service of a series of blades found to have non-discernible grain areas. The one blade from this engine that was identified as having a non-discernible grain area was from a series that were not listed for removal by the service bulletin.

4. Hydraulic pump failure

The failed hydraulic pump was removed and sent to an overseas overhaul agency for assessment and repair. Although it was a requirement, the overhaul agency did not supply a report on the cause of the failure.

The aircraft is fitted with two separate hydraulic systems. The fact that the right engine driven hydraulic pump failed at the same time as the left engine failed is considered to be a coincidence.

The certification requirements for this type of aircraft do not address the failure of multiple systems as experienced on this flight. Accordingly, an engine failure, coupled with a total hydraulic failure leading to a manual extension of the landing gear, will leave the aircraft with high drag profile and low power that would be beyond the experience of most pilots operating this class of aircraft. The aircraft flight manual did not address this problem, nor was it required to, because the event is well outside the certification requirements, and it is unlikely that flight testing was carried out to establish performance parameters that would lead to flight manual recommendations.

5. Bureau determination

Based on the data made available to the investigation, the Bureau was unable to positively determine what initiated the failure of this engine. The Bureau is not convinced that an air foil separation due to non-discernible grain areas was the primary cause. However, if that was the primary cause, and if the manufacturer's service bulletin is properly addressing the removal from service of discrepant blades, then the incidence of engine failures due to this factor should rapidly decrease.

It should be noted however, that the one blade from this engine that was identified as having a non-discernible grain area, was from a series that were not listed for removal by the service bulletin.

If the failure was due to second stage low pressure turbine stator outer casing fatigue, then it is to be expected that more failures of this type could occur, with more definite evidence of fatigue able to be identified, resulting in a fleet rectification program.

VH-AWL was operating over the central business district (CBD) at 2000 feet when VH-AGS called Essendon tower at Point Ormond (PTO) requesting clearance also to operate over the CBD at 2000 feet. Due to workload the clearance was denied and VH-AGS was asked to sight VH-AWL. Later, a radar return was observed entering the control zone north of Point Ormond flying at the same level and in the opposite direction to VH-AWL. The pilot of VH-AGS was asked to sight VH-AWL which he did at a range of about one mile. There was insufficient time to alert VH-AWL. VH-AGS was about 1.5 miles inside the control zone at the time of the confliction.

The pilot of VH-AGS later claimed that he thought he was cleared to enter the control zone once he had sighted the other aircraft. He also thought he had been told the other aircraft was at 3000 feet. He had the other aircraft in sight from the time he entered the control zone until the confliction occurred. He estimated the two aircraft passed with 200 feet of vertical separation and about 200 metres of lateral separation.