1. The aircraft maintenance organisation had not replaced deteriorated parts with improved parts as suggested by the aircraft manufacturer.
2. A cable guard had deteriorated to the extent that it failed and resulted in high control forces in the lateral control system.
3. The operating crew were not aware of the high control inputs required to overcome the load limiter in the lateral control system.

The deteriorated condition of the plastic cable guards, and the use of tape to effect a "repair", suggests that the manufacturer's advice regarding replacement of the guards had not been heeded during major maintenance inspections.

It is likely that, when the plastic cable guard failed, a piece or pieces of plastic lodged in the left side cable run aileron control pulley, restricting the cable movement in one direction. The debris probably dislodged when the aircraft was at about 400 ft on final approach.

Pilot information

The helicopter pilot held a commercial pilot licence, a valid class 1 medical certificate and a NVFR rating. He was endorsed to fly Hughes 500 helicopters (also known as Hughes 369HS) and was current at night flying. He held a helicopter float endorsement and had successfully undergone helicopter underwater escape training on 30 April 1995.

At the time of the accident the pilot had a total flying experience of 8,462 hours, of which 7,882 were in helicopters, including 1,408 hours in Hughes 500 helicopters. He had flown a total of 545 hours at night, and his total instrument flight time was 10 hours. A biennial flight review had been conducted on 8 March 1996, and his most recent company flight check was conducted on 14 January 1997. In the four years prior to the accident, the pilot had flown in excess of 800 marine pilot transfers at night. Most of those had been in Hughes 500 helicopters. The pilot described the helicopter landing sites on the sister ships as more than adequate for a Hughes 500 helicopter, which had a main rotor diameter of 8 m.

The pilot had been rostered for duty in accordance with an exemption against Civil Aviation Order (CAO) 48 - Duty Times which applied to company pilots engaged solely in marine pilot transfer operations. At the time of the accident, the helicopter pilot had been on call solely for marine pilot transfers for the previous two days, following two days off. After awaking at 0730 on 25 February 1997, he flew the first marine pilot transfer for the day between 1830 and 2030; the flight time was about 0.4 hours. He then slept, before departing Gladstone at about 0058 on the accident flight. The pilot reported that he had not engaged in any strenuous activities during the rostered duty period and had flown only 0.7 hours in the 24 hours before the accident.

In recounting the accident, the pilot expressed the view that conditions on the night of the accident were such that no part of the flight would have been considered difficult for an experienced marine transfer helicopter pilot in a Hughes 500 helicopter. He reported that the flight to the ship was normal. After the passenger had boarded the helicopter, the pilot checked for obstructions in the intended direction of taxi and takeoff and noted the crane jib. He planned to depart to the north-east after clearing the left side of the ship and noted that two ships positioned to the north, as well as the lights of Gladstone to the west, would give him a good visual reference.

The pilot's position in the left front seat of the helicopter provided an excellent view forward across the deck, to the left and above, and above to the right. His view immediately to the right would have been slightly restricted by the passenger.

The pilot reported that he took-off and manoeuvred the helicopter into a low hover, then taxied across the hatch towards the left side of the ship. He yawed the helicopter slightly right in anticipation of weathercocking as the helicopter cleared the ship to the left side. The helicopter weathercocked as expected. The pilot said that he then stabilised the aircraft alongside the ship briefly, maintaining a constant altitude and keeping pace with the ship. This was in accordance with company procedures for departing from ships with obstructions. With the ship to his right and the lights of two other ships in the forward left quarter of his field of view, he reported that he established a zero-bank/zero-yaw attitude in preparation for transferring to flight by sole reference to the cockpit instruments. To be sure the helicopter would move away from the ship, he yawed 10 to 15 degrees left, then simultaneously increased power and moved the cyclic control forward to accelerate and climb away. Within a few seconds of initiating this sequence, he felt a jolt and the helicopter pitched nose-up and rolled to the right. Despite flight control inputs, he was unable to counteract the roll to the right. The helicopter then struck the water. The pilot believes that the ship's crane was swung into the helicopter's rotor arc as he took off from the ship.

Aircraft information

The Hughes 369S is equipped with an articulated main rotor system, which permits the rotor blades to feather (change pitch angle), flap (move up and down vertically) and to lead and lag in the plane of rotation. The blades also have washout to equalise lift across the blade. When the blades are rotating, aerodynamic and centrifugal forces act on the rotor disc. These forces are finely balanced to keep the rotor disc stable and acting in the desired manner. If a critical component, such as a rotor blade, is damaged, the rotor disc is likely to become immediately unstable and its action unpredictable. During the investigation the pilot supplied a report, which proposed a mathematical model to verify his evidence. The report's author was not qualified as a helicopter aerodynamicist or as an accident investigator.

The helicopter was loaded within its approved centre-of-gravity and gross weight limits at the time of the accident.

The approved flight manual for the Hughes 369 states that controllability during hovering downwind, and both sideward and rearward flight, has been demonstrated to be adequate in winds up to 20 kts. The Hughes 369 also has a reputation for being fully controllable in much stronger crosswind and tailwind conditions.

Wreckage and survivability information

Examination of the wreckage did not reveal any fault that might have contributed to the accident. All flight control system damage was typical of main rotor and drive train sudden stoppage. The main rotor head assembly had suffered extensive damage, indicative of main rotor blade contact with a solid object while the rotors were being driven. The four main rotor blades and their grips were torn off the rotor head at the strap pack as a result of the blades impacting the pulley block, and the helicopter's subsequent contact with the sea. Both tail rotor blades, the tail rotor gearbox, and part of the tail rotor drive shaft had separated from the aircraft when the aft portion of the tail boom fractured during the impact sequence. Those items were lost at sea.

The engine-to-transmission drive shaft suffered an overload fracture typically caused by sudden stoppage forces. Smearing of the metal fracture surfaces on this drive shaft indicated that it had continued to rotate after the fracture occurred. An in-depth examination of the fuel system was not considered necessary, due to the physical evidence that the engine was performing at a high power setting when the main rotor strikes occurred.

The main damage to the fuselage occurred on the right side, where the fuselage skin exhibited extensive lateral/inward crushing deformation as a result of impact from one or more main rotor blades, as well as from water impact. Both front seat pans were crushed downward, consistent with the high g-loading experienced by both occupants when the helicopter impacted the sea. Both forward cabin doors separated from the aircraft. Most of the fibreglass engine intake fairing was missing after the accident.

Both flight attitude indicators fitted to the helicopter had recently been overhauled. Notwithstanding the extent of impact and salt-water damage, no fault was found with the instruments.

The carrying capacity of the crane was 25 tonnes, with a maximum outreach of 28 m. Marine surveyors subsequently advised that the design of the cables and the pulley block counteracted any tendency for the block to turn and twist the cables. Consequently, the block face that was struck by the rotor blades, was probably facing out to sea at the time of the accident.

The vertical face of the pulley block struck by the helicopter was approximately 1.4 m high by 1 m wide. Contact between the main rotor blades of the helicopter and the pulley block resulted in several distinct impact marks on the face of the block. Four of the marks displayed features that were consistent with contact by the main rotor blade leading edge abrasion strips and threaded tip weights. The sequence of the blade strikes could not be established. However, the presence of the tip weight impact marks on the block face indicated that the main rotor blades had not contacted any solid object before hitting the block.

Multiple scratch marks were found on the opposite face of the block to the main rotor blade strike marks. Those marks were considered to have been a result of contact with the tail rotor, the tail boom or the stabilisers.

No evidence was found of rotor strike marks on the hook, the swivel, the chain, or the cables above the pulley block, nor were any marks found on the upper or lower edges of the block, or on the narrow vertical edges. However, the narrow vertical edge of the block nearest the trailing end of the main rotor strike marks showed evidence of white paint and fibreglass consistent with the engine intake fairing contacting the block. Wreckage evidence indicated that the main rotor blades probably dislodged the intake fairing. Other fibreglass items attached to the airframe were relatively undamaged and showed no evidence of contacting the block.

The helicopter manufacturer reported that, "Once the first main rotor blade struck the pulley block, all blades would have been affected by the tremendous forces generated. The sudden stoppage forces imparted and damage done to the main rotor system would have resulted in severe main rotor imbalance and caused the blades to go divergent in the lead/lag axis and possibly in the flapping and feathering axis as well; in other words the blades would no longer 'fly' as you would expect normal rotor blades to. The drive train and fuselage would also have been affected by these same forces. Engineering and/or mathematical modelling of the accident scenario then becomes a wild guess, as the performance and/or actions of the fuselage, main rotor system (to include main rotor blades) and drive train are no longer predictable."

The helicopter was fitted with utility floats. A life raft was stowed in the rear passenger compartment. Both the helicopter pilot and the marine pilot wore life vests. Both occupants also wore full harness seat restraints and remained strapped in their seats during the accident sequence. Examination of the wreckage indicated that a main rotor blade had penetrated the cabin area on the right side of the aircraft, fatally injuring the passenger. The pilot was able to escape unaided from the helicopter after the accident.

Other information

Police spoke to the ship's captain by telephone two hours after the accident. The captain reported that the helicopter was almost out of the confines of the ship when it started turning left and the main rotors then struck the crane hook which was hanging in the air.

The ship's crew subsequently reported that the crane operator had turned the jib to the left side of the ship and raised both the jib and the hook before vacating the crane tower and standing on the deck for the landing and take-off of the helicopter. They said that the helicopter initially rose into a hover about 1 m above deck level, where it paused briefly before accelerating across the deck, climbing and turning left at the same time. They reported seeing the helicopter then collide with the pulley block and begin to rotate, before the tail rotor also struck the block. The helicopter then fell into the sea. Shortly thereafter, crewmembers saw the helicopter floating inverted about 15 m from the left side of the ship.

The company operations manual (page D8.10) stated:

"During each take-off, when established in the hover over the deck, a check should be made of power available, centre of gravity and temperatures and pressures before moving clear of the landing area.

At night a climb to 500 ft is to be completed before any substantial turns are made. All turns are to be made at the standard rate. Steep turns are not to be carried out".

CAO part 95, section 95.7.3: Exemption of Certain Helicopters from Compliance with Provisions of Sub-regulation 174B (2) of the Civil Aviation Regulations provides for special requirements for helicopters engaged in charter operations at night for the purpose of marine pilot transfers to/from ships. No evidence was found that the operator or the pilot had not complied with the requirements of CAO 95.7.3.

The factors contributing to the accident could not be determined with certainty.

Although there were main rotor blade strike marks on the seaward face of the pulley block, there was no strike damage on the leading or trailing edges of the block. This evidence appeared consistent with the main rotor hub being forward of a line perpendicular to the seaward block face, and the main rotor blade leading edges impacting the block face at an angle of less than 90 degrees. The ship-facing block face had no evidence of main rotor blade strike damage, however it had sustained impact damage from the tail rotor assembly. Information provided by the pilot and the ship's witnesses indicated that the main rotor disc struck the block before the helicopter began rotating. The evidence was consistent with the main rotor blades hitting the seaward block face and, as a result of the impact, the helicopter rotated clockwise, and the tail rotor assembly then struck the ship-facing block face.

Once one of the main rotor blades was damaged, it was probable that the main rotor system became unstable and its subsequent motion unpredictable due to the variables involved. A complete understanding of the relative movement between the helicopter and the pulley block requires accurate data on the motion of each object with respect to a known frame of reference. Both the ship and the helicopter were moving independently of each other. Although the speed and heading of the ship were reported, as was the sea swell, none of the information was calibrated or recorded to sufficiently fine tolerances. While the extent of the rolling and/or pitching motion of the ship was probably negligible, any movement would affect the variables in any calculation. As a consequence, the movements of the pulley block, as a result of the ship's rolling or pitching, or the influence of wind, could not be precisely determined.

These considerations, associated with the expected unstable behaviour of damaged main rotor blades, precluded an accurate assessment of the relative motion between the rotor disc and the pulley block. As a result, the strike marks on the pulley block provided insufficient physical information to reconcile the significant differences between the accounts of the ship's crew and that of the pilot. Determination of the attitude and position of the helicopter at the moment of collision was therefore not possible and any attempt to do so would be, at best, speculative. The investigation considered that, although the mathematical model proposed by the pilot and his advisors was possible, there was insufficient physical evidence to preclude other scenarios.

The helicopter was engaged in the ship-to-shore transfer of a marine pilot at night, and was operating in accordance with the night visual flight rules (NVFR). The ship was reported to be steaming at 14.5 kts, steering 044 degrees M. Weather conditions at the time were reported to be fine, with visibility of 4-5 NM. The night was also reported to have been very dark, with some haze. The sea was almost calm, with a swell ranging between 0.25 m to 0.5 m. The wind was about 5 kts from the ENE and the temperature was about 27 degrees Celsius. The moon was waning, and its bearing relative to the accident site was 027 degrees M, at an elevation of 60 degrees above the horizon.

The deck landing area measured approximately 25 m fore and aft, and 20 m across the ship, which was 23 m wide. The landing surface consisted of steel cargo hatch covers. Sea containers were stacked to a height of 5.2 m, immediately forward of the landing area.

A potential obstacle to the operation was a crane immediately aft of the landing area. In its stowed position, the crane jib would normally be aligned along the fore/aft axis of the ship, above and parallel to the landing area surface. The ship's crew reported that the marine pilot requested the crane's jib be turned 90 degrees to the left side of the ship, and elevated to its upper limit. The pulley block of the crane's hook assembly weighed an estimated 1.1 tonnes and was painted with yellow and black stripes. A lifting hook hung below the block. The crane assembly and the deck landing area were floodlit. A light on the jib illuminated the pulley block. The height of the pulley block at the time of the occurrence could not be positively established. If the crane was hoisted as reported by the ship's crew, its height above the ship's deck was about 20 m. Had the crane been swung as reported by the pilot, the height of the pulley may have been lower than 20 m. The ship's crew reported that the crane operator had vacated the crane tower and was on the deck for the arrival and departure of the helicopter. The helicopter approached the ship from the right side and landed on the forward right hatch cover, facing towards the left side of the ship. The marine pilot boarded the helicopter through the right front door and occupied the right seat. During the subsequent take-off, there was a collision between the helicopter and the pulley block, and the helicopter fell into the sea, where it floated inverted, supported by the buoyancy of its utility floats. Small pieces of rotor blade debris were found on the ship's deck. A fisherman heard the ship's master report the accident on the marine radio frequency and, after searching for about 25 minutes, located the wrecked helicopter. The passenger was fatally injured, and the pilot sustained minor injuries.

This was the first time the ship's master had accepted a helicopter marine pilot transfer from this ship. He was familiar with helicopter operations onto larger ships. He was hesitant to agree to a helicopter transfer until the ship's agent and the marine pilot convinced him that the size of the proposed landing site was adequate. This was also the first occasion that the helicopter pilot had conducted a marine pilot transfer with this ship. However, he had previously conducted two marine pilot transfers at night onto the sister ship, the most recent being on 11 February 1997.

Pilot and eyewitness descriptions of events immediately prior to the accident suggest that the engine lost a substantial degree of power, but did not completely fail.

It is likely that the engine power loss was gradual. This was evidenced by the pilot's report of the aircraft's increasing nose-down tendency and its extended flight profile. Examination of the engine indicated that it was producing limited power at impact. There was no evidence of any pre-existing condition that may have adversely affected the serviceability of the engine.

A significant amount of water and other contaminants were found in the firewall fuel filter.

Inspection and testing indicated that the water had been in the filter a significant length of time. The water in the carburettor bowl and the fuel pump was clear and the lack of corrosion or other damage inside the carburettor and the pump indicated that the water had only recently been introduced to those components, possibly during the post-accident firefighting operations. No factors, other than water contamination and the deterioration of the fuel filter, were disclosed during the investigation. Consequently, it is considered that the deterioration of the filter most likely contributed to the engine's loss of power.

There was no evidence to indicate that the fuel filter had been removed and inspected during the periodic inspection conducted for the initial issue of an Australian aircraft maintenance release or at any subsequent time prior to the occurrence. The aircraft maintenance schedule elected by the Certificate of Registration holder would not have required a further inspection of the filter until 100 flight hours had been completed after the initial inspection.

During the investigation, it became apparent that there was a general misunderstanding of the intent of certain aspects of the CASA maintenance schedule contained in CAR42B Schedule 5.

Two aircraft were being used to suppress bushfires in the Perth basin. At approximately 1027 WST, the aircraft were called to a fire near Maddington, an outer suburb of Perth approximately 6 NM south-east of Perth airport. The pilot of VH-JTD reported that no problems were identified during the pre-take-off checks.

Both aircraft departed at approximately 1035 towards the south. Air traffic controllers in the tower noted that JTD, the accident aircraft, flew a shallower departure profile than the other aircraft. The pilot later reported that the profile resulted from his operating technique and not from any problem associated with the aircraft.

Because the bushfire was relatively close to the airport, both aircraft climbed to only 1,200 ft before descending to conduct the inspection circuit at about 1,000 ft. While communicating with the bushfire ground controllers, the pilots positioned their aircraft in a left circuit pattern to reduce the time during which the aircraft were climbing out over populated areas. The pilots then approached the fire from the south-east. The first aircraft dropped its load of fire retardant without incident.

The pilot of JTD reported that he set up the approach slightly tighter and steeper than usual due to the proximity of housing. Just prior to the base turn, he selected full flap and set the propeller to a fine-pitch setting.

During the base turn, the pilot reported that he needed more elevator control back pressure than usual to maintain the aircraft's nose attitude. The pilot initially thought that the aircraft's trim setting needed adjusting, until the nose dropped further. The pilot levelled the wings and fully opened the throttle: this appeared to have little or no effect and he decided to immediately dump the load of retardant. He could not recall any engine instrument indications. Because the dump system had been set up for a partial dump only, he decided to use the manual lever rather than the electric switch to dump all the fire retardant. Following the loss of engine power, the pilot turned on the emergency fuel pump as required in the emergency checklist, but did not apply carburettor heat.

The pilot reported that the aircraft appeared to travel further than expected during the descent. He flew the aircraft close to the stall at approximately 60 kts during the forced landing, using rudder to control wing drop. Having passed over several streets, the pilot decided to attempt to land the aircraft on a small street, parallel and close to the aircraft's flight path.

The aircraft severed a set of powerlines at the northern end of the street before its right wing struck and broke a power pole. The aircraft then veered right, glanced off the roof of an occupied house, and embedded itself into the next house, which was unoccupied at the time.

Once the aircraft had come to rest, the pilot switched off the fuel and electrical systems. There was no fire. He then exited the aircraft and signalled to the other pilot that he was okay. The pilot was wearing a four-point harness and a flying helmet, both of which were undamaged. However, the pilot received minor injuries to his left leg during the impact sequence. The nature of the impact sequence appeared to have progressively slowed the aircraft, thereby lessening the peak forces suffered by the pilot.

When the other pilot saw that JTD had crashed, he advised air traffic services of the accident and gave directions to enable emergency services to locate the accident site. Emergency services arrived about 5 minutes later. The pilot switched off the operating emergency locator transmitter.

Wreckage examination

The aircraft's forward fuselage was substantially damaged and the main landing gear had collapsed. The wings remained attached to the fuselage although powerlines had severed 2 m of the left wing. The cabin was penetrated by two pieces of wood from the house roof trusses. The engine was torn from its mounts, moving 1 m forward and rotating approximately 90 degrees to the left.

The on-site examination of the engine revealed no abnormality. The carburettor was firmly attached and all four butterfly valves were fully open although their positioning may have resulted from the impact. The propeller remained attached to the engine with all blades bent backwards. The general appearance of the blades indicated that the propeller was rotating at low power at impact.

Meteorological information

The Bureau of Meteorology reported that there was 2 octas of cumulo-form cloud at 4,500 ft. The temperature was approximately 35 degrees C, dewpoint was 16.5 degrees C, and there was a relative humidity of 38% at Perth airport. There was a light wind from the north-east.

Fuel system

The fuel system included two wing tanks connected to a header tank. Fuel from the header tank was drawn by the engine-driven fuel pump and fed through a firewall-mounted filter before entering the carburettor.

The operator reported that the aircraft departed with a fuel load of approximately 700 L and that approximately 40 L of fuel was drained from the left tank soon after the accident. It was estimated that about another 20 L was spilled during the aircraft wreckage recovery. The right wing tank contained no fuel when it was inspected after the accident; however, the fuel would have drained from the fuel lines that were damaged during the collision sequence.

A sample of fuel was drained from the left wing tank and analysed. The analysis confirmed the fuel as Avgas 100/130. Apart from an extremely low Reid Vapour Pressure, it complied with the fuel specification requirements. The fuel sample was taken almost 24 hours after the accident. The fuel manufacturer reported that due to the hot weather in Perth at the time, the lighter fuel fractions may have evaporated from the ruptured tank, resulting in the low vapour pressure.

A small amount of clear water was found in the fuel sample drained from the fuel pump.

Fuel lines to and from the fuel filter, mounted on the firewall, were severed at the filter. The fuel filter did not have an external or cockpit warning to indicate when the filter was blocked and operating in bypass. The aircraft's flight manual noted that the fuel pressure indication dropped when the filter was blocked, but the pilot reported that he did not observe such an indication. A mixture consisting of 50% fuel and 50% black-coloured water was drained from the filter. Chemical analysis of the water found the presence of dissolved impurities.

Examination of the filter element found the lower part saturated with water and heavily contaminated with rust particles to the point that the lower half of the filter was blocked. On the fuel inlet side, there was a significant accumulation of large, flat rust pieces. The lower half of the steel coil inside the element was severely corroded. Analysis indicated that the large rust pieces probably originated from a fuel storage tank. Specialist examination of the filter indicated that the water and subsequent corrosion had been present for an extended period.

Approximately 250 mL of fuel and a similar quantity of water were drained from the carburettor bowl. Both samples were free of sediment. The carburettor was a standard float type, where the pressure difference between the chamber and venturi regulated the amount of fuel drawn. The carburettor was opened and inspected. No sediments, or evidence that water had been present for an extended period, were found. The examination of the carburettor found no defect that may have adversely affected its operation.

The aircraft was manufactured in 1993 and operated in the USA before its purchase by the Australian owner in 1996. The aircraft Certificate of Registration holder made a valid election to use the Civil Aviation Safety Authority (CASA) maintenance schedule, CAR42B Schedule 5, as the aircraft maintenance schedule and the aircraft had undergone a periodic maintenance inspection prior to the issue of an Australian maintenance release. A subsequent periodic maintenance inspection would then be required at the completion of 100 flying hours or 12 calendar months.

An option available to the aircraft Certificate of Registration holder was to elect to use the aircraft manufacturer's maintenance schedule, which would require that the fuel filter be inspected every 50 (+/- 5) flight hours.

On 5 February 1999, CASA informed the Bureau that the aircraft maintenance schedule, CAR 42B Schedule 5 Part 1, required that a fuel sample drain of aircraft fuel filters be conducted each day the aircraft was flown, except when not applicable to the aircraft. In the case of the PZLM18A aircraft, CASA stated that the engine firewall fuel filter would not be required to have a fuel sample drain carried out during the daily inspection.

The Civil Aviation Safety Authority confirmed that the firewall fuel filter should have been inspected during the initial periodic inspection for the issue of the aircraft maintenance release.

At the tiine of the accident, the aircraft had accumulated approximately 53 flying hours since the completion of the periodic inspection. The investigation could find no evidence that the firewall fuel filter had been inspected at any time prior to or during flying operations in Australia.