Collision with terrain

Preparation for take-off

The pilot was qualified, appropriately endorsed and reported to be fit and well for the flight. The available aircraft and maintenance documentation, results of the recent 50-hourly inspection and subsequent engine ground run, and the successful 2-hour flight to El Questro indicated that there were no known pre-existing aircraft anomalies that might have contributed to the occurrence.

Although it was reported that the pilot most likely observed the El Questro Aircraft Landing Area (ALA) in June 2003, the investigation was unable to determine the means by which the pilot satisfied himself prior to arriving on 28 August 2004 that the ALA was suitable for his operation in accordance with the requirements of Civil Aviation Regulation 92. The runway width and length satisfied the guidance provided by Civil Aviation Advisory Publication (CAAP) 92-1(1) for the aircraft. While the tree that was initially struck by the aircraft may have infringed the lateral transitional slope, the pilot had indicated his acceptance of the suitability of the ALA for his operation and had landed without incident on 28 August 2004.

The results of the fuel tests, and the large number of other aircraft that had also used the Broome fuel supply without any reported problems, indicated that the quality of that fuel supply had not contributed to the development of the occurrence. The reported conduct by the pilot of the left underwing fuel drain test and observed presence of the pilot underneath the right wing tended to suggest that the pilot had completed the fuel drain and sump checks. That supported the consistent reports from all witnesses that the aircraft's engines sounded 'normal' throughout the take-off, and decreased the likelihood that water or other particulate contamination of the aircraft's fuel supply to the engines had any effect during the take-off.

The pilot's extensive and recent experience in the operation of this type of aircraft, and apparent intimate knowledge of its systems and procedures, would have reduced the likelihood of the pilot omitting to unlock the aircraft's controls prior to commencing the take-off. In any case, the ability of the pilot to fly the aircraft from the runway meant that the control column and optional rudder gust locks were unlocked at that time. Therefore, the pilot should have been able to react to the reported initially slight left bank after take-off.

The Broome engineer's report that the engines had started and run normally after the 50-hourly inspection, and the lack of any indication by the pilot during his stay at El Questro of any problems affecting the flight to El Questro indicated that the aircraft probably did not have a recurring engine starting problem. The action by the pilot to continue the engines start, taxi and take-off without interruption indicated that the pilot was unconcerned by the reported initial difficulty starting the left engine, or by the 'frequency vibration' that was reported by the passenger witnesses at the runway 32 parking area. In addition, given the history of extensive efforts by the pilot during the flight to Australia to at all times ensure the serviceability of the aircraft, it was likely that the pilot would have attended to any engine(s) start or after-start anomaly as soon as it became evident. The investigation concluded that the start and after-start passenger witness reports probably resulted from observations that were accepted as 'normal' by the pilot and also by the pilot witness who was also located at the runway 32 parking area.

Take-off

The action to conduct a rolling take-off was consistent with an attempt by the pilot to minimise the potential for damage to the propeller blades and underside of the aircraft as a result of the dislodgement from the runway surface of pebbles and other potentially damaging debris. The investigation considered whether the pilot may have attempted to lift off the runway earlier than normal in an attempt to minimise the possibility of such damage. However, this was considered unlikely given the report from the pilot witness, that the lift-off appeared to reflect other twin-engine aircraft take-offs that he had observed, and the approximation of the length of the take-off roll to that predicted by the Pilot's Operating Handbook (POH) performance charts.

The reports from the witnesses at the runway 32 parking area that there had been no smoke or fumes observed emanating from the aircraft and that nothing fell from it during the take-off, corroborated the results of the after-accident runway inspection. That, along with the identification of all control surfaces at the accident site, and the distribution of the aircraft wreckage, indicated that there had been no aircraft break-up or detachment of the aircraft's control surfaces prior to the impact with the trees and ground.

There was no evidence of any pre-existing failure or anomaly in either engine or their supporting assemblies or components that were able to be tested, that might have adversely influenced the performance of the engines during the take-off. In addition, the nature of the damage to the propellers and propeller hubs was consistent with the delivery of comparable amounts of power to both propellers from within the engines' upper operating range at the time of ground impact. There was no evidence of an enduring power failure of either engine during the take-off.

The investigation considered a number of possible factors that could have had the potential to affect the lateral control of the aircraft. Those factors included: the possibility of a lateral weight imbalance; an aerodynamic influence on the performance of the aircraft; and whether the pilot might have either intentionally or inadvertently manoeuvred the aircraft to the left of the runway immediately after lift-off.

Given that all of the baggage was observed being stowed in the aircraft cabin, there was minimal potential for the baggage distribution on take-off to have adversely affected the lateral control of the aircraft. The pilot's high and recent experience operating the aircraft minimised the possibility that he might have neglected to check the correct engine fuel selections on the three occasions (at least) required by the POH during the flight to El Questro. Similarly, had the pilot inadvertently mis-selected the engines' fuel supply, the nature of the aircraft checks meant that it was unlikely that a mis-selection would remain undetected by the pilot for the duration of the flight to El Questro. Therefore, it was considered that the potential for a lateral fuel load imbalance, and therefore aircraft weight imbalance to have developed during the flight to El Questro was minimal.

In any case, given the short taxi to runway 32 from the parking area and the rolling take-off, any disparity that may have existed in the lateral fuel distribution for the take-off was estimated to have been within 50 pounds of that for the landing at El Questro. That meant that the pilot would have been aware of the potential for any associated aircraft roll during the take-off and would have been ready to compensate. In addition, as indicated in the POH, the relatively benign nature of the ambient conditions and reported 'normal' appearance of the take-off ought to have allowed the pilot to control the aircraft even had there been a lateral fuel imbalance.

The aerodynamic factors with the potential to have affected the lateral control of the aircraft included the effects of asymmetric or 'split' flap and a wing stall. The damage to the left flap meant that, in the first instance, the investigation could not discount that split flap might have occurred during the lift-off. However, given that the right flap was confirmed retracted at ground impact, the only possibility was that the left flap might have self-extended. In that case, the aircraft would have rolled to the right after lift-off. That was not consistent with the witness reports that the aircraft banked to the left after lift-off before striking the trees to the left of the runway. On that basis, the investigation discounted that split flaps had occurred.

It was considered most likely that the experienced and proficient pilot would have raised the nose wheel at 95 knots indicated air speed (KIAS) and lifted the aircraft from the runway at 100 KIAS in accordance with the procedures recommended in the POH.

The pilot witness's description of the take-off and initial slight left bank after take-off could be construed as normal manoeuvring. In that case and given the reported lack of a very steep pitch angle after take-off, it ought to have been difficult for the pilot to have stalled the aircraft. In addition, the pilot's high experience and proficiency in the aircraft, and intimate knowledge of its systems meant that it was highly unlikely that he might have: allowed the speed to decrease to between about 88 to 93 KIAS in order for the stall warning to have activated; to have then reacted inappropriately, or not at all to that warning; to have finally not, or inappropriately, reacted to any developing mild and then increasing aerodynamic buffet; and, instead allowed a continuing reduction in airspeed until reaching about 83 KIAS whereupon the aircraft stalled. On that basis, it was concluded that an aerodynamic stall had most likely not occurred.

Given the pilot's predominant recent experience during the flight to Australia of landing and taking off from major or international aerodromes, the occurrence take-off was potentially his first from an airstrip and surface such as that at the El Questro ALA for some time. In addition, as a result of the pilot having landed the aircraft on runway 14, the take-off from runway 32 was potentially his first opportunity to have comprehended the relative proximity of the trees to the left of the runway. That comprehension was probably only possible during the conduct of the occurrence take-off and may have influenced the pilot's awareness of the need to maintain the aircraft overhead the runway centreline during the take-off.

The absence of bird or other animal remains along the runway surface, and the reported absence of any thermal activity or dust devils, indicated that the pilot most likely was not required to intentionally manoeuvre the aircraft as a result of those potential influences on the take-off. Also, the nature of the pilot's recent flying experience, the number and variety of the exotic locations visited during his flight to Australia, the pilot's age and total flying experience, and his probable desire to not compound any apprehension on the part of the passenger meant that it was most unlikely that the pilot was performing the early stages of any sort of intentional low-level manoeuvre.

The investigation could not discount that the pilot might have been momentarily distracted during the lift-off from the runway to the extent that the aircraft developed an unintended slight left bank and drift. The probable small amount of time for the pilot to have reacted once he perceived any unintended movement of the aircraft would have minimised the possibility for the pilot to have avoided impacting the trees to the left of the runway.

Summary

There was no documentary, physical or witness evidence identified during the investigation that indicated that an anomaly or failure in the aircraft or its systems contributed to the development of the occurrence. In addition, there was no evidence to indicate that the reported slight left bank after lift-off from the runway was the result of a lateral imbalance of the aircraft, an aerodynamic effect or an intentional control input by the pilot. However, the investigation was unable to determine whether the pilot might have been distracted during the lift-off by an unidentified event to the extent that he did not notice or was unable to react to any unintentional left bank and drift of the aircraft in sufficient time to prevent the aircraft impacting the trees to the left of the runway.

CONCLUSIONS

Findings

Pilot

• The pilot was qualified and appropriately endorsed for the flight.
• The pilot was reported to have 975 hours experience flying the Cessna 421 B and C models over the preceding 10-year period and at least 2,100 total flying hours.
• The pilot held a valid Class 2 medical certificate.
• The pilot was reported to be fit and well and in good spirits prior to the flight.
• There was no evidence of any pre-existing medical disease, sudden illness or incapacitation that may have affected the pilot's ability to control the aircraft.

Aircraft

• Export Certificate of Airworthiness number 3588/04 was issued for the Swiss-registered aircraft by Swiss regulatory authorities on 27 February 2004.
• The quality of the fuel that was supplied from the Broome supplier was not a factor in the occurrence.
• There was sufficient fuel onboard the aircraft to complete the flight to Broome.
• The aircraft weight and balance was estimated to be within the published limits at the time of the take-off.
• The ability of the pilot to lift the aircraft from the runway meant that the control column lock and optional rudder gust lock were unlocked at that time.
• There was no break-up of the aircraft or detachment of the aircraft's control surfaces prior to the impact with the trees and ground at the accident site.
• There was no evidence of any pre-existing failure or anomaly in either engine or their supporting assemblies or components.
• Comparable amounts of power were being delivered to both propellers from within both engines' upper operating range when the aircraft impacted the ground.

Other findings

• Witness reports indicated that, shortly after lift-off from the runway, the aircraft banked and drifted to the left slightly, before striking the trees to the side of the runway and impacting the ground.
• There was no evidence that the ambient conditions contributed to the circumstances of the occurrence.
• The pilot indicated that the El Questro ALA was suitable for his operation.
• There was no distress radio transmission by the pilot.
• Based on the available evidence, it was considered most unlikely that the pilot was performing the early stages of any sort of intentional low-level manoeuvre.
• The investigation could not discount that the pilot might have been momentarily distracted during the lift-off, resulting in the development of an unintentional slight left bank and drift of the aircraft.
• The relative proximity of the trees to the left of the runway would have adversely affected the time available for the pilot to have reacted to the development of any unintentional left bank and drift.
• The destruction of the aircraft cockpit and cabin from the combined effects of the impact forces and post-impact fire rendered the accident non-survivable.

Sequence of events¹

On 30 August 2004, shortly before 1200 Western Standard Time, the owner-pilot of a twin-engine Cessna Aircraft Company 421C Golden Eagle (C421) aircraft, registered HB-LRW, commenced his take-off from runway 32 at El Questro Aircraft Landing Area (ALA). The private flight was to Broome, where the pilot intended resuming the aircraft delivery flight from Switzerland to Perth. The available documentation indicated that the flight segments en route to Australia had all been to international or major aerodromes.

The pilot of a Cessna Aircraft Company 210 (C210) and his two passengers in the runway 32 parking area witnessed the take-off. Those witnesses reported that the C421 pilot carried out a pre-flight inspection of the aircraft prior to boarding for the take-off. During that inspection, he was observed preparing for and conducting a fuel drain check under the left wing, and to have removed some weed-like material from the right main wheel. He then loaded a small amount of personal luggage into the aircraft cabin, before he and the sole passenger boarded.

The C210 pilot witness, who reported having observed a number of twin-engine aircraft operations at another aerodrome, did not comment on the nature of the pilot's start and engines run-up checks. The passenger witnesses reported that the pilot of the C421 made a number of unsuccessful attempts to start the left engine, before reverting to starting the right engine. He then started the left engine and moved the aircraft clear of the C210 in order to conduct his engine run-up checks. The passenger witnesses reported that during those checks they heard a 'frequency vibration' as the C421 pilot manipulated the engines' controls.

The witnesses at the parking area reported that the C421 pilot taxied the aircraft onto the runway and applied power to commence a rolling take-off.² They, together with a hearing witness³ located to the north of the ALA indicated that the engines sounded 'normal' throughout the take-off. Witnesses who observed the take-off reported that the aircraft accelerated away 'briskly'. The pilot witness stated that the take-off roll and lift-off from the runway appeared similar to other twin-engine aircraft take-offs that he had observed.

The witnesses at the parking area also stated that, shortly after lift-off from the runway, the aircraft banked slightly to the left at an estimated 10 to 15 degrees angle of bank and drifted left before striking the trees along the side of the runway and impacting the ground. There was no report of any objects falling from the aircraft, or of any smoke or vapour emanating from the aircraft during the take-off. The aircraft was destroyed by the impact forces and post-impact fire. The pilot and passenger were fatally injured.

Personnel information

The 60-year-old pilot was appropriately licensed and held the relevant aircraft and other endorsements to conduct the flight. It was reported that the pilot had accumulated more than 975 hours experience flying the Cessna 421 B and C model aircraft over the preceding 10 year period and had at least 2,100 total flying hours. The pilot had flown about 50 to 60 hours in the aircraft since March 2004 and held a valid Class 2 medical certificate. He last underwent an electro cardiogram examination (ECG) in support of the revalidation of his medical certificate, on 22 September 1999. That included an annotation by the consulting doctor that the ECG was 'normal'. The pilot's family indicated that the pilot's personal logbook would have been in the aircraft at the time of the accident.

The passenger had accompanied the pilot for the majority of the flight from Switzerland to Australia, but was reported to have been a little nervous about take-offs and landings.

The pilot and passenger arrived at the El Questro ALA at about 1330 on 28 August 2004 and landed on runway 14. They were reported to have spent the next two days relaxing in the tourist resort and homestead. During that time, the pilot was observed by staff to have retired for bed by about 2200 and appeared from his room by about 0930 each day. During his stay the pilot ate regularly, drank alcohol only socially, and recounted many of his experiences during the delivery flight to staff and other guests. That did not include the discussion of any difficulty starting the aircraft engines, of any anomalies during the after-start checks and procedures, or during the flight to El Questro. The pilot was reported to be fit and well and in good spirits on the morning of the accident.

Aircraft information

General information

All of the aircraft's original maintenance documentation was reported to be on board the aircraft for the flight to Australia and was subsequently destroyed in the post-impact fire. The loss of the aircraft's maintenance documentation and historical records precluded a thorough review of the aircraft's documentation concerning compliance with applicable airworthiness directives and service bulletins. Aircraft and engine maintenance and airworthiness-related issues were reconstructed from available secondary documentation, including: pilot and other relevant party e-mails and facsimile messages, and data from international regulatory and other agencies.

E-mail correspondence from the pilot dated 16 September 2003 indicated that the aircraft was equipped with the Robertson Short Take-off and Landing (R-STOL) Kit. That kit included the following changes to the configuration of the aircraft:

• Replacement of the existing trailing-edge split flaps with R-STOL slotted flaps for use with the flaps extended 10°. The effect was to reduce the aircraft stall speed and the best single-engine climb speed.
• Introduction of a scissors-type aileron bell crank that allows symmetrical aileron droop with extension of the flaps. This had the effect of further reducing the aircraft stall speed.
• Introduction of a spring/cable flap/elevator interconnect to minimise pitch trim changes with flap extension or retraction.

In this instance, and as discussed in the wreckage examination discussion at page 13, and the asymmetric or 'split' flap discussion at page 19, the aircraft flaps were retracted. In that case, the R-STOL kit would have had no effect on the aircraft take-off performance, and the aircraft's performance would have been in accordance with a standard C421C in a flaps retracted configuration.

The aircraft's take-off weight and centre of gravity were estimated by the investigation to be within the limits published in the C421C's Pilot's Operating Handbook (POH). The POH also included a Normal Take-off Distance prediction chart. Application of the estimated aircraft take-off weight and reported ambient conditions for a normally configured C421C to that chart resulted in a predicted take-off roll that approximated the estimated take-off distance reported by the witnesses located at the runway 32 parking area.

Aircraft history

The aircraft was manufactured in the United States (US) and was US-registered until 1992, when it was exported to Switzerland. It was operated in Switzerland in the private category until purchased by the pilot in December 2003. E-mail correspondence from the pilot and dated 4 December 2003 indicated that a Swiss maintenance organisation would commence a 'new annual/200-hours check' on 10 December 2003. The available aircraft maintenance records indicated that the total aircraft flight hours were 3,192 hours 40 minutes as at 10 December 2003. That corresponded with the initial entry made by the pilot in the aircraft maintenance record after a flight in the aircraft on 2 March 2004. The last entry in the aircraft's maintenance record, that was available to the investigation, was a total of 3,233.6 aircraft flight hours at Ahmedabad, India on 9 August 2004.

A number of aircraft anomalies or maintenance requirements needed resolution while the aircraft was en-route from Switzerland to Australia. They included:

• In Malta, where an approved maintenance facility:
• replaced leaking right engine push rod tube seals
• replaced a faulty left engine vacuum pump
• identified a leaking right main landing gear oleo and disassembled the landing gear oleo and installed new packings and 'o' ring seals before reinstalling the oleo on the aircraft.
• In Cyprus, where a crack was discovered in the right engine crankcase that necessitated the replacement of that engine.
• In the United Arab Emirates and Oman, where a number of attempts were made by local engineering companies to resolve a problem with the operation of the aircraft's landing gear. That ultimately required the replacement of a selector valve and hydraulic line.

In addition, the pilot established communications with a twin-engine Cessna owners group in an effort to fault analyse ongoing problems with:

• the left and right hydraulic pumps that supplied the necessary hydraulic pressure to extend and retract the landing gear, and
• the right alternator warning light, which was reported to commence flickering after about 30 minutes flight time. That anomaly was reported to have commenced in July 2004.

The resolution or otherwise of those two anomalies was not documented by the pilot.

On arrival in Australia, a 50-hourly inspection of the aircraft was carried out by a Broome aircraft maintenance company on 26 August 2004. The total airframe hours at the time of that inspection were not noted on the inspection work sheet. At the conclusion of the inspection, a company engineer carried out a ground run of the engines with the pilot accompanying him in the right front seat. The engineer reported that during that ground run, both engines started on the first attempt and ran normally without any anomalies being noted.

Engines and propellers

The aircraft engine details were as follows:

The available right engine documentation indicated that it had been certified by a US Federal Aviation Administration-approved maintenance organisation. In addition, the documentation confirmed that the engine complied with all of the engine manufacturer's service bulletins and service letters that affected the engine up to and including 17 May 2004.

The details of the propellers, including their relationship to their respective engines, were determined from examination of the pilot's e-mail correspondence and the aircraft sales brochures, and included:

Meteorological information

A Bureau of Meteorology (BoM) post-accident assessment of the wind at the ALA at the time of the accident was that it would have been a light southerly at around 5 kts. An estimation of the ambient temperature and humidity at the ALA was not included in the BoM assessment. The BoM indicated that any difference between the BoM estimation of the prevailing wind and that reported by any witnesses could have been due to local topographical effects at the site. The BoM advised that a possible influencing factor on the aircraft's take-off could have been the presence of a dust devil⁹, but that the presence of that phenomenon would also require confirmation by any witnesses at the scene of the accident.

The witnesses at the landing area estimated that the wind affecting the runway was south-easterly at 5 to 10 kts at the time of the take-off. There were no dust devils reported in the vicinity of the runway at that time, and another pilot who was conducting charter work in the vicinity of the landing area indicated that there was minimal thermal activity.

Aerodrome and communications information

Aerodrome

The El Questro ALA, designation YEQO, is located at 16°00.5'S, 127°58.5'E and is at an elevation of 300 ft above mean sea level. The dirt runway is aligned south-east (runway 14) to north-west (runway 32) and is 1,400 m long and about 15 m wide. Windsocks are located at the northern side of the threshold to runway 14 and at the tourist resort homestead, which is located about 1 km south-east of the landing area.

The manager of the tourist resort indicated that when making bookings with the resort, visiting pilots generally included that they were 'self-fliers'. That was the case with the occurrence pilot. In addition, it was reported that the pilot telephoned the resort manager on the morning of 28 August 2004 and nominated a SARTIME¹⁰ for his arrival at El Questro. During that call, the pilot confirmed that he was comfortable with the location and details of the ALA. Other twin-engine aircraft of similar size to the C421 Golden Eagle were reported by the resort manager to have operated to, and continue to operate to, the El Questro ALA.

It was reported that in mid-March each year, just prior to the commencement of each tourist season, resort staff conducted a routine inspection of the runway and environs. As a result of those inspections, any newly growing vegetation was cleared from the runway and its surrounds, runway markers were repainted as required and other actions were undertaken by resort staff as and when required. In addition, the resort manager stated that he routinely consulted with aircraft operators who regularly fly to the resort, in order to confirm the ongoing suitability of the ALA for aircraft operations.

Communications

The charter pilot indicated that just prior to 1200, he exchanged a number of radio transmissions with the pilot of an unknown aircraft on frequency 126.7 Mhz in order to coordinate that pilot's take-off from the ALA. The charter pilot reported observing a plume of smoke from the vicinity of the ALA shortly thereafter, and that he did not hear a distress radio transmission. There was no facility at the ALA to record pilots' radio transmissions.

Wreckage information

The impact forces and post-impact fire sustained by aircraft structures in occurrences of this type can result in erroneous control position indications. In general, the position of the flight controls after impact cannot be relied upon as evidence of the aircraft's pre-impact configuration.

Overview of accident site and aircraft wreckage

The accident site was located on level ground, alongside a dry creek bed about 106 m to the left of the runway centreline and abeam a point on the runway about 888 m from the runway 32 threshold. A photograph of the general location of the accident site is at Figure 1. Groupings of rocks were located about the site and the surrounding light scrub was interspersed with isolated larger Boab and other trees.

Figure 1: General location of the accident site

A number of trees to the left of the runway were struck by the aircraft before it impacted the ground. Those trees were oriented along a line at about 15° to the left of the runway heading. Laser range equipment was used to measure the distances of those tree strikes from the aircraft wreckage. Trigonometry was then applied to the laser ranges in order to estimate the height of the strikes above ground level as follows:

• The initial tree strike was to a tree located about 66 m to the left of the runway centreline and at an estimated height of about 8.2 m (27 feet (ft)).
• The final tree that was struck prior to ground impact was located about 97 m from the runway centreline. That tree was struck at an estimated height of about 10.7 m (36 ft). The location of the left-wing tip and remnants of the left navigation light in the immediate vicinity of that tree indicated that tree strike had been by the left wing.

The aircraft impacted the ground about 33 m beyond the last tree, in a left-wing low attitude and cartwheeled counter-clockwise. The right wing and tail then struck the ground. During the impact sequence, the right engine separated from the airframe mounts and was thrown about 26 m from the main wreckage and the tail section separated from the aircraft. The main wreckage came to rest upright, with the nose of the aircraft facing south-east. All structural components and flight control surfaces were accounted for in the vicinity of the impact point. A severe post-impact fire destroyed the majority of the aircraft's fuselage, wings, tail section, cockpit and cabin, and damaged the left engine. The right engine sustained minor fire damage from a scrub fire that was started by the aircraft fire. A photograph of the aircraft wreckage is shown at Figure 2.

Figure 2: Aircraft wreckage

Wreckage examination

The investigation conducted a post-accident inspection of runway 32 from the threshold of the runway to a point on the runway abeam the ground impact point. That inspection found no evidence of any: bird or other animal remains; gouges, scrapes or other abnormal ground marks; or the presence of any detached aircraft items or components.

Very few ground impact scars or marks were able to be examined at the accident site due to them having been partially obliterated by the vehicles and personnel involved in the initial firefighting and rescue response. However, a number of rocks located in the dry creek bed had evidence of propeller impacts at substantial propeller revolutions per minute (RPM).

Both wing structures were destroyed by the fire. The right-wing forward attachment point was intact, and the aft spar was fractured in overload consistent with upward loads in excess of design limits. The left-wing structure had separated at the wing spar outboard of the engine nacelle, having also failed in overload due to the ground impact.

The fire severely damaged the aircraft fuel system. Both wing tanks were destroyed, and their associated auxiliary pumps were severely damaged. The left- and right-over-wing filler caps were secure. Damage during the impact, and the post-impact fire precluded the recovery of a fuel sample from the wreckage.

Both of the engines' air boxes and both turbocharger compressor turbines, together with their associated valves were fire damaged. On-site examination of the turbochargers did not reveal any anomalies, or foreign object, or other damage that might have adversely affected their operation. Deformation damage to the right engine's exhaust system was consistent with engine operation at ground impact. The left and right engines were recovered from the accident site and transported to an authorised overhaul facility for subsequent inspection under the supervision of the Australian Transport Safety Bureau (ATSB).

All six propeller blades separated from their respective hubs during the ground impact. Four intact propeller blades and segments from the remaining two blades were recovered from the accident site for subsequent technical examination.

Both engine propeller hubs flanges displayed indications of rotation at high RPM at the time of ground impact and were recovered from the accident site for subsequent technical examination. Severe impact damage to the propeller static stops prevented any determination of the propeller pitch settings at the time of ground impact. All of the propeller counterweights were recovered and examined on site. That examination identified overload of the threaded inserts, to the extent that a number of the counterweights had separated from their housing. That corroborated the earlier evidence of high engine RPM at the time of ground impact.

All of the cockpit and cabin seats and structures, along with the seat belts and their attachments, were destroyed by the fire. Most of the cabin fittings and cockpit, including instrumentation and switches were also destroyed. The nature of the damage to the switches was such that their position prior to the ground impact could not be ascertained. The control columns and flap actuator were destroyed in the fire and the engine controls, and the cockpit instruments and radios were severely damaged. While that prevented the examination of most of the instruments, the attitude indicators and annunciator panel were recovered for subsequent technical examination.

Pre-impact flight control continuity was confirmed for the elevator and rudder control surfaces. Flight control continuity was evident for the ailerons, from the cockpit controls aft to the point where the wing impact damage occurred. A continuity check of the engine controls was not possible as a result of the fire damage. The nature of the damage to the right flap indicated that it was retracted at the time of ground impact. The more extensive damage to the left flap precluded a definitive assessment of its position at ground impact. The landing gear was fully extended. The tyres were destroyed by the fire. The severe disruption of the tail structure rendered the determination of the aircraft trim measurements inconclusive.

Examination of components recovered from the wreckage

The left and right engines were disassembled and inspected at an authorised overhaul facility under the supervision of the ATSB and with the engine manufacturer's representative in attendance. That inspection found no evidence of internal mechanical failure within either engine, or of their associated accessories or components that would have prevented the normal operation of either engine prior to the accident.

Technical examination of the propellers and propeller segments indicated multiple high-energy hard object impact signatures on all blade surfaces. Several of those impacts were of sufficient force to have caused ductile shear of the outer airfoil sections. In addition, there was backward curling or loss of material from the blades' leading edges, with associated chordwise scoring and gouging across the airfoil sections. That was consistent with each propeller being actively driven by a comparable amount of power from within the respective engine's upper operating range at the time of ground impact.

Both engine propeller hubs sustained similar multiple fractures to their aluminium alloy housings that was consistent with ductile overload during the accident sequence. There was no indication of any pre-impact cracking or manufacturing defects. The propeller hubs were exposed to gross bending loads through the blade sockets, which was assessed as being consistent with the magnitude of the impact forces that damaged the propellers.

The technical examination of the primary attitude indicator proved inconclusive due to the extensive heat damage to the instrument. There was evidence of rotational scoring to the inside of the secondary attitude indicator's instrument case and to the gyro armature, which indicated that pneumatic drive was available to the aircraft's vacuum instruments at the time of ground impact.

The filaments from the annunciator panel globes were distorted and encased in molten glass as a result of the fire. That prevented the analysis of whether any of those lights had been illuminated at the time of ground impact.

Medical and pathological information

A review of the pilot's aviation-related medical records and the results of the pilot's postmortem examination found no evidence of any pre-existing medical disease, sudden illness or incapacitation that may have affected his ability to control the aircraft.

Fire

There was no report by the witnesses to the take-off, or evidence, of an in-flight fire. The tourist resort volunteer fire-fighting crew responded to the accident site and scrub fires.

The source of the intense post-impact fire was fuel that had spilled from the ruptured wing fuel tanks. The ignition source of the fire could not be confirmed but was likely the hot engine exhausts.

Survival aspects

The emergency locator transmitter (ELT) was destroyed in the post-impact fire. There was no report from the charter pilot, or from the search and rescue authorities to indicate that the ELT had activated on ground impact.

The destruction of the cockpit and cabin from the combined effects of the impact forces and fire rendered the accident non-survivable.

Tests and research - aircraft fuel

The last recorded refuel of the aircraft was the addition of 594 litres of aviation gasoline 100 at Broome on 27 August 2004. It was reported that the pilot refuelled the aircraft's tanks to capacity. The investigation team quarantined a sample of that fuel for subsequent analysis by an approved National Association of Testing Authorities facility. That analysis indicated that the fuel:

• was clear and bright
• was free from water and sediment
• conformed to specification for aviation gasoline 100.

Examination of the Broome fuel supplier's records confirmed that 18 other aircraft were refuelled from that source after the occurrence aircraft on that day. There were no reports from the pilots of those aircraft of any fuel-related problems.

There were no aircraft refuelling facilities at El Questro.

Additional information

Use of aerodromes

Civil Aviation Regulation 92 places responsibility for ensuring that an aircraft landing area is suitable for landing or take-off with the pilot in command. In addition, the regulation requires the pilot to have regard to the prevailing weather conditions and other circumstances affecting the proposed landing or take-off. The determination by a pilot of which other circumstances should be considered is not stated in that regulation.

Civil Aviation Advisory Publication (CAAP) 92-1(1): Guidelines for Aeroplane Landing Areas includes guidance on the factors that may be considered by a pilot when determining the suitability of a potential landing area. While there was no evidence that the pilot had considered the requirements of the CAAP prior to planning his arrival at the El Questro ALA, the investigation applied the minimum landing area physical characteristics recommended by the CAAP to the pilot's take-off from runway 32 until the point at which the aircraft first struck a tree to the left of the runway. That examination determined the following relevant recommended parameters for the take-off:

• minimum runway width - 15 m
• required runway length - about 624 m
• suitable lateral transitional slope, which the CAAP notes could allow for a desirable area of increased lateral clearance during the take-off and may reduce wind shear if near tall trees - maximum obstacle height of about 7.2 m at 66 m from the runway centreline.

The pilot's family indicated that the pilot had operated at a gravel airstrip in the south-west of Western Australia on a number of occasions over the previous 3 years. In addition, the family reported that the pilot drove to El Questro in June 2003 and, during that visit, most likely observed the ALA.

Preparation for flight

The pilot submitted a flight notification to Airservices Australia on the morning of the accident, for a 2 hours 15 minutes flight under the Visual Flight Rules from El Questro to Broome. The investigation estimated that 740 lbs of fuel remained after the reported 2-hour flight from Broome to El Questro, and the POH stated that 50 lbs of fuel was required for 'taxiing for take-off'. That, and the endurance nominated by the pilot in the flight notification, indicated that sufficient fuel was carried for the planned flight to Broome.

The POH stated that the aircraft equipment included a control column lock that restricted control column movement and held the ailerons in a neutral position and the elevators at about 10° trailing edge down. The aircraft manufacturer indicated that the design of the control column lock was such that, if inadvertently left engaged by a pilot, the aircraft would be unable to take-off. The available documentation indicated that an optional rudder gust lock was also included in the aircraft equipment. The POH stated that engagement of that lock required the rudder to be centralised and the elevators to be moved to the fully 'down' position. Disengagement of the rudder lock was possible either manually during the aircraft pre-flight, or automatically as the elevator was moved up through the 6° 'down' position. Due to the damage to the aircraft, the investigation was unable to confirm the position of these locks.

Manufacturer data

The POH promulgated the necessary checks to be carried out by the pilot when operating the aircraft. That included confirmation of the selection of the left and right engines to the left and right main fuel tanks respectively as part of the following checks: before start, before take-off and during the descent. In addition, the POH stated that:

A take-off with one main tank full and the opposite tank low on fuel creates a lateral unbalance. This is not recommended since gusty air or premature lift-off could create a serious control problem.

The published take-off technique included the requirement for the pilot to raise the nose wheel at 95 kts indicated air speed (KIAS) and lift the aircraft from the runway at 100 KIAS. In addition, the POH included a description of the aircraft stall including that:

• the stall characteristics are conventional
• there is an aural stall warning device that operates at 5 to 10 KIAS above the stall in all configurations
• the stall is preceded by a mild, aerodynamic buffet, which increases in intensity as the stall is approached
• the power-on stall occurs at a very steep pitch angle, either with or without flaps extended
• it is difficult to inadvertently stall the aircraft during normal manoeuvring.

Stall speeds were published in the POH for a number of aircraft configurations. None of those configurations reflected the aircraft's take-off configuration. At the estimated aircraft weight and with wings level, the maximum stall speed for the published configurations was calculated as 83 KIAS. The increase in stall speed at 15° angle of bank for all published configurations was about 2 KIAS.

The manufacturer's Pilot Safety and Warning Supplements identified a rare, but potentially serious problem known as 'split wing flaps'. Split or asymmetric wing flaps may result from a mechanical failure in the flap system and cause the flap position on one wing to differ from that of the opposite wing flap. The result is a tendency for the aircraft to roll in the direction of the retracted flap. Depending on the experience and proficiency of a pilot, the manufacturer indicated that any rolling tendency caused by a split flap situation may be controlled with opposite aileron. In addition, there was the potential for a pilot to apply differential power in a multi-engine aircraft to assist in managing the condition. This is discussed further in the Analysis under 'Take-off'.

1. Only those investigation areas identified by the headings and subheadings were considered to be relevant to the circumstances of the occurrence.
2. A take-off commenced by a pilot without pausing an aircraft in a stationary position on a runway, or decreasing the speed of an aircraft on arriving at a runway intended for use for a take-off.
3. A witness that heard, but did not observe the take-off.
4. On 14 November 2003, the pilot submitted an application to the Australian Civil Aviation Safety Authority (CASA) to reserve an Australian aircraft registration in anticipation of registering the aircraft in Australia. That reservation was granted by CASA.
5. Based on the 50-hourly inspection that was carried out in Broome on 26 August 2004 being conducted within the potential 10-hour extension period that had been authorised by the Swiss regulatory authorities. That was between 50 and 60 hours after an annual/200 hours check that was reported as being commenced in Switzerland on 10 December 2003, and included the 2.0 hour flight to El Questro ALA.
6. Derived from e-mail correspondence from the pilot dated over the period 16 September 2003 to 24 August 2004 and the available aircraft and engine documentation.
7. Derived from the right engine Export Certificate of Airworthiness of 19 May 2004 and e-mail correspondence from the pilot dated 29 July to 19 August 2004.
8. Precise date not available.
9. A miniature whirlwind with the potential to be of considerable intensity, and to pick up dust and perhaps other items and carry them some distance into the air.
10. The time nominated by a pilot for the initiation of Search and Rescue action if a report has not been received by the nominated unit.

ANALYSIS

The circumstances of the accident are consistent with dynamic rollover occurring after the right wheel penetrated the helipad surface and acted as a pivot point about which the helicopter rolled. As the helicopter rolled, the main rotor blades contacted the walkway. Because the timing and order of the events that occurred during the rollover were not available, a detailed analysis of the accident sequence was not possible. However, the pilot's response of applying opposite cyclic control, but not lowering the collective pitch control when the helicopter began to roll was not the most effective method of preventing the development of the rollover. Whether the rollover may have been prevented and/or the consequences limited had the pilot lowered the collective pitch control was not able to be determined.

The pilot did not establish before the landing that the load capability of the helipad was appropriate to the requirements. He appeared to have two bases to support his decision to operate the Agusta A109C onto the helipad. One was that neither the helipad owner, nor CASA, had told him that he could not. Such reasoning was inadequate and demonstrated a low level of risk awareness. The second was that other helicopters of operating weight similar to the Agusta A109C had used the helipad. That reasoning was also inadequate because it did not take into account that those helicopters were skid-equipped, and would not subject the helipad deck to the point loads that would exist with the Agusta A109C. It demonstrated an inadequate level of knowledge.

The location of the deck failure was where the plywood would be subject to the greatest bending for a given load, and consequently, where a failure was most likely to occur. The potential for a similar failure would have been present since changes were made to the plywood sheeting size and grade. However, variability between factors such as helicopter wheel positions, dynamic movement of the helipad, and the pilot's flight control inputs would have changed the effect on the deck of each landing. It is likely that the crack in the underside of the plywood in the area of the deck opposite the failure location was an example of the effect of those factors during a previous landing by the helicopter.

Documentation regarding maintenance of the helipad indicated that the installation of the smaller size sheeting most likely occurred during repairs that were completed in September 2000. At that time, other than by way of the design drawings prepared in 1984, full-width plywood sheeting was not specifically addressed either by the oversighting government department or the contractor conducting the repairs.

The strength of the helipad deck compared with the load imposed by the helicopter can be demonstrated from the following information:

It is apparent that the strength of the plywood used in the decking (61 MPa) was approximately 12 times greater than required by the design (5.138 MPa). By comparison, the load applied by each main wheel (928 kg) was approximately 13 times greater than the design load (73 kg per panel).

It is unlikely that the helipad owner calculated the load bearing capability of the helipad with respect to specific helicopters that used the helipad, or their type of landing gear. In particular, the reason for the 2,000 kg load limit, as advised to users of the helipad, appears to be unsubstantiated. The potential for the plywood to be subjected to point loads by helicopter wheels, was not taken into account during the design of the replacement deck. Together with the dynamic movement of the deck by wind and wave action, that effect would have significantly increased the effective ratio of wheel loading to deck design load.

The load limit issues, the installation of plywood deck sheeting of size inappropriate given the existing supporting structure, and the poor condition of some areas of the deck indicated shortcomings in the maintenance and management oversight of the helipad.

Prior to April 2004 there was apparently no awareness within the helipad owner's organisation of the significance of helicopter landing gear configuration to the load bearing capability of the helipad. Nor were there indications that the owner was aware of CAAP 92-2(1), or the guidance it contained. It is likely that those issues were unrecognised due to a lack of aviation technical expertise within the helipad owner's organisation.

CONCLUSIONS

Factors

1. Management and maintenance of the helipad did not encompass all the aspects necessary to ensure that the actual load bearing capability of the helipad, including for skid and wheel type landing gear equipped helicopters, was known.
2. The actual load bearing capability of the helipad was less than that required for safe use by the Agusta A109C helicopter.
3. The pilot did not check before the flight, whether the helipad was capable of safely accepting the loads imposed by the Agusta A109C helicopter.

FACTUAL INFORMATION

History of the flight1

On 30 July 2004, at about 1120 Eastern Standard Time, an Agusta A109C helicopter, registered VH-ZZN, was being operated on a private flight from the Gold Coast to Brisbane Airport via the Brisbane River Helipad. The pilot and two passengers were on board. The intention was to disembark one of the passengers at the helipad, without shutting-down the engines, and then to continue the flight to Brisbane Airport.

The pilot reported that he established the helicopter in a low hover above the helipad and was satisfied that it was correctly positioned above the painted white landing circle. He then allowed the wheels to settle on the helipad, with the right side of the helicopter facing the river bank. The pilot said that he kept his left hand on the collective pitch control lever and his feet on the tail rotor pedals. He then gripped the cyclic control with his knees so that he could open the cockpit door with his right hand. He leaned out and checked that the helicopter was centrally positioned in the landing circle and that the right main wheel was not causing any visible stress to the helipad surface. He then applied the parking brake and closed the door before telling the passenger to disembark through the right cabin door.

The pilot recounted that when the passenger was safely clear of the helicopter, he applied a small amount of collective pitch control in preparation for lift off. He then felt the helicopter lurch and begin to roll right. The pilot immediately applied full left cyclic pitch control and maintained the collective pitch control position. However, the helicopter continued to roll rapidly right and the main rotor blades struck the walkway safety rails. The helicopter came to rest inverted between the helipad and the riverbank, with the forward fuselage against the junction between the helipad and the connecting walkway.

The helicopter was not fitted with a flight data recorder or cockpit voice recorder, nor was it required to be by relevant aviation regulations. Consequently, there was no information available regarding the events or timing of the accident sequence, including the rate of helicopter movement and any control inputs made by the pilot.

Injuries to persons

The pilot and passenger received minor injuries but were able to egress from the helicopter unassisted.

Damage to the helicopter

The helicopter was substantially damaged during the rollover sequence. Contact with the walkway destroyed the main rotor blades and damaged the forward fuselage. The fuselage was partly immersed in water. There was no fire.

Debris from the helicopter was found up to 200m from the helipad, both on and adjacent to the pedestrian/bikeway on the riverbank. There was potential for debris to have been thrown onto the adjacent road complex.

Other damage

Initial examination of the helipad revealed that the helicopter's right main wheel had broken through the deck sheeting on the edge of the white painted landing circle (Figure 1). See also page 7, 'Helipad examination'.

The helicopter's main rotor blades caused substantial damage to the walkway linking the pontoon to the riverbank.

Figure 1: The helicopter as it came to rest. The arrow points to the location of the deck surface failure

Pilot in command

The pilot, who was the chief pilot of the organisation that operated the helicopter, held a valid commercial pilot licence (helicopters). His total flying experience (all rotary wing) was 1,720 hours, of which approximately 250 hours were on the Agusta A109C type (all on VH-ZZN).
The pilot had operated onto the Brisbane River Helipad on numerous occasions since 1 July 2003 in Hughes 500 and McDonnell Douglas 600 type helicopters, and on 19 occasions in the Agusta A109C type2.

Aircraft information

The Agusta A109C was a twin turbine-engine helicopter capable of carrying up to seven passengers. Its maximum operating weight with an internal load was 2,720 kg. The helicopter was equipped with a retractable landing gear, which comprised a nose wheel and two main wheels.

The estimated weight of the helicopter at the time of the accident (including approximately 300 L aviation turbine fuel) was 2,471 kg. Under those conditions, and taking account of the centre of gravity position, the static load on each main wheel was about 930 kg.
According to documentation provided by the operator, there were no outstanding maintenance requirements for the helicopter at the time of the accident. The pilot reported that the helicopter was operating normally in the period prior to the accident.

Brisbane River Helipad

The 9.4 m square helipad was owned and administered by the Queensland State Government. It was located adjacent to the northern bank of the Brisbane River at the end of Margaret Street in Brisbane City (Figure 2). It was mounted on a tethered floating pontoon that was linked to the riverbank by an 11 m walkway. Safety rails approximately one metre high were attached to either side of the walkway. There was a pedestrian pathway and motor vehicle road infrastructure along the riverbank adjacent to the helipad.

Figure 2: The Brisbane River Helipad location

Documentation relating to the operation and maintenance of the helipad was contained in separate files that were administered by various Queensland State Government departments and branches of departments during the period 1981-2004. Through most of that time, operational control and maintenance responsibility for the helipad were held by separate departments, with maintenance aspects being handled predominantly by the department that had responsibility for maritime safety and/or harbours and marine. No record was found on the files of the original design criteria for the helipad, including the helicopter type(s), operating weights, and type of landing gear (skid or wheels) for which it was intended to be used.

Helipad examination

The helipad deck (Figure 3) consisted of 1,200 mm x 2,400 mm plywood sheeting nailed to 100 mm x 75 mm hardwood joists 600 mm apart. The joists were attached to a series of aluminium 'I' beams mounted on two pontoons. The longer (2,400 mm) sides of the plywood sheets had tongue and groove edges and butted over the approximate mid-point of every alternate joist. The shorter (1,200 mm) sides of the plywood sheets did not have tongue and groove edges and were not structurally connected. There was no supporting structure beneath those sides except at the midpoint position at the location of a joist. The central area of the helipad surface (where the helicopter landed) appeared to have been in relatively good condition.

Figure 3: After removal of the wreckage. Individual plywood decking sheets are evident. The arrow points to the location of the deck failure

Figure 4: Location of the deck failure and its relationship to the supporting framework

The location of the deck failure coincided with the corner of one of the plywood decking sheets (Figure 4). The nature of the failure was consistent with the right wheel of the helicopter bearing down on the plywood sheet midway between two adjacent joists, and near the edge of the sheet. The edge formed the shorter (1,200 mm) side of the sheet and was not structurally connected to the adjacent sheet edge.

Specialist examination concluded that the final failure occurred at the point of highest bending moment between the joists. The individual layers within the decking sheet all failed at the point of highest bending, suggesting that the plywood adhesive and the timber layers had not deteriorated due to water ingress.

Laboratory tests carried out on samples taken from the failed section of the deck revealed a bending strength of 61 megapascals (MPa). The size of the test samples was smaller than stipulated by the standard. However, that bending strength exceeded the bending strength of marine grade plywood3.

Examination of the underside of the deck revealed a significant longitudinal crack on the side of the deck opposite the failed section. The deck upper surface at that location showed no sign of damage. The crack displayed damage to the lower layers of the plywood midway between two longitudinal joists. The cracking was not adjacent to the edge of a plywood sheet and was not observed when the helipad was inspected in April 2004. Specialist assessment of the damage concluded that it was likely to have been caused by a point loading in the order of 1,200 kg.

Some areas of the deck on the river and upstream sides of the helipad were significantly degraded due to water ingress into the layers of the plywood, probably as a result of wave action. Examination of those areas indicated that both the ply adhesive and the wood itself had deteriorated significantly. Deterioration of that nature was consistent with continuous contact with, or immersion in, water for an extended period. Those areas, however, were not related to the occurrence.

Helipad load bearing capability

There were numerous references in the helipad owner's documentation regarding the pad's load bearing capability. The earliest of those appeared in drawings that were completed in 1984 for the installation of a new deck consisting of plywood sheets mounted on aluminium support beams (Drawing No. BN-5-43-1). The design strength of the replacement deck was based on a pressure of 3 kilopascals (kPa) evenly distributed over the entire deck area, although the origin of the selected value was not clear. A pressure of 3 kPa is equivalent to 306 kg evenly spread over each square metre of decking. The layout of the plywood sheets over the aluminium support beams resulted in a design panel size of 600 x 400 mm (0.24 sq.m), equivalent to 73 kg evenly distributed on each design panel. The designer's calculations showed that this loading would create a bending stress of 5.138 megapascals (MPa) in the plywood decking, which was well within the allowable working stresses for the F114 grade structural plywood selected for the repairs. There was no indication in the documentation that the design of the repairs included any consideration for point loads, such as would be experienced from wheel equipped helicopters.

In May 1982, a request to land a helicopter weighing up to 3,630 kg on the helipad was refused because 'it would be unsafe to place this weight on the helipad'. However, there was no supporting documentation for that statement.

A July 1997 document authorising a private operator to use the helipad included the statement that the applicant 'should be aware that there is a load limit for the HLS [helicopter landing site], and aircraft in excess of 1,400 kg should not access the HLS'. There was no information regarding the origin of that figure.

In late 1996, the Queensland Government commissioned a feasibility study by a town planning consultant on the future options for the helipad. The study, completed in late 1997, noted that there was:

… no evidence at hand which would indicate it's [the helipad's] safe load capabilities but it appears to be beyond the operation of the Queensland Emergency Services helicopter - a Bell 412 with a maximum take-off weight approaching 5,400 kg - but comfortably accommodates most small to medium-sized turbo driven helicopters in use in the region. The practical limit appears to be approximately 3,000 kg.

There was no supporting evidence to indicate how that conclusion was reached.

In March 1997, there was internal correspondence regarding the establishment of a deed of indemnity that was to apply to users of the helipad. A deed was developed and in early 1999 was sent to users of the helipad for signature. In June 1999, in response to a request to use the helipad, a list of 'requirements that apply to the use of the helipad' was included, for the first time, in the covering letter to the deed. Those requirements included the following:

1. The helicopter landing site is intended for the use by helicopters within the "small" utility class only i.e. helicopters up to 2,000 kg in mass, and
2. The permittee, or their nominated pilot, are responsible for ensuring that the helicopter landing site is in suitable condition before intended use.

There was no indication in the documents regarding the origin of those 'requirements'. Subsequently, there were other instances where prospective users of the helipad were provided with a copy of the requirements when they were asked to sign the deed of indemnity.

In April 2004, a helicopter operator not associated with VH-ZZN contacted the helipad owner requesting confirmation that an Agusta A109C helicopter could land on the helipad. The operator noted that the weight of the Agusta A109C was between 2,000 and 3,000 kg and that, because the helicopter was equipped with wheel instead of skid type landing gear, a 'wheel footprint' of about '1,200 kg per mainwheel' would exist. A reply was sent to the operator advising that the helicopter weight exceeded the 2,000 kg limit that formed part of the standard terms of use for the helipad. The operator replied, noting that skid equipped helicopters weighing greater than 2,000 kg, such as the AS350 Squirrel5, had landed on the helipad.

In May 2004, the Civil Aviation Safety Authority (CASA) contacted the helipad owner regarding the helipad. CASA requested the weight restrictions and the conditions of use of the helipad, and asked whether the weight limit was clearly marked on the helipad. That request, along with the earlier questions and information regarding the suitability of the helipad for the Agusta 109 helicopter, generated activity by the helipad owner including the proposal that a safety audit of the helipad be conducted as a matter of priority.

The justification for the proposed audit was that there may be safety issues associated with the helipad that were not being adequately addressed, such as appropriate usage and signage, and the proximity of the helipad to pedestrian and motor vehicle traffic adjacent to the helipad. At the time of the accident, those issues had not been progressed.

Helipad maintenance history

The helipad owner conducted major maintenance on the helipad in 1984, 1997, 2000 and 2004. In the intervening periods, minor maintenance was conducted. The maintenance activities generally involved work on the flotation pontoons, mooring chains, and the walkway linking the pontoon to the riverbank, and the helipad deck. For the purposes of this report, however, discussion of maintenance activities will focus on the helipad deck.

The drawings for the helipad prepared in 1984 (Drawing No. BN-5-43-1) indicated nominal deck dimensions of 9,400 mm x 9,400 mm and specified a series of hardwood plywood 'rot proofed' decking sheets 9,400 mm x 1,200 mm, with minimum stress grade F11.

At the commencement of the work phase of the repairs conducted in 1984, the contractor proposed standard size plywood sheeting 2,400 mm x 1,200 mm for the decking but that was deemed not acceptable by the helipad owner because the sheet size was not in accordance with the drawing specifications. Subsequently, a special order of 9,400 mm x 1,200 mm plywood sheeting was obtained. The repairs to the helipad, including the installation of the new 9,400 mm long plywood decking sheets on aluminium support beams, were completed in March 1985.

The major maintenance undertaken on the helipad in 1997 included the replacement of two plywood decking sheets and painting of the deck. The documentation indicated that the replacement plywood sheets were 9,400 mm long.

Inspection of the helipad in May 2000 revealed that, in addition to other items requiring rectification, most of the plywood decking was badly deteriorated and beyond repair. On 4 May 2000, the company that subsequently won the contract to conduct the repairs was asked to provide a price:

… to replace any damaged structural plywood sheeting of the deck of the helipad (both a price to replace each sheet and a price to replace all the sheeting). The plywood shall [be] in accordance with Drawing No. BN-5-43-1 and classified as Marine Grade Structural Plywood.

The contractor provided a price to remove and replace all sheets/deck and to coat all surfaces with a product that he stated would waterproof all surfaces and provide total protection to the pontoon/helipad for many years. The contractor was awarded the contract on 16 May 2000. Aside from the reference contained in drawing No. BN-5-43-1, the documents did not refer to the plywood decking sheet size. Documents dated early September 2000 indicated that the rectification works had been inspected as being completed. There was no confirmation in those documents regarding the grade of plywood or sheet size used on the deck.

Another inspection of the helipad was conducted in December 2003. Among the defects identified were four areas of deck sheeting that had loosened, and that the deck needed repainting. The pontoon was removed from the river in March 2004 and the repairs undertaken. There was no reference to any plywood deck panels or supporting beams being replaced during that work. The helipad was returned to service on 21 April 2004.

Operator's information regarding the helipad

The helicopter operating company's operations manual included a register of helicopter landing sites that were used by the company, and pertinent details regarding those sites. The Brisbane River helipad was included in that register. The register details included a caution note regarding river craft during approach and departure, and local pedestrian traffic during approach termination. The register contained no information regarding the load capability of the helipad.

An email dated 1 December 2003 was sent by the helipad owner to the company operating the helicopter regarding access by that operator to the Brisbane River Helicopter Landing Site. The email stated that the deed of indemnity applying to the operator was attached to the email along with:

a copy of the cover sheet that accompanies the Deed of Indemnity as this cover sheet provides details of access restrictions etc.

However, the investigation was unable to establish conclusively that details of the access restrictions had been attached to the email. The deed of indemnity, signed by the helicopter pilot under the company seal, was returned to the helipad owner by facsimile on 1 December 2003 under the signature of the company operations manager.

The pilot reported that he had not sought any information from the helipad owner regarding the load capability of the helipad and claimed that he was not aware that the helipad had a load rating. He stated that he had never seen any document regarding a load rating for the helipad and that the helipad owner had never advised him of a weight limit for operating on the helipad. His rationale for not seeking information regarding the load capability of the helipad was that both the helipad owner and CASA were aware that he was operating the helicopter onto the helipad and neither organisation had advised him against such operations. He had therefore presumed that it was safe to land on the helipad.

The pilot believed that a Bell 412 helicopter operated by the helipad owner had landed on the helipad, and that 'countless other helicopters of larger sizes' than the Agusta 109 had operated onto the helipad. (Personnel involved in the operation of the helipad owner's Bell 412 helicopters reported that they had never operated the Bell 412 helicopter type onto the Brisbane River helipad. No record was found to indicate that any helicopter larger than the Agusta 109, or any other wheel equipped helicopter, had landed on the helipad.)

Operator's procedures

The operator advised that its standard procedures for operating onto a marine helipad included the following after-landing checks:

• visually determine if the helicopter is actually centred on the heli-pad;
• visually observe the deck of the pad to make sure [sic] it is holding the aircraft well; and
• Apply the handbrake and allow passenger(s) disembarkation.

The pilot reported that he followed those procedures for all landings on helipads with both wheel and skid equipped helicopters, and did so on this occasion. In the case of the Agusta A109C, it was necessary to open the door to view below, and fore and aft of, the helicopter. When he checked after landing whether the right main wheel was causing any stress to the helipad surface, it was as part of his standard practice and not because he was unsure as to whether the deck would support the helicopter.

Regulatory aspects

The Civil Aviation Safety Authority issued the Civil Aviation Advisory Publication (CAAP) 92-2(1), in January 1996 titled Guidelines for the establishment and use of helicopter landing sites (HLS). The CAAP included the following introductory statement:

The information contained in this publication is advisory only. There is no legal requirement to observe the details as set out. The Civil Aviation Regulations detail the legal requirements that must be complied with in relation to use of areas for takeoff and landings by a helicopter. While there may be a number of methods of ensuring that the requirements of the Civil Aviation Regulations are met, this CAAP sets out criteria which ensures compliance with the regulations. The CAAP must be read in conjunction with the Civil Aviation Regulations.

The following text appeared under the heading 'Purpose of this CAAP':

Civil Aviation Regulation 92 (1) states that: "an aircraft shall not land at, or take-off from, any place unless: … (d) the place …. is suitable for use as an aerodrome for the purposes of the landing and taking off of aircraft; and, having regard to all the circumstances of the proposed landing or take-off (including the prevailing weather conditions), the aircraft can land at, or take-off from, the place in safety."

Regulation 92 (1) does not specify the method of determining which "circumstances", other than the prevailing weather conditions, should be considered in any particular case. These matters are the responsibility of the pilot in command and, in some circumstances, are shared with the aircraft operator.
These guidelines set out factors that may be used to determine suitability of place for the landing and taking off of helicopters. Experience has shown that, in most cases, application of these guidelines will enable a take-off or landing to be completed safely, provided that the pilot in command:
• has sound piloting skills; and
• displays sound airmanship.

Under the heading 'Factors that should be considered prior to using an HLS', the CAAP stated, in part, that:

The pilot of the helicopter operating to, from or at an HLS should ensure that:
• the HLS is clear of all:
- persons, other than persons essential to the helicopter operation; and
- objects and animals likely to be a hazard to manoeuvring the helicopter, other than objects essential to the helicopter operation; and
• no person outside the helicopter, other than a person essential to the operation is within 30 m of the helicopter; and
• appropriate permission from the owners and authorities has been given.

The CAAP also listed recommended criteria for two types of HLS - Basic and Standard.

The CAAP defined a basic HLS as 'a place that may be used as an aerodrome for infrequent, opportunity and short-term basis for all types of operation, other than RPT, by day under helicopter VMC6'. The criteria for a basic HLS included that it should be large enough to accommodate the helicopter safely and should have a surface capable of withstanding the static and dynamic load imposed by the helicopter.

The CAAP defined a standard HLS as a place that may be used as an aerodrome for helicopter operations by day or night. The recommended criteria for a standard HLS were more comprehensive than those for a basic HLS.

Documents from the helipad owner revealed that in 1992 the question arose as to whether the helipad complied with the then Civil Aviation Authority (CAA) regulations. In 1994, that issue was raised again and there was concern that if the helipad was moved a further 11 metres away from the riverbank to comply with CAA regulations, such action could conflict with the requirements of the Harbours Act.

The consultant undertaking the feasibility study commissioned in late 1996 advised the helipad owner at that time that the CAA requirement impacting on the distance of the helipad from the riverbank had been removed, and that it was now up to the pilot to determine the suitability of a particular location for landing.

The consultant's report, completed in October 1997, expanded on the advice given in late 1996 and discussed some aspects of CAAP 92-2(1). The study noted that 'the type of HLS to be authorised for use at this site under the Civil Aviation Regulations is a little obscure but would appear to fit the 'Marine HLS' model'. The study concluded that the 'existing facility' did not conform to the Marine HLS criteria due to the proximity of the adjacent infrastructure. The study did not address the basic or standard HLS load bearing criteria contained in the CAAP. There was no evidence on the files examined that there had been any follow-up by the helipad owner regarding the criteria of the helipad against those of the CAAP.

Dynamic rollover

A phenomenon known as dynamic rollover can occur during helicopter operations. It arises when the helicopter is placed in a situation where it is pivoting or rolling around a wheel or skid that is in contact with a fixed object.

Dynamic rollover can occur during or prior to the hover when any part of the helicopter acts as a pivot and the helicopter exceeds a critical angle of roll. This angle is dependant upon control limits and, in most helicopters, is in the order of 15 degrees.

Rough ground or obstructions that pin a wheel or skid to the ground can contribute to dynamic rollover. Several rollover accidents have been caused by hitting an obstruction with the landing gear or by attempting a take-off with an obstruction next to the gear.

Regardless of rotor design, actions needed to correct a roll rate are the same and should be instinctive: simultaneously reduce collective and adjust cyclic to maintain lateral trim.7

Once started, dynamic rollover cannot be stopped by application of opposite cyclic control alone. For example, if the right skid or wheel of a helicopter contacted an object and became the pivot point while the helicopter started to roll to the right. Even with full left cyclic applied, the main rotor thrust vector and its moment would follow the aircraft as it continued rolling to the right. Quickly applying down collective is the most effective way to stop the development of dynamic rollover. The phenomenon can occur in both skid and wheel equipped helicopters, and all types of rotor systems.8

Dynamic rollover has been identified in Australia9 and overseas as a contributing factor in helicopter accidents.

The helicopter manufacturer advised that the critical dynamic rollover angle for the A109C had not been determined. The static lateral rollover limit for the helicopter was 27.8 degrees at minimum flight weight (the most critical condition). The maximum slope from which the helicopter was able to takeoff and land with adequate safety margins was 10 degrees.

1. Only those investigation areas identified by the headings and subheadings were considered to be relevant to the circumstances of the occurrence.
2. The Hughes 500 and McDonnell Douglas 600 were small utility type helicopters whose normal maximum operating weights were less than 2000 kg. Both were equipped with skid type landing gear.
3. Table B1 of AS/NZS 2272 1996 Plywood Marine stated that 'The stress grade of plywood manufactured to this standard is a minimum of F14. The sectional properties, if required, shall be calculated by the method specified in AS/NZS 2269'. Table 4.1 of AN/NZS 2269 Plywood-Structural stated that the characteristic bending strength of F14 plywood was 40 MPa.
4. 'Stress grade' refers to the classification of a piece of timber to indicate, for purposes of structural design, a set of basic working stresses and stiffness appropriate to that piece. The grade is designated by a number preceded by the letter F eg. F8. The letter stands for "force" and the number is the amount of stress, in megapascals, that the timber will withstand without bendin beyond acceptable limits in a test situation.
5. It was reported that a single-engine Eurocopter Squirrel helicopter, which had a maximum all-up weight of 2,100 kg and a twin-engine Eurocopter Squirrel helicopter, which had a maximum all up weight of 2,540 kg had operated onto the Brisbane River Helipad. However, the operating weights of the helicopters on those occasions were not established.
6. Visual Meteorological Conditions (VMC) define the in-flight conditions (visibility and distance from cloud) under which aircraft should operate to conduct flight under the visual flight rules (VFR).
7. Rotorcraft Dynamic Rollover. A new look at the problem. Flight Safety Australia, April 1999.
8.Rotorcraft Flying Handbook (FAA-H-8083-21) 2000, Federal Aviation Administration, Flight Standards Service.
9.ATSB report 200300982.

The pilot landed the helicopter on the Brisbane River Helipad to disembark a passenger. As the pilot prepared to lift off, the right main wheel penetrated the pontoon deck, causing the helicopter to capsize. The circumstances were consistent with dynamic rollover. The pilot did not lower the collective pitch control (the recommended response to dynamic rollover) when the helicopter began to roll.

The helipad consisted of a plywood deck mounted on a framework that formed a floating pontoon moored to the river bank. Prior to the flight the pilot had not checked the load bearing capability of the deck, but assumed that the helipad was capable of accepting the Agusta 109 helicopter. Examination revealed that the point load imposed by the main wheels of the helicopter was close to the tested strength of the plywood pontoon decking material. The additional dynamic load effects due to helicopter movement, and wind and water action, were likely to have increased the loads on the deck surface, causing failure of the plywood.

Documentation relating to the management and maintenance of the helipad revealed that the actual load bearing capability had never been established. Potential users of the helipad had been advised of a load limit of 2,000 kg, although the origin of that figure was unknown. The pilot claimed that he was not aware of that limit.

The engineering drawings for the deck called for plywood sheets that extended the full width of the deck. However, smaller sheets were used when the deck surface was last replaced. That had the effect of introducing weaker areas where sheet ends butted together but were unsupported. The deck failure occurred at one of the sheet ends.

FACTUAL INFORMATION

On 15 June 2004, a Hiller Aviation UH-12E helicopter, registered VH-HMT, crashed during an agricultural operation, after the pilot reported a loss of tail rotor authority. The pilot was not injured in the accident. The horizontal stabilizer was found about 150 m from the wreckage site, indicating that it had separated from the helicopter in-flight. The operator reported that damage to the stabiliser and the tail rotor blades was consistent with the separated stabiliser coming into contact with the tail rotor blades, as it was still connected to the helicopter by the taillight wiring. The operator reported that an engineering examination found no evidence of tail rotor system failure.

The horizontal stabiliser spar tube failed in the area where the tube passes through a collar in the doubler attached to the inner stabiliser rib (refer to figure 1).

Figure 1: Horizontal stabiliser and detail of the spar tube and doubler.

Helicopter

The helicopter was manufactured in 1978 and had accumulated 6,903.50 hours time in service at the time of the accident. In September 2002, at 6,386.25 hours in service, it sustained major damage following a sudden main rotor stoppage. The helicopter was rebuilt and had since accumulated approximately 517.20 hours in service. It was maintained in accordance with the maintenance requirements applicable at the time and had a valid Maintenance Release. It had flown approximately 31 hours since the last periodic inspection, a 100 hourly inspection completed in May 2004.

Horizontal stabiliser

The operator advised the Australian Transport Safety Bureau (ATSB) that the horizontal stabiliser is not a serial numbered item and its time in service since new was not known. The stabiliser was fitted to the helicopter when it was acquired from its previous owner. During the helicopter rebuild in September 2002, the stabiliser was inspected in accordance with the 100 hourly component inspection guide and the Hiller UH-12E maintenance manual, before being re-installed. The component inspection guide required the stabiliser tube to be inspected for corrosion and security and the maintenance manual detailed the spar tube inspection requirements and wear limits. No anomalies with the horizontal stabiliser were reported at that time.

Spar tube failure

The failed horizontal stabiliser was forwarded to the ATSB for detailed examination. That examination revealed that the stabiliser spar tube fractured as a result of fatigue cracking. The crack initiation occurred at a number of locations where the tubing had been reduced in wall thickness by wear. It was apparent that the wear of the tubing was associated with small scale repeated movement between the stabiliser tube and the doubler attached to the stabiliser's inner rib. There was no evidence of any pre-existing defect present.

A search of the ATSB, United States National Transportation Safety Board and Federal Aviation Administration databases found no reported events involving an in-flight separation of the horizontal stabiliser on similar helicopter types.

The Australian Transport Safety Bureau did not conduct an on-scene investigation of this occurrence.

3 Conclusions

3.1 Findings

• There was no witness or mechanical evidence to indicate any problems with the flight controls, fuel system or engine prior to the event.
• Failure of the forward flex plate resulted in drive to the main rotor gearbox being partially disconnected.
• The failed forward flex plate coupling punctured one or both the helicopter fuel tanks and the stainless steal engine compartment firewall.
• The helicopter settled heavily and impacted with terrain.
• The helicopter was destroyed by impact forces and a post-impact fire.
• Egress from the helicopter was hampered by the unusual attitude of the cabin which caused extended exposure of the survivor to the post-impact fire.

3.2 Significant factors

One of the two bolted joints linking the flex plate to the main rotor gearbox yoke in the forward flexible coupling was previously assembled incorrectly, resulting in a lack of clamping force and subsequent fatigue failure of the flex plate and loss of drive to the main rotor gearbox.

2 ANALYSIS

2.1 Helicopter

The initiating event in the occurrence sequence was the failure of the flex plate in the forward flexible coupling. That event was the source of the loud noise that the passenger reported hearing. Once the plate failed, the clutch shaft yoke was retained at one end only.  As a result, the diameter of the yoke and flex plate effectively doubled, allowing those rotating components to contact and puncture the fuel tanks and the horizontal stainless steel fire wall, as observed in the wreckage examination. Fuel leaking from the punctured tanks and firewall, would have fed directly into the engine compartment. The evidence indicates that the fuel then contacted either a hot engine component or some other source of ignition, which could have resulted in the helicopter being on fire before impact. The soot pattern on the ground confirmed that fuel from the punctured tanks was being ejected from the helicopter at first impact with the ground. The fuel may have ignited while the helicopter was still airborne or flash back could have occurred after impact.

One of the bolted joints in the forward flexible coupling had been assembled incorrectly, resulting in a lack of clamping force and subsequent fatigue failure of the flex plate in the forward flexible coupling and loss of drive to the main rotor gearbox.

The build records and digital photographs supplied by the manufacturer showed that the assembly of this joint at the factory had been correct.

Examination of the maintenance history of the helicopter revealed that, while maintenance had been performed on the drive system on a number of occasions since factory assembly, there was no recorded documentary evidence found that any subsequent maintenance had been performed on the particular bolted joint where the failure occurred. The stated procedure employed by the engineer who carried out AD/R22/51 was appropriate and consistent with accepted practices and did not require disassembly of the bolted joint that failed.

Examination confirmed that both tail rotor blades failed as a result of contact with the ground during the impact sequence.

At the time of the incident the helicopter was being operated close to the ground with low forward air speed. That would have left the pilot with little time to respond effectively to failure of a critical flight system such as the main rotor drive system.

The manner in which the failure of the forward flex plate occurred resulted in an unusual emergency situation for the pilot. As well as the obvious noise and out of balance caused by the still rotating but partially separated forward flexible coupling, symptoms would have included a failure of the drive to the main rotor gearbox resulting in decreasing main rotor speed and a nose-left yaw. That was consistent with some of the symptoms normally associated with an engine failure. However, in this case the engine had not failed and initially would have presented the pilot with additional and potentially confusing symptoms. Those symptoms would have included an initial indication of engine overspeed, and continuing tail rotor drive.

The normal response to an engine failure, and also relevant in this case, would be for the pilot to lower the collective control in order to recover any loss of main rotor RPM. The pilot would then allow the helicopter to enter an autorotation descent. If there was insufficient height for the descent to stabilise, main rotor RPM would not have recovered before the pilot commenced the touchdown phase of the autorotation. Low main rotor RPM during the touchdown phase would have decreased the possibility for the pilot to reduce the helicopter's rate of descent and carry out a safe touchdown. It is evident that fatigue cracking in the flexplate initiated at the bolthole and propagated under the washers of the bolted joint towards the edge of flexplate. It was also evident from the fracture surface features that crack growth had occurred over a number of flights prior to the accident flight. Because cracking occurred under the washers in the bolted joint, the opportunity to detect cracking by a pre-flight visual inspection would have been limited to the detection of a crack at the edge of the flexplate arm.

The reliability of a general visual inspection (for example, a pre-flight inspection) is affected by lighting, the ability to get close to the component (proximity), and dirt and dust from the operating environment. Reliable detection of a specific defect would require a directed detailed inspection.

2.2 Survival

The crush damage observed to the right side of the cabin structure showed that impact forces were greater on the right side of the helicopter than the left. As a result, the pilot, who occupied the right seat, was subjected to higher impact forces than the passenger in the left seat. The impact forces in this accident were not directly aligned with the helicopters fore-aft and vertical axes. Hence, the level of protection afforded by the seat/harness combination was reduced and probably contributed to the injuries sustained by the pilot. The extent of the post-impact fire precluded any definite conclusions as to what, if any, material may have been carried under the pilot's seat.

When the helicopter came to rest on its right side, the seat in which the passenger was trapped was in an elevated position and was severely affected by the fire. It was also likely that the difficulty the passenger experienced in escaping from the helicopter was due, at least in part, to him being suspended by the safety harness. The time it took the passenger to release himself contributed to the extent of his burn injuries.

The passenger, although seriously injured himself, walked approximately 8 km to summon help. Unfortunately, due to the accident site's remote location, the pilot succumbed to his injuries before medical assistance could reach him.

2.3 Communication equipment

No record of fitment of a fixed Emergency Locator Transmitter (ELT) could be found in the logbooks for the accident helicopter and no ELT unit remnants for either a fixed or personal unit were identified within the main wreckage. Given the intensity of the post-impact fire, it was probable that, had a fixed ELT been fitted, it would have only operated for a very short period of time before being destroyed.

If the pilot was carrying a personal ELT at the time, it is possible that it separated from his person either during the impact or during egress from the wreckage and was destroyed in the subsequent fire.  In addition, if the occupants carried a satellite telephone, it may have reduced the time taken to summon assistance.

1 FACTUAL INFORMATION

1.1 Sequence of events1

At about 1100 Central Standard Time on 30 May 2004, a Robinson R22 Mariner II helicopter, registered VH-MIB, crashed and caught fire while being operated on a fence inspection flight at Tobermorey Station, NT. The pilot sustained fatal injuries, and the passenger was seriously injured.
The passenger reported that the purpose of the flight was to inspect a fence line bordering the property and then to conduct cattle mustering operations. The helicopter was refuelled to full tanks prior to departure. The weather was fine and sunny, with a slight breeze from the south-east.
The passenger reported that while overflying a section of the fence line about 45 km south of the homestead at 30 to 40 ft above ground level, the pilot initiated a turn back to the left to enable a closer look at a particular section of fence. Part way through the turn there was a loud bang from behind the cabin, followed by "horrendous vibration" and the helicopter immediately began to yaw left and descend. The ground marks showed that the tail rotor blades contacted the ground first and then the forward section of the helicopter's right skid struck the ground, disrupting the front section of the cabin. The helicopter then came to rest on its right side and fire rapidly spread to engulf the cabin area.

The pilot was able to free himself from the helicopter through the broken front section of the cabin, but the passenger experienced difficulty undoing his safety harness and remained trapped. He was eventually able to free himself and joined the pilot at some distance from the wreckage. The passenger reported that he assessed the pilot to be badly injured and directed him to a nearby water hole. The passenger then walked to a water bore approximately 8 km from the accident site, where he met other station personnel. The pilot was deceased when medical assistance arrived at the accident site some hours later.

The passenger reported that he had flown in the helicopter several times. During the last few flights before the accident, and during the accident flight itself, he had detected a vibration that he considered abnormal. The passenger advised the investigation that he had conveyed these concerns to the pilot, who advised that he conducted a good check of the helicopter and was satisfied that there were no problems.

1.2 Wreckage examination

Examination of the wreckage by the Australian Transport Safety Bureau (ATSB) investigation team at the accident site (Figure 1) confirmed that the helicopter was yawing left and moving forward when it struck the ground. The tail rotor blades contacted the ground first. The helicopter then contacted the ground slightly nose down and heavily onto the right skid, causing it to separate from the helicopter. The ground impact marks showed that the helicopter continued to yaw left though about 180 degrees after it struck the ground, before coming to rest on its right side. The majority of the cabin and engine bay, including the entire floor area and cabin structure beneath the pilot and passenger seat positions were destroyed by fire.

Figure 1: Helicopter wreckage (arrow indicates bulkhead deformation behind pilot seat position)

The extent of fire damage meant that a complete examination of the helicopter was not possible. Some aluminium components such as tubing in the flight control system had been destroyed. However, all steel components in the control systems for the main and tail rotor were identified and damage to all of these components was consistent with impact forces or fire. The main rotor blades exhibited damage consistent with low rotor energy at impact. Both tail rotor blades had fractured approximately 1/4 span outboard of the rotor hub centre drive (Figure 2). The failed section of one blade was found adjacent to the main wreckage. The failed section of the other blade was found subsequently about 70 m from the main wreckage. Both blade sections were taken to the ATSB laboratories for further examination.

Figure 2: Tail rotor damage

All of the engine drive system components were identified within the wreckage. The flex plate for the forward flexible coupling of the main rotor drive system was fractured at one of the two attachment points to the main rotor gearbox yoke (Figure 3 and 4). The flex plate, including the clutch shaft, were retrieved from the accident site, for further examination. The flex plate for the intermediate flexible coupling was intact and showed evidence of rotational damage consistent with partial drive system power at impact. The rear flex plate and coupling components were also found intact.

Figure 3: Forward flexible coupling as found in wreckage

Figure 4: Components of failed forward flexible coupling

The fuel system and engine ancillaries were destroyed by the fire. There were vertical cuts puncturing the inside wall of the right fuel tank and the horizontal stainless-steel firewall above the engine. The cuts in the right tank and the firewall aligned with the forward flex plate plane of rotation (Figure 5). The left fuel tank was destroyed by fire.

Figure 5:  Flex plate puncture of stainless-steel firewall above engine.

1.3 Personnel information

The pilot held a commercial pilot (helicopter) licence and was appropriately endorsed on the R22. He was issued with a private pilot (aeroplane) licence in 1974 and a commercial pilot (helicopter) licence in 1990. The pilot was issued with a commercial pilot (aeroplane) licence in 1995. He held a stock mustering rating and a valid class 1 medical certificate. He completed a flight review in the occurrence helicopter on 19 April 2004. At the time of the accident, the pilot had approximately 10,400 hours aeronautical experience. He flew 31 hours in the occurrence helicopter between 17 and 30 May 2004.

1.4 Medical and pathological information

Postmortem and pathology reports did not indicate that the pilot was suffering from any condition that might have affected his performance during the flight. The most significant injuries sustained by the pilot were the result of impact forces rather than fire.

1.5 Fire

There was a fire affected area (sooting) on the ground that extended up-slope from the wreckage (Figure 6). The sooting formed a swirl pattern of decreasing radius in the direction the helicopter was yawing when it contacted the ground.

Figure 6:  Sooting pattern adjacent to wreckage (arrow indicates approximate direction of flight at impact)

1.6 Survival aspects

Three-point lap/sash type safety harnesses were fitted to both seating positions in the helicopter. The passenger reported that both he and the pilot had their harnesses fastened during the flight. Fire damage precluded a detailed assessment of the seats and performance of the crush zones beneath them as well as the seat belt harnesses.

Severe crush damage to the lower cabin bulkhead was evident immediately behind the pilot's seat. (Figure 1).

Following the accident, no Emergency Locator Transmitter signal was received (refer section 1.7.6). There was no mobile telephone coverage in the area and the passenger reported that they did not carry any other communications aids, such as a portable satellite telephone.

1.7 Helicopter information

1.7.1 Helicopter data

The helicopter was manufactured in August 2002 as Serial No 3357M. The most recent maintenance release for the helicopter could not be located. It was reported to have been kept in the helicopter. If so, the maintenance release would have been destroyed in the post-impact fire. Based on other maintenance records and information contained in the pilot's personal diary, the total time in service of the helicopter on 30 May 2004 was estimated to have been 506 hours.

1.7.1 Main and tail rotor drive system

In the R22 helicopter, power to drive the main and tail rotors was transmitted from the engine to the rotor drive train via a multiple Vee belt drive and clutch system. A shaft transmitted power forward from the clutch to the main rotor gearbox and aft to the tail rotor gearbox (Figure 7).

Figure 7:  Main components of main and tail rotor drive systems²

A forward flexible coupling, which includes a flex plate, connected the drive shaft to the main rotor gearbox.  The tail rotor drive system also included an intermediate and a rear flexible coupling. Yoke assemblies at the end of each drive shaft section connected the shaft to the flex plate via bolted joints (Figure 8). The purpose of these flex plates was to accommodate small differences in shaft axial alignment during drive shaft rotation. The flex plates and the bolted joints were critical elements in drive system integrity.

Figure 8:  Components of the forward flexible coupling

The helicopter manufacturer published procedures for assembling and aligning the drive system components, including the allowed tolerances. The design loads of components could be exceeded if those tolerances were not met.

1.7.3 Helicopter manufacture

The helicopter arrived in Australia partly disassembled for ease of shipment. The main drive system components within the engine compartment were assembled during manufacture and not subsequently disturbed for this method of international shipment.

In order to assist the investigation, the manufacturer supplied the investigation with the itemised build records, which included digital photographs of the rotor drive system, for the occurrence helicopter. Those records were reviewed as part of the laboratory analysis of the flexible coupling failure. The review of the build records and photographs for the forward flexible coupling in the occurrence helicopter showed that NAS6605-6 bolts had been used, and a spacer washer had been included in each bolted joint and one thin washer had been installed under the nut, with a palnut (locking nut) fitted to each. The build records showed that the build-up of the bolted joint at the time of manufacture of the helicopter was correct and in accordance with the assembly procedures.

The 'Daily or Preflight Checks' section 4-2 and 4-3 of the manufacturers Pilot's Operating Handbook identified the requirement for a visual check of the flex coupling to ensure there are 'No cracks and 'Nuts tight'. Also required is a check of the yoke flanges for cracking. It further advised, in part, that 'During the following inspection, check the general condition of the aircraft and also look for any evidence of leakage, discolouration due to heat, dents, chafing, galling, nicks, corrosion and especially for cracks. Also check for fretting at seams where parts are joined together. Fretting of aluminium parts produce a fine black powder, while steel produces a reddish brown or black residue'.

1.7.4 Maintenance history

Maintenance records indicated that an Australian certificate of airworthiness for the helicopter was issued on 11 October 2002, after assembly in Australia, following manufacture and acceptance flights in the US.  At that time, the total time in service was 5.1 hours. A summary of subsequent maintenance conducted on the helicopter is as follows. All references to drive system adjustments and/or maintenance have been included.

• 11 March 2003. Total time in service 55.1 hours.  50 hourly engine inspection
• 27 March 2003. Total time in service 98.5 hours.  100 hourly inspection.  Maintenance carried out included adjustment of the engine sheave alignment.³
• 30 June 2003. Total time in service 198.2 hours.  100 hourly inspection.  Maintenance carried out included checking and adjustment of the engine sheave alignment and intermediate flex plate shimming to within limits.
• 15 September 2003.  Total time in service 296 hours.  100 hourly inspection.
• 17 February 2004. Total time in service 384.1 hours.  Civil Aviation Safety Authority Airworthiness Directive (AD) R22/51 'Main Rotor Clutch Shaft', dated 12 November 2003, was incorporated. AD/R22/51 was applicable to all R22 helicopters. It required disassembly of the main rotor yoke (A907) to the clutch shaft joint (A166) (see Figure 6) and inspection of the shaft and yoke for damage including fretting⁴ of bolt holes, cracking in the area of the bolt holes, and the presence of an unapproved jointing compound in the mating surfaces. The helicopter maintenance worksheet indicated that no fretting was evident but that the incorrect jointing compound had been used. The worksheet stated that the AD had been complied with and that the clutch shaft and yoke were reassembled in accordance with the maintenance manual. The worksheet also recorded that a duplicate inspection of the clutch shaft installation and the yoke (A907) assembly had been performed. The licensed aircraft maintenance engineer who carried out the AD reported that he disconnected the yoke (A907) from the forward flex plate, but did not disconnect the flex plate from the main rotor gear box yoke (A908). He stated that he did not perform any maintenance on the bolted joints at the connection between the main rotor gear box yoke and the flexible coupling.
• 27 March 2004. Total time in service 396 hours.  100 hourly inspections.  Maintenance carried out included engine sheave alignment.
• 12 May 2004. Total time in service 476.1 hours.  100 hourly inspections.

The documentation showed that maintenance had been performed on 27 March 2003, 30 June 2003, and 17 February 2004 in the vicinity of the forward flexible coupling that, while it did not specifically necessitate bolt removal, provide opportunities for the forward flexible coupling bolts to be disturbed.

1.7.5 Forward flex plate bolted joint component specification

A review of the diagrams contained in the manufacturer's Maintenance Manual and the Illustrated Parts Catalogue (IPC) revealed a difference in the specifications of the parts in the bolted joints. Notes contained within the IPC explained that bolts of different grip lengths and washers of different thickness were to be used in the flex plate bolted joints to expose between two and four threads beyond the end of the nut. A table comparing the different specifications between the Maintenance Manual table and the IPC is provided in section 4.1.3 of the ATSB technical analysis report attached as Appendix 1.

The manufacturer advised that the bolt length identified in the Maintenance Manual was for use in an earlier version of the manual and was out of date. Corrective action to update this information was scheduled by the company for November 2005, but at the time of writing of this report had not been accomplished.

The manufacturer advised that the company did not publish any warning to maintenance organizations about the discrepancy in bolt length between the Maintenance Manual and the Illustrated Parts Catalogue.  The discrepancy was not considered by the manufacturer to be critical in that the use of either a NAS6605-5 or a NAS6605-6 bolt with the appropriate combination of spacer and washers would give the correct clamp up for proper joint integrity. In the few cases where the -5 bolt did not allow proper installation of the B330-16 palnut, the problem would be self-evident. The manufacturer believed that any engineer performing the installation where the bolt was too short to install a palnut would install a longer bolt or make inquiries to resolve the problem.

1.7.6 Emergency locator transmitter

The maintenance records indicate that the helicopter was imported from the US and subsequently operated by various owners without a fixed Emergency Locator Transmitter (ELT) unit being fitted. This fact was noted on the maintenance releases issued at 5.1 airframe hours total time in service (TTIS) on 11 October 2002 and 98.5 hours TTIS on 27 May 2003, which required the pilot to observe the requirements of CAR 252 and carry a personal ELT. No further entries of this nature were found on maintenance releases issued after this date, nor could evidence be found in the aircraft logbooks that an ELT had been fitted. A search for both a fixed and personal ELT within the wreckage and surrounding accident site was conducted but nothing was found.
The passenger stated that the pilot normally carried a personal ELT on him. No personal ELT was identified among the pilots clothing or personal effects and no emergency signal was received by AusSAR from that location on the day.

1.8 Specialist examination of the failed components

The forward flex plate and the broken sections of the tail rotor blades were subject to detailed examination by the ATSB. The report on those examinations and analysis of the failures is attached as Appendix 1.

The metallurgical evidence confirmed that the failure mode of both tail rotor blades was very similar and was the result of contact with the ground during the impact sequence. The rocky material embedded in the blade tips provided clear evidence that the blades had struck the ground while rotating. The blade that was found about 70 m from the wreckage was thrown that distance as the result of tail rotor rotational energy.

The specialist examination found that the flex plate in the forward flexible coupling fractured as a result of the propagation of a fatigue crack at one of the bolted connections between the plate and main rotor gearbox yoke. Final fracture of the flex plate occurred during operation and not as a result of the collision with the ground. No crack growth or wear damage was observed at the three remaining boltholes. Examination of the bolt installed at the failure location revealed that extensive fretting wear had occurred around the entire circumference of the bolt, in the region adjacent to the flex plate and the regions adjacent to the reinforcing plates. Fretting wear was also evident on the washer surface adjacent to the bolt head.

The bolted joint at the flex plate failure location was found to have a single thin washer under the bolt head and nut, and no spacer washer between the yoke and flex plate. This spacer and washer combination was different from that specified by the manufacturer for use with a NAS6605-6 bolt.

1. Only those investigation areas identified by the headings and subheadings were considered to be relevant to the circumstances of the accident.
2. Diagram with permission of Robinson Helicopter Company.
3. Drive Vee belts sometimes stretch when new and adjustments are then necessary to maintain the correct drive system alignment. Engine sheave alignment is part of that adjustment process.
4. The AD defined major fretting as 'any evidence of the machining marks in any of the bolt holes being partly or fully obliterated'.

The crew of the Robinson R22 helicopter were undertaking a fence line inspection at about 30 to 40 ft above ground level. The crew had initiated a turn back along the fence line for a closer look at a particular section of fence. During the turn, a loud bang was heard, and the helicopter began to rotate quickly before striking the ground.

Both occupants were able to exit the helicopter unaided after it came to rest but sustained serious injuries and burns as a result of a post-impact fire. The pilot subsequently died of his injuries.

The investigation found that one of the bolted joints linking the forward flexible coupling flex plate to the main rotor gearbox drive shaft yoke had been assembled incorrectly. This resulted in subsequent fatigue failure of the flex plate and loss of drive to the main gearbox. Control of the helicopter was then lost at a height from which it was difficult to recover.

The crew of the Robinson R22 helicopter were undertaking a fence line inspection at about 30 to 40 ft above ground level. The crew had initiated a turn back to the left, along the fence line for a closer look at a particular section of fence. During the turn, a loud bang was heard, and the helicopter began to yaw quickly before striking the ground.

Both occupants were able to exit the helicopter unaided after it came to rest but sustained serious injuries and burns as a result of a post-impact fire. The pilot subsequently died of his injuries.

Examination of the helicopter wreckage revealed that both tail rotor blades had failed due to contact with the ground.  In addition, the flex plate in the forward flexible coupling of the main rotor drive was found fractured at one of the two attachment points to the main rotor gearbox yoke. The tail rotor blades and several components from the main rotor drive were recovered for detailed analysis in order to resolve the mechanism of fracture and the sequence of failure.

The flex plate in the forward flexible coupling fractured due to the propagation of a fatigue crack at one of the bolted connections between the plate and main rotor gearbox yoke. Final fracture of the flex plate occurred during operation and not because of the collision with the ground. There was no crack growth or wear damage evidence at the three remaining boltholes. Examination of the bolt installed at the failure location revealed that extensive fretting wear had occurred around the entire circumference of the bolt, in the region adjacent to the flex plate and the regions adjacent to the reinforcing washers. Fretting wear was also evident on the washer surface adjacent to the bolt head. This type and degree of wear damage was indicative of operation with insufficient clamping force in the bolted joint.

A review of the manufacturer's original build-sheets for the forward flexible coupling in the occurrence helicopter revealed that NAS 6605-6 bolts were used, and a spacer washer had been included in each bolted joint and one thin washer had been installed under each nut with a palnut (locking nut) fitted to each.

Examination of the forward flexible coupling retrieved from the accident site, found that the bolted joint had been assembled with a washer and spacer combination that was different from that identified by the manufacturer's original build records. These differences indicated that it was likely that the joint had been disassembled and reassembled during a maintenance action subsequent to assembly in the manufacturer's facility.

The circumstances of the accident are consistent with a loss of control due to insufficient main rotor RPM being maintained, and incompatible control inputs from the instructor and the student following the initiation of the simulated engine failure by the instructor. The reported actions by the instructor indicate that he was attempting to recover the situation and allowing the student to follow him through on the controls. The student also recalled attempting to manipulate the helicopter's controls during the descent. It was unlikely that the instructor could have maintained effective control of the helicopter with both pilots manipulating the controls. Procedures for clarifying who is in control at all times, should be established and followed.

The helicopter manufacturer warned that to recover lost main rotor RPM, the pilot must immediately roll on throttle and lower the collective simultaneously. Both pilots reported that they could not lower the collective to the full down position. The activation of the low rotor RPM warning horn during most of the descent confirms that the collective was seldom in the full down position. The instructor reported attempting to increase the throttle position, but it felt like the student had frozen on the throttle. There were no defects found in the examination of the helicopter that would have explained why the collective was not able to be lowered to the full down position or the throttle increased. The manufacturer cautions that once the main rotor RPM decreases below 80%, pilots may not be able to recover control even if the flight controls are correctly positioned. Both pilots recalled seeing the rotor RPM needle in the vicinity of 80% during the descent. The student's recollection suggested that the rotor RPM may have reduced to below 80%.

The investigation was unable to resolve the differences between the statements by the instructor and the student with reference to the way in which the throttle was reduced.

The seat structures are designed to deform during a high G vertical impact, reducing the load transmitted to the seat occupant and increasing survivability. However, in deforming, the seat structure loses significant strength. In this case, the seat structure lost sufficient strength to allow the left anchor point of the left seat lap belt to tear free, increasing the risk of injury to the seat occupant.

An instructor and student were conducting a training flight from Caloundra aerodrome in a Robinson Helicopter Company Model R22 (R22) helicopter, registered VH-HBI. The weather was fine with a light north-west wind.

The instructor reported that shortly after the helicopter reached the intended initial cruising altitude of 1000 ft, and over a suitable area that was clear of other traffic, he reduced the throttle setting to idle to simulate an engine failure. The instructor said the purpose of the exercise was to test the student's alertness and ability to enter and maintain a stabilised autorotation. He intended to terminate the practice engine failure by introducing engine power at about 500 ft. During interview, the instructor said that he rolled the throttle off quickly but gradually. In a subsequent letter, he advised that he slowly reduced the throttle setting to idle. The instructor advised that he did not announce the simulated failure to the student, and that he had previously initiated 'unannounced' engine failure exercises to the student as part of the student's training.

The instructor reported that the student correctly applied right anti-torque pedal and pulled the cyclic control rearwards, but did not lower the collective lever. He also reported that the low rotor RPM warning horn came on within a few seconds and the airspeed reduced rapidly to between 30 and 40 kts. The instructor said that the helicopter then started shuddering and he lowered the collective as far as it would go and pushed the cyclic forward. He said that, as they approached 500 ft he attempted to increase engine power, but it felt like the student had 'frozen' on the throttle and he was not able to rotate the twist grip. The instructor said that the low rotor RPM horn was on for most of the descent. He said that the rotor RPM was just above the horizontal, around 83%, and the engine RPM was at idle on the engine and rotor RPM indicator (see Figure 1). He said that he did not advise that he was taking control of the helicopter, but provided 'spoken instruction, supported by directive pressure on the controls', because he wanted the student to feel the control movements. Although the instructor recalled applying up collective to cushion the helicopter for landing, the pilots were not able to terminate the helicopter's descent.

The helicopter impacted tidal mudflats near the northern end of Bribie Island in a nearly level attitude with a high rate of descent and low main rotor RPM. The floor of the helicopter sustained significant deformation and physically trapped the student's feet. Both occupants of the helicopter sustained serious injuries during the accident.

An examination of the wreckage did not identify any defect that would have prevented normal operation of the helicopter prior to the accident. The collective was able to be moved to the full down position and there was no indication of any restriction.

The student was unable to recall much of the sequence of events during the occurrence. He said that the instructor did not mention the possibility of unannounced engine failures during the flight. In an interview with the ATSB, he said that the instructor had rolled off the throttle. Later, in a letter he said that the instructor snapped the throttle off very rapidly. The student described the correct response to a simulated engine failure, but said that on this occasion, the collective did not feel like it went all the way down. The student recalled seeing the engine RPM and rotor RPM needles below the horizontal position on the engine and rotor RPM indicator (see Figure 1), when the helicopter was descending through about 700 ft. The student said that during the descent the collective was about halfway up and although he tried to push it down, it didn't feel like it moved.

The helicopter manufacturer advised that deformation of the seats provided additional absorption of vertical energy beyond that required for certification. Both crew seat structures significantly deformed during the impact with the right seat being more affected. The right seat had a plastic first aid kit under it, which had been crushed. The left anchor point of the left seat lap belt had torn free of the seat pan. The rivets attaching the anchor point had stretched, but not separated, however the sheet metal had failed, allowing the anchor point to come free.

The helicopter was operating with a valid maintenance release, and had accrued 387 hrs total time in service. The instructor had accumulated approximately 925 hrs total helicopter experience, of which about 158 hrs was in the R22. He had completed his instructor rating with the same operator, and had a total of about 105 hrs instructional time, all in R22 helicopters. The student pilot held the equivalent of a helicopter student pilot license. He had accrued about 94 hrs in the R22, of which about 16 hrs was in command.

The R22 Pilot's Operating Handbook stated that during an autorotation (prior to the flare), the collective should be adjusted to maintain the rotor RPM in the green arc between 97% and 104%, or approximately 90% to obtain maximum glide distance.

Section 4, page 10 of the R22 Pilot's Operating Handbook stated:

CAUTION - During simulated engine failures, a rapid decrease in rotor RPM will occur, requiring immediate lowering of the collective control to avoid dangerously low rotor RPM. Catastrophic rotor stall could occur if the rotor RPM ever drops below 80% plus 1% per 1000 ft of altitude.

The engine RPM and rotor RPM needles would have been in the horizontal position when the respective RPM values were 80% (see Figure 1). The low rotor RPM warning horn was designed to activate when the rotor RPM was less than 97%. The warning horn did not activate if the collective was in the full down position.

The R22 Pilot's Operating Handbook also included three Safety Notices pertinent to this accident.

The first, titled 'Surprise throttle chops can be deadly', stated:

The student may freeze on the controls, push the wrong pedal, raise instead of lower the collective, or just do nothing. The instructor must be prepared to handle any unexpected student reaction.

The second Safety Notice, titled 'Fatal Accidents caused by Low RPM Rotor Stall', stated:

No matter what causes the low rotor RPM, the pilot must first roll on throttle and lower the collective simultaneously to recover RPM before investigating the problem. It must be a conditioned reflex.

The third Safety Notice, titled 'Practice autorotations cause many training accidents', stated:

As the aircraft descends through 100 feet AGL, make an immediate power recovery unless all of the following conditions exist:

1. Rotor RPM in middle of green arc,
2. Airspeed stabilized between 60 and 70 KIAS [knots indicated airspeed],
3. A normal rate of descent, usually less than 1500 ft/min,
4. Turns (if any) completed.

This Safety Notice also stated:

Practice autorotations continue to be a primary cause of accidents in the R22 and R44. Each year many helicopters are destroyed practicing for the engine failure that very rarely occurs.

A review of the ATSB database identified 18 accidents between 1985 and 2003 that occurred during autorotation training in Robinson R22 helicopters.

Most flight instructor manuals emphasise the importance of establishing and using procedures that at all times identify which pilot has control of the aircraft. For example, the Flight Instructor Guide - Helicopter (1995) issued by Transport Canada stated:

CONTROL OF AIRCRAFT

2. There should never be any doubt as to who has control of the aircraft. ...:

(d) when the student has control, you must not "ride" the controls. Your student may feel that you are taking control and this could lead to a dangerous situation. Additionally, you may rob your student of the feeling of accomplishing the manoeuvre independently. This is particularly difficult during critical manoeuvres, such as full-on autorotations, when there is little time available to the instructor to correct errors. This procedure must be adhered to at all times.

The Ted Smith Aerostar 601 aircraft, registered VH-WRF, departed Coolangatta at 1301 ESuT with a flight instructor and a commercial pilot on board. The aircraft was being operated on a dual training flight in the Byron Bay area, approximately 55 km south-south-east of Coolangatta. The aircraft was operating outside controlled airspace and was not being monitored by air traffic control. The weather in the area was fine with a south-easterly wind at 10 - 12 kts, with scattered cloud in the area with a base of between 2,000 and 2,500 ft.

The purpose of the flight was to introduce the commercial pilot, who was undertaking initial multi-engine training, to asymmetric flight. At approximately 1445, the operator advised Australian Search and Rescue that the aircraft had not returned to Coolangatta, and that it was overdue. Recorded radar information by Airservices Australia revealed that the aircraft had disappeared from radar coverage at 1335. Its position at that time was approximately 18 km east-south-east of Cape Byron. Search vessels later recovered items that were identified as being from the aircraft in the vicinity of the last recorded position of the aircraft. Those items included aircraft checklist pages, a blanket, a seat cushion from the cabin, as well as a number of small pieces of cabin insulation material. No item showed any evidence of heat or fire damage. No further trace of the aircraft was found.

Air traffic control received normal radio transmissions from the aircraft during the departure from Coolangatta and the transit to the operating area. No other transmissions from the aircraft were received. In particular, there were no radio transmissions from the aircraft around 1335.

The instructor was the chief flying instructor of the organisation that owned the aircraft. He held a Grade 1 flight instructor rating. He had extensive civilian and military experience as a pilot and flight instructor, most of which was on single-engine fixed and rotary wing aircraft. At the time of the accident, his total flying experience was 7127.2 hours, with 447.7 hours on multi-engine aircraft. The instructor had accrued 294 hours on Aerostar aircraft of which 194 were as a flying instructor.

The student pilot held a commercial pilot's license and had about 283 hours flying experience at the time of the accident. He commenced flying training on 23 April 2001 on a Mooney M20J aircraft, and first flew solo on 23 July 2001, after 38.3 hours dual instruction. On 21 December 2002, he gained an unrestricted private pilot's license, having flown a total of 181.1 hours. On 19 September 2003, with a total flying experience of 257 hours, the student pilot failed his first commercial pilots license test. On 17 October 2003, he passed the second attempt with a total flight time of 276.9 hours. Excluding the accident flight, the commercial pilot's flight time on Aerostar aircraft was 3.5 hours.

The accident flight was the commercial pilot's fourth in the Aerostar aircraft and was the third flying exercise sequence in the operator's multi-engine training syllabus. The objectives of the exercise included controlling the aircraft after the failure of an engine, recovering from a stall in the take-off configuration, and entering and recovering from a minimum control speed (Vmca) situation. (Vmca is the minimum control speed in flight with one engine inoperative.)

Recorded radar data for the flight showed that the aircraft proceeded from Coolangatta to the east of Byron Bay where it conducted a series of manoeuvres in an area approximately 18 km square, at between 2,500 and 3,000 ft, with occasional brief excursions below 2,500 ft. Between 1313:00 and 1328:00 ESuT, there were three instances, about 4 minutes apart, where the recorded groundspeed of the aircraft decreased rapidly from approximately 150 kts to between 100 and 110 kts before increasing again. Each speed reduction was accompanied by an altitude loss of 200 - 400 ft. Between 1328:40 and 1329:05, the groundspeed decreased from 140 to 118 kts. It then fluctuated between 121 and 112 kts for the next 1 minute and 25 seconds while the recorded altitude reduced from 2,600 to about 1,900 ft. The recorded altitude and ground speed then steadily increased for the next 3 minutes to a maximum of 2,800 ft and 123 kts respectively. During the next 60 seconds, the recorded altitude reduced to 2,500 ft while the groundspeed decreased to 110 kts at 1335:00. At 1335:29, the recorded altitude was 2,600 ft and the groundspeed 108 kts. The last valid radar data was at 1335:37 when the recorded altitude was 2,100 ft, and the groundspeed was 100 kts. The aircraft was tracking in a south-easterly direction from about 1331 until radar contact was lost. There were no sudden or significant changes in the recorded track during that period.

The aircraft was being operated on a valid maintenance release and there were no maintenance items outstanding at the time of the accident. The aircraft was flown from Coolangatta to Sydney and return on the night before the accident and was reported to have operated normally during those flights. The aircraft was not fitted with a stall warning system.