History of the flight

The Lancair, registered VH-CIV, was a homebuilt aircraft operating under an experimental certificate of airworthiness. The owner had recently built the aircraft and had commenced a test flight program. The aircraft departed with a test pilot occupying the pilot seat, and the owner/builder occupying the right seat. Witnesses reported seeing the aircraft flying in the Drysdale area before descending steeply. It subsequently impacted the ground and the two occupants were fatally injured.

The pilot departed from Point Cook and flew along the coast, southwest of Point Cook aerodrome, at altitudes between 3,000 ft and 4,000 ft. After crossing Port Phillip Bay to the Bellarine Peninsula the pilot climbed the aircraft to between 5,000 ft and 6,000 ft. Recorded flight data indicated that the aircraft was slowed, and stalled at an altitude of 5,200 ft. The pilot recovered from the stall, but the aircraft entered another stall during the recovery at 4,950 ft. The pilot recovered the aircraft from that stall, and then repositioned it and entered a third stall at an altitude of 6,200 ft. The aircraft rolled at the initiation of the stall, and continued to roll as it then descended rapidly, accelerating to approximately 150 kts at an angle of approximately 40 degrees from the horizontal, with low engine power. The engine power increased shortly before the aircraft impacted the ground.

The aircraft impacted the ground upright, with the wings level, at a pitch angle of 40 degrees nose down. The landing gear was retracted at the time of impact, and the flaps were either completely or nearly retracted. There was no indication of any mechanical failure prior to impact. The accident was not survivable.

Flight crew

The pilot was an experienced ex-military pilot with 6,500 hours total aeronautical experience, including experience over many years in aircraft with a wide range of performances.

The owner/builder had a Commercial Pilot (Aeroplane) Licence issued 16 October 2001. He had 352 hours total aeronautical experience in a range of general aviation aircraft with a level of complexity and performance up to four-seat single-engine aircraft, with a constant speed propeller and retractable landing gear.

Aircraft

The aircraft was originally being constructed to comply with the requirements of Civil Aviation Safety Authority (CASA) Civil Aviation Order (CAO) 101.28. During construction, the experimental designation concept was developed and the aircraft construction was subsequently changed and completed in accordance with the experimental designation under Advisory Circular (AC) 21.4(2) issued September 2000. The certification requirements in the experimental category were less prescriptive than those in CAO 101.28 as experimental aircraft are not type-certificated. 'Experimental' is not a category but rather a designation. It is also important to note that an experimental certificate does not attest to an aircraft being fully airworthy (Civil Aviation Safety Authority Advisory Circular 21.1(1) - Aircraft Airworthiness Certification Categories and Designations Explained - Revised December 2000).

The aircraft was based on a Lancair IV-T kit plane that was originally designed to use a high-performance piston engine. The owner/builder decided to replace the engine with a turbo-propeller engine and sought assistance from a number of people to redesign the aircraft to accommodate the new engine.

Documentation indicated that during construction, numerous changes had been made to the original design, including the engine type and the design of the aircraft from the firewall forward. The propeller had been modified by removing 20 cm of the propeller tips. No evidence was found to indicate that any form of risk assessment had been undertaken to consider the safety implication of these aircraft design changes. A risk assessment was not required for aircraft constructed under the experimental designation.

Aircraft equipment

The aircraft was equipped with an integrated electronic flight information system (EFIS) comprised of a number of data acquisition units and three liquid crystal multi-function display (MFD) units, that could be configured to present operational information in different ways. The instrumentation system received data from a variety of sources, and integrated that data to present operational information to the pilot. Each MFD also contained a memory unit that recorded information that was sent to it. This information was used by the ATSB to determine the flight profile prior to the accident.

The MFDs could be selected by the pilot to display primary flight information, navigation information, or engine system information. The primary flight information consisted of an electronically generated artificial horizon display upon which a number of other information displays were overlayed. Indicated airspeed (IAS) was also displayed on the left side of this display. The IAS was presented as a vertical tape display of airspeed with a range of approximately 100 kts, in increments of 10 kts. The actual airspeed was also displayed numerically on the left side of the artificial horizon display, adjacent to the IAS speed tape display. The airspeed tape display contained a number of coloured regions to indicate various speed ranges and included the stall speed of the aircraft. The stall speed was dynamically generated and changed with varying flight loads on the aircraft. The displayed stall speed was known as the G-corrected stall speed. The EFIS derived the G-corrected stall speed from the aircraft's static 1G stall speed. The 1G stall speed had to be manually entered into the EFIS. The aircraft instrumentation system had been programmed with a 1G indicated stall speed of 69 kts with flaps and landing gear retracted.

The aircraft was also equipped with a video camera mounted on the ceiling just behind the two front seats, which was aimed at the instrument panel. This camera operated for some of the test flights, and some of the recordings were recovered from the camera after the accident.

The original aircraft design (with a piston engine) had a declared indicated stall speed of 84 mph, or 73 kts with flaps and landing gear retracted. The Comparative Aircraft Flight Efficiency (CAFE) foundation test flight of this version of the aircraft type recorded a stall speed of 78 kts with flaps and landing gear retracted. That test flight was conducted with a calibrated pitot/static boom, so the indicated airspeed would have been accurate in that instance. A theoretical aerodynamic calculation for the accident aircraft indicated a stall speed of 82 kts. Recorded information indicated that the aircraft stalled at 82 kts at the commencement of the accident sequence.

A flight path marker symbol, in the centre of the primary flight display, indicated the current pitch attitude of the aircraft. At an airspeed of 20 kts above the G-corrected stall speed, an amber pitch limit indicator symbol appeared above the flight path marker. The pitch limit indicator symbol moved down toward the flight path marker symbol as the airspeed decreased towards the stall speed. At 5 kts above the stall speed, the pitch limit indicator symbol turned red. At the point of stall, the pitch limit indicator symbol overlayed the flight path marker symbol. Coincident with the pitch limit indicator symbol changing to red, an aural annunciator would repeat a voice warning of 'stall' and a red stall flag symbol would display in the bottom left corner of the primary flight display.

The aircraft was also equipped with an angle of attack sensing device that used pressure information from two ports on the upper and lower sides of the left wingtip to derive the angle of attack. This instrument could have indicated when the aircraft was approaching a stall. The investigation did not determine if it had been calibrated before the accident flight.

During a test flight on 9 December 2002, the test pilot had decelerated the aircraft until it was approaching the stall as a part of the test flight program. Written records from that flight noted:

'A/c [aircraft] becomes laterally unstable below 80 Kts.'

Recorded flight data indicated that the aircraft stalled three times during this flight, from speeds of 72, 75 and 76 kts. The test pilot had also noted:

'Close to stall at 75 Kts with 10 [degrees] flap
Close to stall at 72 Kts with full flap'.

Video recordings from the flight indicated that the aircraft had stalled, and rolled 45 degrees to the left, and then 45 degrees to the right of horizontal during the recovery from the stall. Stalls were not a part of the test flight program for the flight.

Test program

A friend of the owner/builder developed a program for the flight testing of the aircraft. That program followed the guidelines in the Federal Aviation Administration (FAA) Advisory Circular 90-89A - Amateur-built Aircraft and Ultralight Test flighting Handbook. The CASA Advisory Circular 21.4(2) - Amateur Built Experimental Aircraft - Certification, advised that CASA:

'...most strongly urges [Amateur-built experimental aircraft builders] to make detailed reference to [this document], prior to their flight programs commencing, and [to] follow the guidance provided.' (section 14.4).

The test flight program was detailed, and was clearly developed in accordance with the recommendations of Advisory Circular 90-89A. Neither the advisory material, nor the test flight program considered action to be taken if aircraft handling or performance produced unexpected results.

CASA Advisory Circular 21.4(2) stated:

'14.5 Those undertaking test flight programs may also derive benefit in consulting the following additional references, as applicable to the class of aircraft involved:

'(a) CAA publication dated January 91, Flight Test Guide for Certification of CAO 101.28 Category Aeroplanes...'

This document recommended calibration of aircraft flight instruments, so that the limits of the flight envelope could be accurately determined. The test flight program for the accident aircraft did not include the in-flight calibration of aircraft flight instruments, including the airspeed indicator. Comparison of the recorded airspeed from the satellite navigation system, fitted to the aircraft as apart of the EFIS, with the recorded airspeed from the pitot/static system showed no inaccuracy in airspeed indication.

There was no evidence that any significant risk assessment was undertaken during construction of the aircraft and in the development of the test flight program. Such a risk assessment could have examined the planned activities and considered any potential hazards for their likely impact on the aircraft's safety during test flights.

The friend who developed the test program stated that he conducted the first two flights of the aircraft, but subsequently did not participate any further in the test program. Prior to those initial flights, he conducted an operational pre-flight briefing with the owner. These briefings included hazards and potential actions. The investigation did not determine whether the test pilot for the accident flight conducted similar operational pre-flight briefings for subsequent flights.

There was no evidence that any significant re-evaluation of risk was done during the subsequent conduct of the test flight program. Such a risk assessment process could have examined the results of test flights for hazards that became apparent from analysis of observations and data from each flight. This could have allowed for a considered assessment of any risk mitigators for their likely impact on the aircraft's safety during subsequent test flights, as the aircraft flight envelope was expanded. For example, if unexpected handling characteristics had been encountered during a stall sequence, then previously identified mitigation procedures, such as moving the centre of gravity forward, could have been considered.

Aircraft design

An aircraft operated as an experimental aircraft does not have to comply with any specific design. Builders may comply exactly with a design, or may deviate from that design as much as they wish, or may build an aircraft independent of any previously developed design.

If an aircraft is built mainly in accordance with an established design, but with some design differences, then it may perform differently from an aircraft built exactly in accordance with that established design. If design differences are incorporated one at a time, then it is possible to measure the effect of any single design difference. If many design differences are incorporated at one time, the effect of a single design difference on the behaviour of the aircraft may be impossible to predict due to the compounding effect of other incorporated design differences. The accident aircraft had been constructed with many differences compared with the original Lancair IV-T design.

Required persons on board during test flights

A CASA delegate had issued a special certificate of airworthiness authorising flight in accordance with the test flight program. The approval included operating limitations for the test flight program, such as geographical limitations, minimum weather conditions for flight and the maximum number of persons to be on board the aircraft.

Test flights are hazardous compared with normal flight. Accordingly, there is normally a requirement for only operational persons to be on board an aircraft during a test flight. The special certificate of airworthiness under which the accident flight was made stated 'Only personnel essential for the conduct of the testing may fly on board the aircraft. The carriage of passengers is prohibited.'

This aircraft type was normally flown as a single crew operation, and the cockpit of this aircraft was configured for single crew operation. There was no evidence that the test flighting required two persons on board.

Following take-off from Los Angeles International Airport (LAX), the crew of the B747-438 aircraft noticed a severe airframe jolt while conducting a climbing left turn. The cockpit instruments indicated that the number 1 engine exhaust gas temperature was rising through 900 degrees C. Passengers also reported flames emanating from the number 1 engine tailpipe.

The crew shut down the number 1 engine and returned the aircraft to LAX for a one- engine inoperative landing.

An initial investigation carried out by the operator's maintenance personnel revealed that there had been an apparent failure within the engine's high-pressure compressor (HPC) assembly. The engine was removed from the aircraft and transported back to the operator's engine maintenance facility in Australia, where a more detailed examination was carried out. An Australian Transport Safety Bureau (ATSB) metallurgist was present for that examination.

The engine was a Rolls Royce RB211-524G2-T-19/15 turbofan engine. The designation `T', in the engine model number, indicated that the engine had been manufactured with a core engine from the larger Rolls Royce `Trent' engine series. The inclusion of the `Trent' core had enabled the engine to be more fuel-efficient and operate at a lower exhaust gas temperature.

The `Trent' core engine was split into numbered modules. The three modules of interest to the investigation were: Module 33, the intermediate case module; Module 41, the high-pressure system module; and Module 51, the intermediate and low-pressure turbine assembly module.

The engine had a nominal overhaul life of 30,000 hours or 4,000 cycles. At the time of the failure the engine was well within its overhaul life, having been in operation for a total of 13,922 hours and 1,395 cycles. It had not undergone any major maintenance.

The ATSB Technical Analysis report on the engine failure, (see Appendix A), indicated that the engine failure had resulted from the liberation of a single blade from the first-stage HPC rotor in Module 41. The blade release had resulted in extensive damage to the engine. The friction from the liberated blade impacting the surrounding blades on the HPC rotor resulted in a titanium fire within the compressor assembly.

A close inspection of the remains of the liberated blade root stub showed evidence of fatigue cracking and loss of the forward trailing edge corner of the blade dovetail root block. None of the other blades within the first-stage HPC assembly showed any visible evidence of cracking when inspected with the unaided eye. Minor collateral damage was also evident to components in Module 51 and Module 33 resulting from the blade failure.

The manufacturer was aware of three similar failures of the HPC blades in the RB211-524G/H-T series of turbofan engines. In those failures, cracking in the blade root area was believed to have resulted from uneven friction on the blade root bedding surfaces due to a breakdown in blade root lubricant. The manufacturer further indicated that damage in the blade root area that led to local stress concentrations, such as scores and sharp edges, might also have contributed to those blade failures.

The evidence from the ATSB investigation indicated that the failure mode in this incident was the same as the three other known failures in RB211-524G/H-T turbofan engines.

The British Aerospace 146-100A (BAe 146) was being prepared for a regular public transport service from Perth to Newman, WA. As the copilot boarded the aircraft to conduct pre-flight checks he detected strong fumes in the cabin and the flight deck. He noted that the auxiliary power unit (APU) was supplying bleed air to one of the aircraft's two air conditioning packs that in turn supplied conditioned air to the flight deck and cabin. In an attempt to clear the fumes, he directed the bleed air supply to the alternate pack.

Bleed air from the APU was generally used when air conditioning was required during ground operations or during the take-off and landing phases. During flight, the air conditioning packs receive bleed air from the engines.

As the copilot continued the pre-flight checks, he noticed the cabin fumes becoming stronger and so shut down the air conditioning pack and opened the flight deck windows.

Maintenance engineers were requested to investigate the source of the fumes and subsequently discovered an oil leak in the APU generator drive adaptor pad. Rectification work, including the replacement of a carbon seal, was carried out 11 days later, on 13 December. To enable the aircraft to continue in service on the day of the incident, the APU was isolated from the air conditioning system in accordance with the terms of the aircraft's Minimum Equipment List (MEL) that permitted operation of the aircraft in non-standard configurations. The operator reported that the maintenance engineers then addressed the defect in accordance with the Civil Aviation Safety Authority (CASA) airworthiness directive AD/BAe146/86, effective 3 April 2001, and the British Aerospace Systems Inspection Service Bulletin (ISB) 21-150. That ISB called for certain actions to be performed whenever a cabin air quality problem was identified, which was suspected of being associated with oil contamination of the air supply from the air conditioning packs. No oil contamination was found. The engineers then operated both packs using bleed air from engines 1 and 4 until they were satisfied that there were no fumes and the aircraft was then released for service.

The copilot had been exposed to the fumes for approximately 30 minutes. The two cabin crew, who boarded the aircraft shortly after the copilot, were exposed to the fumes for approximately 20 minutes. All three were eventually affected by the fumes, became unwell, and were removed from flight duty. In accordance with company standard practice they underwent medical examinations. The pilot in command was exposed to the fumes for less than 5 minutes and was not affected. Impaired performance due to the inhalation of contaminated air is considered a potential threat to flight safety. For that reason, company procedures emphasise the importance of flight crew donning oxygen masks if poor air quality is suspected during flight.

The pilot in command and a replacement crew subsequently departed in the same aircraft for the flight to Newman. The cabin crew reported a slight smell of fumes toward the rear of the aircraft during the first sector. On the return sector, both cabin crew reported feeling unwell, with symptoms consistent with fumes inhalation.

Follow-up inspections of the air conditioning system, engines and APU were carried out in accordance with the CASA airworthiness directive and no contamination was evident. On 6 December, the operating crew indicated that a smell was apparent and it appeared to be consistent with the operation of the APU. As the APU was still isolated, engineers doubted that it could be contaminating the air conditioning system. All four engines and the regenerative ducts were again checked, with no contamination evident. Follow up inspections were scheduled in accordance with the ISB.

A further cabin air quality event occurred on 12 December, when the flight deck crew detected fumes shortly after departure. The flight crew proceeded to identify the source of the fumes using a contamination source location schedule. That schedule involved selecting different combinations of engine air and air conditioning packs. The fumes were traced to the number 3 engine, which was isolated, and the flight continued as planned. Subsequent inspection revealed oil wetness in the number 3 engine high-pressure compressor; the result of a worn number 1 bearing seal. Trend monitoring had not indicated abnormal oil consumption for that engine. The engine was replaced and airworthiness directive AD/BAe146/86 was complied with. No further contamination was evident and fumes were not reported during subsequent flights.

Evidence from previous incidents of air system contamination on this aircraft type had indicated that fume events were often intermittent in nature and were associated with engine or APU oil contamination of the air conditioning system. The air supplied to the air conditioning packs was protected from contamination by oil seals in the engines and APU. A defect in one of those seals could result in oil entering the cabin air conditioning system, with the first sign of the defect being an awareness of fumes by passengers or crew members.

The investigation of cabin fumes incidents on BAe 146 aircraft has typically been characterised by a difficulty in precisely locating the original source of the oil leak that led to the creation of the fumes. That has been especially so if there was more than one engine/APU leak combination. The failure of oil seals has been a common factor in the majority of those incidents.

The crew carried out the take-off in the British Aerospace Plc BAe 146 (BAe 146) with number 4 engine air bleed selected on, and engine numbers 1, 2 and 3 and the auxiliary power unit air bleeds selected off in compliance with an item in the discrepancy log. Shortly after take-off, at approximately 700 ft above ground level, the copilot switched on the remaining engine air bleeds and both air conditioning packs. Shortly after selecting engine anti-ice on, the pilot in command (PIC) asked the copilot if he could smell fumes. The copilot agreed that he too had detected a smell. The engine anti-ice system was then switched off.

A short time after commencing the cabin service, a flight attendant (FA) called the flight deck and informed the PIC of fumes in the cabin and that they were particularly bad near the rear of the cabin where another FA had donned an oxygen mask. This mask was of the diluter type that supplies a mixture of the incoming oxygen with the ambient air that is then delivered to the user. The copilot then indicated to the PIC that `he felt he should go onto oxygen' and donned an oxygen mask, but the PIC did not feel he needed to perform the same action at that time.

The PIC later stated that he did not don his oxygen mask at this time, as he was considering if the fumes were oil related in accordance with a notice to pilots (NOTOP) from the operator. This NOTOP required the PIC to make a diagnosis as to the source of contamination `wherever it is safe and practicable to do so'. The PIC's findings were required by the operator to determine the level of response required to later rectify the problem and return the aircraft to service.

The PIC stated that the aircraft type had a history of fumes related problems and not donning his oxygen mask was a normal practice for himself and, he believed other aircrew employed by the operator. He said `most smells and odours were considered the normal environment of the day to day operation of the BAe 146' and he would have discontinued his NOTOP diagnostic action if he perceived a flight hazard issue and would have reverted to the emergency checklist action. He also stated that the copilot, being on oxygen, could confirm a successful isolation procedure by occasionally removing his mask and comparing pure air with the ambient air of the flight deck.

This was at variance with the emergency checklist for SMOKE/FUMES/FIRE IN COCKPIT/CABIN. This list takes priority over any other action. The first item on the checklist is `Oxgen masks and goggles...Flight crew don, check 100%'. The imperative in relation to fumes events was also highlighted in an all operator message (AOM) from the manufacturer, which states in part `pending the definition of any necessary corrective actions, oil leaks and cabin/flight deck smells must be regarded as a potential threat to flight safety and not just a nuisance'.

The copilot completed fault isolation checks that appeared to improve the air quality on the flight deck. The PIC then asked the FAs if they could come to the flight deck so that he could better assess the situation in the cabin. The FAs came forward in turn, opened the flight deck door and entered. This action was at variance with the operations manual actions for flight attendants in the event of smoke/fumes in the cabin. The manual stated that, in the event of smoke/fumes, the FAs were to inform the PIC via the intercom and were not to open the flight deck door.

In his original report, the PIC stated that `each time when they opened the flight deck door, we noticed that the odour intensified'. The FAs' cabin crew reports to the operator also stated that the odour and fumes were still evident in the passenger cabin during the remainder of the flight. The PIC described the odour to be unlike any odour previously encountered and then decided the safest option was to return to the departure airport.

The incident operating crew underwent medical examinations that evening as directed by the operator after the event. The PIC stated that the medical practitioner they visited told him she knew very little about the effects of odours on crew and was unaware of any specific blood testing requirements for such an event. Medical testing information printed by the aircraft manufacturer in Service Information Letter (SIL) 21/45 Issue Number 1, dated January 2001, details specific test requirements. The PIC stated that it was sometimes difficult to find a medical practitioner at short notice (especially late at night) who was familiar with the required testing procedures.

After advising the medical practitioner that he was unsure if he would be able to work the following day, the PIC was given a medical certificate excusing him from flight duties for the following 24 hours. Even though he donned his oxygen mask, the copilot was similarly affected and was also excused from flight duties for the same period as the PIC. The operator reported that the remaining crew did not exhibit any residual effects from the incident.

A maintenance investigation by the operator included compliance with the latest airworthiness directive and service bulletins. All engines and the auxiliary power unit were checked. The airconditioning regenerative ducting and the delivery ducting to the rear cabin were also dismantled and inspected. That investigation determined that the number-3 engine was the likely source of the fumes and the engine was changed. The aircraft was returned to service with subsequent operating crews reporting no further fumes problems.