The crew lost directional control of the aircraft on the runway after landing due to thunderstorm-related strong wind and reduced visibility in heavy rain.

The crew did not refer to the on-board weather radar during final approach.

The crew was not provided with relevant information regarding the Emerald weather of which BoM and Airservices were aware.

BoM processing of METAR and SPECI information, along with Airservices procedures for passing operational information to flight crew during flight, reduced the likelihood that SPECI reports would be transmitted to the crew.

The recorded flight data indicated that the influence of the left to right lateral acceleration immediately prior to, and just after, touchdown was sufficient to overcome the countering influence of the aircraft's orientation and the rudder position. Information regarding the prevailing weather conditions, along with the recorded flight data, indicated that the aircraft encountered significant windshear immediately before touchdown. The crew were unable to prevent the aircraft from touching down prior to their intended touchdown point. The windshear was accompanied by very heavy rainfall which reduced visibility to such an extent that the crew lost visual reference with the runway centreline and were unable to maintain directional control of the aircraft. Both the windshear and the heavy rain were characteristic of thunderstorm activity. The wet runway would have reduced the coefficient of friction between the aircraft's tyres and the runway surface.

The crew was surprised by the sudden and significant change in the weather conditions at the time of landing. Had they been aware of the proximity of the thunderstorm to the aerodrome, they would probably have delayed their landing until conditions improved. Against that background, it is relevant to discuss the information the crew was aware of regarding the thunderstorm.

1. The 1827 SPECI that the crew received at 1831 indicated that the Emerald weather had changed significantly from that reported in the 1800 Emerald METAR. Subsequent events indicated that the crew took a number of steps to gain further information regarding the Emerald weather.
2. The dark night conditions and absence of lightning activity when the aircraft was on final approach prevented the crew from visually assessing the location, size, and movement of the thunderstorm. The crew therefore had to rely on other sources of information regarding the storm.
3. The crew confirmed the presence of weather, consistent with a thunderstorm, near Emerald from the aircraft's weather radar, which provided an explanation for the change in conditions advised in the 1827 SPECI.
4. The dark conditions may have limited the value of the company agent's assessment regarding the movement of the thunderstorm. Nevertheless, the agent's advice that the storm would pass west of the aerodrome probably contributed to the crew's decision to continue the approach.
5. The crew's report that they could see the runway lights throughout the approach indicated that there was little, if any, rain or cloud between the aerodrome and the aircraft during that period. Those conditions would have strongly reinforced the crew's apparent perception that the thunderstorm posed no hazard to the progress of the flight.

Considering what the crew knew about the weather conditions at Emerald, their decision to continue the approach was reasonable. However, BoM and Airservices were aware of additional information about the weather conditions at Emerald, and further information was probably available from the aircraft's onboard weather radar. This additional information was potentially significant in the context of the flight crew's decision-making.

1. Despite the lack of detailed information available to the investigation regarding the speed and direction of movement of the thunderstorm, it seems very likely that it was close to the aerodrome during the aircraft's final approach. Therefore, it is possible that additional use of the aircraft weather radar during that period may have enabled the crew to better assess the conditions ahead.
2. Had the crew received all the SPECI reports as transmitted by the Emerald AWS (including all the wind gust data), their awareness regarding the significance of the thunderstorm activity may have been greater. In turn, that may have prompted them to seek more information, for example by reference to the weather radar, and/or amend their arrival procedure.
3. Because the controller did not have visibility or cloud height data for Emerald (the two parameters that generally determine alternate minima), the information available to the controller to assess whether the SPECI reports justified a hazard alert was incomplete and reduced the likelihood that a hazard alert would be issued. Even if visibility and cloud height information was available, the additional information the controller required to determine if a weather report such as a SPECI indicated that conditions had deteriorated below the alternate minima was not easily accessible, and reduced the likelihood that a hazard alert would be issued.
4. BoM removed from SPECI reports data on wind gusts that were not 10 kts or more greater than the mean wind speed. Therefore, information was not available to the air traffic controller, which may have influenced the controller's decision whether or not to pass the SPECI reports to the flight crew.
5. The 1842 and 1846 SPECI reports indicated a rapid and significant change in wind direction, characteristic of thunderstorm activity. The MATS requirement that a hazard alert should only have been issued if conditions were unexpected and critical reduced the likelihood that such SPECI reports would be passed to flight crew, even though there are occasions when such information constitutes a safety hazard. For example, if the wind had backed after 1827 rather than veered, the aircraft would have encountered a significant tailwind during the landing.
6. At 1852, four minutes after the crew advised that they were changing to the Emerald MBZ frequency, and coincident with the aircraft's estimate for arrival at Emerald, Airservices received a SPECI report which indicated that the wind had veered a further 20 degrees and the gust strength was 31 kts. The controller did not pass that SPECI report to the crew. The MATS requirement that hazard alerts be directed to aircraft maintaining continuous communications with ATS reduced the likelihood that the SPECI report would have been passed to the crew. The crew were still monitoring the area frequency, even though they had earlier advised that they were changing to the Emerald MBZ frequency.

The occurrence resulted in minor damage to the aircraft. Had the aircraft encountered the windshear slightly earlier in the approach, at a point where there was insufficient altitude available for the crew to establish a positive rate of climb and go around, the consequences of this occurrence could have been far more serious.

The recorded rainfall rate was within the BoM classification of `violent'. Visibility in such conditions would have been very limited and the benefit of windscreen wipers in such conditions was probably marginal at best, as indicated by the loss of visibility reported by the crew when they encountered the heavy rain. Therefore, the failure of the first officer's wiper when high wiper speed was selected was unlikely to have contributed significantly to the loss of directional control.

Flight crew are generally in the best position to assess the significance of operational information such as weather reports. However, currently the responsibility for assessment and transmission of such information remains with air traffic controllers. This situation is likely to exist until advanced technology systems for the provision of operational information direct to flight crews are available. One such system, an airborne internet communication system, is part of the US National Aeronautics and Space Administration-led Small Aircraft Transportation System project.

On 1 May 2003, the de Havilland Canada Dash 8-200 aircraft was conducting a scheduled public transport flight from Brisbane to Emerald, with a night landing at Emerald. The crew was aware of a thunderstorm near Emerald, but did not refer to the weather radar during final approach. As the captain flared the aircraft for landing, the crew lost virtually all external visual reference as the aircraft encountered very heavy rain. After the aircraft touched down at 1856:17, the right main gear and nose gear departed the right side of the runway.

The captain regained the runway, but was unable to identify any markings associated with the runway centre line, and the aircraft departed the left side of the runway, with all three landing gear leaving the sealed surface. The crew subsequently returned the aircraft to the runway. The crew was unable to maintain the aircraft on the runway after landing because of thunderstorm-related reduced visibility and strong wind. Between 1842 and 1851, the Emerald automatic weather station issued three special weather reports (SPECIs) that indicated rapidly changing conditions at Emerald aerodrome. Flight crew could not access the SPECI reports directly. The crew relied on the air traffic controller to inform them of any relevant operational information such as SPECI reports, however the controller did not provide any of the three SPECI reports to the crew. Air Traffic Service hazard alerting procedures reduced the likelihood of the controller passing the SPECI reports to the flight crew.

Damage to the main rotor system was consistent with both main rotor blades having failed in upward bending overload, in excess of design limits, and the main rotor diverging from its normal plane of rotation and contacting the tail boom and canopy.

Examination of the helicopter and its transmission and rotor systems found no evidence to indicate any pre-existing defect that could have contributed to the in-flight break-up. Witnesses reported hearing the engine running before and after the break-up occurred.

The pilot was seen waving just moments before the helicopter broke up. The investigation was unable to determine if flight control input by the pilot or passenger, or lack of corrective control, had contributed to the development of the accident. Although either low rotor RPM or abrupt manoeuvring can result in air loads on the blades exceeding their design limit, the reason for the excessive upward bending of the blades could not be determined.

The NTSB special investigation report NTSB/SIR-96/03 - Robinson Helicopter Company - loss of main rotor control accidents, which analysed accident data from 31 fatal accidents, concluded that in the absence of any evidence of defects or component failures, other possible factors such as the sensitivity and responsiveness of the helicopter's flight controls combined with limited pilot skills, proficiency, or alertness, be considered. Although that report concerned a different helicopter type from the Safari, its conclusions were directed to all lightweight helicopters with sensitive and responsive controls, characteristics shared by both types.

The installation of a governor and an aural low rotor RPM warning, as noted in the NTSB special investigation report NTSB/SIR-96/03 - Robinson Helicopter Company - loss of main rotor control accidents, had contributed to the greatly reduced incidence of low rotor RPM related accidents in that helicopter type.

Sequence of events

The pilot and a passenger of a kit-built Canadian Home Rotors Safari (Safari) helicopter, registered VH-VDB, operated as an experimental aircraft¹, were making a private flight from the pilot's property to a nearby airstrip.

At about 1415 Central Summer Time, witnesses reported seeing the helicopter flying in a south-westerly direction. The Australian Transport Safety Bureau did not conduct an on-scene investigation of this occurrence. The report presented below was prepared principally from information supplied to the Bureau.

One witness, who knew the pilot, reported that he saw the helicopter commence a gentle left turn and the pilot waving² to him moments before it broke up in flight. That witness also reported seeing the cockpit bubble shatter and a cloud of white dust appearing from the area of the cabin. Witnesses reported having heard the engine operating immediately prior to and following the in-flight break-up. Other witnesses heard a loud metallic sound, and described the helicopter almost stopping, pitching nose-up and a main rotor blade folding, before pitching nose-down and descending steeply. The helicopter collided with a large tree and a shed before impacting the ground at the rear of a residential house block. There was no fire. Both occupants were fatally injured.

Wreckage information

The main wreckage consisted of the forward part of the helicopter, main rotor and transmission assembly and the forward section of the tailboom. The tail rotor, its gearbox and the rear part of the tail boom had lodged in the roof of the shed beside the main wreckage. The rear section of the damaged tail rotor drive shaft was lodged in a tree.

Debris from the helicopter was distributed over a distance of approximately 100 m. Components from the tailboom, tail rotor drive system, fragments of acrylic windshield and items from the cockpit were found along the wreckage trail. A damaged section of the windshield exhibited rotational scuff marks consistent with a main rotor blade strike.

Evidence at the accident site was consistent with the main rotor blades being stationary at impact. Paint transfer and deformation of the tailboom tubing and corresponding damage to the leading edges of the main rotor blades, were consistent with the tailboom being severed by the main rotor blades at the second rearmost bay. A ballast weight was later found approximately 150 m northwest of the main wreckage site.

Examination of the engine did not reveal any sign of mechanical failure. The carburettor, magneto and ignition harnesses were damaged in the impact sequence and could not be tested. Environmental conditions at the time of the occurrence, when plotted on the Carburettor Icing Probability Chart, were in the moderate zone for cruise power settings.

Pilot information

The pilot held a private pilot licence (helicopter) endorsed with the Robinson R22 helicopter type. Records held by CASA showed that the pilot's class 2 medical certificate had expired on 31 January 2002. The pilot had flown an estimated total of 308 hours of which 241 hours were on the Safari helicopter. Postmortem and toxicology reports did not find any condition that would have affected the pilot's performance.

Examination of components

The flight control system was examined for integrity. Although many of the control tubes were bent and fractured, they were all correctly attached. A friction device was fitted to the pilot's collective control lever. All fracture surfaces, when examined, exhibited failure in overload, consistent with impact forces. One of the high-density plastic 'droop stops', found separate from the main wreckage, exhibited damage that indicated a significant in-flight main rotor grip impact.

The main rotor transmission, main rotor mast and main rotor assembly, including the main rotor blades, were removed for examination at the ATSB's technical laboratory. None of those components exhibited any identifiable pre-existing defect. However, during examination it was noted that neither of the pitch control horns had the indexing ball bearings installed. A detent, machined into each blade grip, allowed a ball bearing to align with a slot machined into the clamping surface of the pitch control horn. That was to facilitate accurate alignment of the flight controls during rigging. Although the reason for the missing pitch control horn indexing balls was not determined, the kit supplier advised that their use was not essential.

Both main rotor blades exhibited extensive permanent upward bending. The blades were constructed of an extruded aluminium leading edge spar to which was bonded an upper and lower carbon fibre skin. The area between the upper and lower skin behind the spar was unfilled, except for the trailing edges which were bonded with an expanded foam adhesive. The deformation of both blades was similar, although the leading-edge extrusion of one blade had fractured at two locations. Examination indicated that bending was consistent with both blades having been subjected to excessive air loads while the rotor system was powered. The examination concluded that when bending of the blades exceeded normal design limits, the rigid, less ductile carbon fibre skins, separated. The upper skins separated almost instantaneously, releasing the cloud of fine dust-like particles from the disruption of the foam bonding material.

Subsequent examination of the blades by an independent composite structure specialist found that the carbon fibre surfaces had separated from the spar due to shearing forces produced by excessive upward bending. The blades did not exhibit any pre-existing manufacturing defect or delamination failure in the composite material of the blades.

Centre of Gravity

The Centre of Gravity (C of G) of the helicopter was calculated to be within normal operating limits.

Experimental aircraft

Civil Aviation Regulation CAR (1998) Part 21.191 (g) permitted construction and operation of the kit-built Safari helicopter as an experimental aircraft solely for the education and recreation of the constructor. As the constructor of the major portion of the helicopter, the pilot was considered to be the manufacturer and was responsible for the acceptance and use of all the helicopter's components. That permitted the use and modification of components without the need for lengthy and expensive development normally associated with certification.

The helicopter kit supplier did not approve any modification or alteration to the Safari helicopter or its components, other than those made by the factory. The kit supplier reported that following any factory modification to the helicopter, an extensive program of hover and flight testing was undertaken before acceptance. They also recommended that when constructors made modifications or substituted components, a similar testing program to that used by the manufacturer should be adopted.

The pilot was reported to have modified and replaced some components that had been supplied as part of the kit. Although the kit supplier reported that many of the changes had been made in consultation with them, the absence of a complete set of component modification records for the helicopter did not allow the investigation to authenticate information about those modifications or the testing of them.

Experimental aircraft were not required to have a Flight Manual, but it was expected that sufficient information for the safe operation of the aircraft was available to the pilot. A copy of a Safari flight manual was with the pilot's documentation provided to the investigation team. The manual was not identified as being specific to VH-VDB and the accuracy of information contained in it could not be validated. The provision and accuracy of information relating to the operation of the helicopter was the responsibility of the constructor.

Main rotor speed

Main rotor speed was one of the critical factors that determined the amount of lift generated by the main rotor of a helicopter. Engine power provided the force to overcome the drag of the rotor blades and attain a specific main rotor RPM. The pilot controlled engine power via a twist grip throttle control on the collective control lever. Although a mechanical linkage, described as a correlator, automatically adjusted engine power to approximate the power requirement for the collective control position, the pilot was required to make minor adjustments with the twist grip throttle. Despite the installation of a collective lever friction lock, the helicopter was not fitted with a governor that automatically maintained engine RPM and therefore main rotor RPM.

The rotor speed limits in the pilot's copy of the Flight Manual for the Safari were a maximum of 520 RPM and minimums of 475 RPM (power off) and 450 RPM (power on). The Flight Manual cautioned pilots that catastrophic rotor stall occurred if rotor RPM dropped below 390 RPM. Low rotor speed could result in excessive rotor 'coning' and if uncorrected, would ultimately lead to blade failure from excessive upward bending. In the event of a low rotor RPM, the kit-supplier reported that 'A strong engine will allow recovery. Lowering collective will lessen the time required. Using less right pedal will also help.'

The pilot was reported to have developed a main rotor RPM warning system, after having experienced difficulty monitoring the kit-supplied dual-pointer, combined rotor and engine RPM gauge. The system was described as providing both aural and visual warnings. Visual warnings were provided by an indicator, installed on the centrally mounted instrument panel. A red light at the top of the display indicated rotor overspeed and an amber light positioned at the bottom of the display indicated rotor underspeed. Four green lights arrayed across the centre of the instrument illuminated sequentially to provide a rotor RPM deviation trend. A single tone was generated to indicate both overspeed and underspeed rotor conditions. The warning indicator was too badly damaged to determine its operating status. The damaged sender unit from the warning indicator was tested and found to provide a low rotor speed signal at 459 RPM and an overspeed rotor signal at 494 RPM.

A Flight Manual limitation stated 'Use maximum power-on RPM [520 RPM] during take-off, climb or level flight below 500 ft AGL or above 5,000 ft density altitude'. That had the potential of a main rotor overspeed warning remaining continuously activated while a pilot adhered to the recommended rotor RPM during the above-mentioned phases of normal flight.

Maintenance

As the constructor of the helicopter, the pilot was approved to carry out his own maintenance. The kit supplier reported that the pilot had received maintenance training on the Safari helicopter from the kit supplier's test pilot when he attended the constructor's property to conduct the initial flight tests. The pilot was reported to have performed his own maintenance, but no maintenance documentation, including the current maintenance release, was found. The kit supplier reported that the pilot seldom asked about the Safari maintenance but would contact them for information when needed. A Rigging and Balance Manual for the Safari was available to owners. That manual contained information about the flight control rigging and balancing and tracking procedures for the main and tail rotors.

The main rotor blades were manufactured in the US by a separate manufacturer and sold to the pilot by the Canadian kit supplier. The pilot's family reported that the main rotor blades were obtained as part of the helicopter kit. When unpacked, one of the blades was found to have a 2 mm depression on the upper and lower surfaces at the mid-chord position along most of the blade length. The pilot had contacted the kit supplier about the problem and the blade manufacturer subsequently advised that the increase in air pressure, between the pilot's near sea-level property and the significantly higher elevation of their US facility, had resulted in the slight depression. The blade manufacturer advised the kit supplier that the pilot should drill a small hole in the outer end of the blade to allow the pressure to equalize and then fill the hole. The pilot was reported to have performed the repair according to the instructions. Although the kit supplier had offered to replace the blades with a matched pair, the pilot subsequently used the repaired blade.

The pilot had recently replaced the pitch control bearings of each blade due to excessive wear. That required the removal and refitting of the main rotor blades. As noted in the Examination of Components, the indexing ball bearings had not been installed. The investigation was unable to determine if any misalignment between the pitch control horns had existed before the in-flight break-up occurred.

The pilot had reported to other constructors that he had experienced vibrations associated with the main rotor with either a near full fuel load, or with low fuel quantities. During the previous three months, four of the flight entries in the pilot's notebook were annotated as 'track and balance'. The pilot was reported to have also weighed and balanced the blades in an attempt to eliminate a possible source of vibration. A family member reported that the pilot was attempting to eliminate all the vibrations, not because of any risk to flight safety, but in order to satisfy a personal ideal. That family member reported that during a flight in the helicopter earlier that week it had flown normally. A damaged hand-held vibration analyser unit was found in the wreckage. When examined, the unit contained no stored data. No relevant record of track and balance data was found. It could not be determined if there had been any tracking and balance testing during the accident flight.

NTSB Accident study of light weight helicopters

In 1996 the United States of America (US) National Transportation Safety Board (NTSB) released special investigation report NTSB/SIR-96/03 - Robinson Helicopter Company - loss of main rotor control accidents. The report reviewed 31 occurrence reports where loss of main rotor control was identified but where no evidence was found of the specific event that caused or allowed the main rotor blades to diverge from their normal plane of rotation and strike the airframe. The report stated that the sensitivity and responsiveness of the helicopter's flight controls when combined with limited pilot skills, proficiency, or alertness could have been a factor in some of the 31 accidents the NTSB reviewed. Although the NTSB report was solely concerned with the R22 helicopter type, it noted the need for continued research to study flight control systems and main rotor blade dynamics in light weight helicopters with highly responsive controls.

The report also noted that in March 1995, the US Federal Aviation Administration (FAA) technical panel recommended that the R22 helicopter type be reconfigured with an electronic engine RPM governor. The report further recommended that in normal flight operations, switching the governor off should be prohibited. The manufacturer adopted those recommendations and fitted governors to their fleet. The NTSB report cited that those measures were some of the many that had subsequently resulted in a reduction to the R22 accident rate.

¹ Civil Aviation Safety Authority (CASA) Advisory Circular 21.1(1) Aircraft Airworthiness Certification - Categories and Designations Explained stated that:

Experimental aircraft by their very nature are not type-certificated. Experimental is not a category per se, rather it is a designation. It is also important to note that an experimental certificate does not attest to an aircraft being fully airworthy, despite being grouped under the special Certificate of Airworthiness (CoA). The Experimental certificate system replaces some elements of airworthiness control previously covered by Civil Aviation Regulation 134 Permit to Fly, and Civil Aviation Order 101.31 developmental category CoA.

² The helicopter was flown from the left seat. When seated on the left side of the helicopter, the pilot's left hand operated the collective control and twist-grip throttle. The cyclic control was operated with the other hand and unlike the collective control, generally required constant input.
³ The Safari helicopter had a relocatable 6.3 Kg ballast weight consisting of a lead filled stainless steel tube. The ballast weight was positioned according to the helicopter loading, to maintain the Centre of Gravity within permissible limits. Mounts were located on the forward section of the right landing skid and on the tail boom tubing in the second bay from the rear. When not required for helicopter trim, the ballast weight could be carried in the cabin.
⁴ Bureau of Air Safety Investigation, Asia Pacific Air Safety Journal No. 22, June 1999.
⁵ Coning is the upward bending of the blades caused by the resultant forces of lift and centrifugal force. At rest, the main rotor blades have a preset Coning Angle. As rotor RPM increases, the coning angle reduces due to the centrifugal force exerted on the blades. When weight is transferred to the blades, the lift produced by the blades increases, increasing the coning angle. Within the range of normal operating RPM, centrifugal force acting on the blades prevents excessive coning.
⁶ Civil Aviation Regulation 42ZC(6) of CAR 1988 permitted authorised persons to conduct maintenance on the aircraft and engine. CASA Instrument 545/00, effective from 1 January 2001, gave that permission to constructors of kit built aircraft who fabricated and assembled the major portion of that aircraft and if the aircraft was used solely for that person's own education or recreation.

The damage to the rod end was consistent with failure during the retraction of the landing gear after take-off at Gove. Disconnection of the left main gear from its push-pull tube meant that the gearbox was unable to extend or retract that gear. It also meant that the emergency gear extension was ineffective. Fatigue cracking within the rod eye section was the principal factor behind the separation of the push-pull tube and the subsequent failure of the landing gear to operate correctly.

The investigation was unable to determine why the rod end failed. It is possible that an increase in transmitted loads resulting from excessive system friction, system rigging problems or the failure of interrelated components could have contributed to the initiation of the rod end cracking.

The pilot was confronted with the high workload of maintaining control during night circuits and trouble shooting a landing gear malfunction. The night conditions meant that people on the ground were of limited help, at least initially, in determining the status of the landing gear. Landing the aircraft with the gear retracted allowed for some directional control during the landing slide and probably limited damage to the aircraft.

The release of the emergency exit on final approach had the potential to inflict serious damage to the tailplane with possible control problems resulting.

The pilot of a Cessna 310R made a night departure from Gove, NT, at 1900 Central Standard Time for Groote Eylandt, NT. After take-off he noticed two unusual thumps near the end of the landing gear retraction cycle. Normal landing gear up indications were observed and there were no unusual handling characteristics during the flight to Groote Eylandt.

The pilot reported that on arrival in the circuit area he selected the landing gear down. The landing gear operated but the left main landing gear down light did not illuminate at the end of the cycle. The pilot changed the down light bulb without change to the indication. He then cycled the landing gear by selecting up then down, which resulted in the left main and nose landing gear down lights not illuminating. The pilot spoke to ground personnel by radio and made a few low passes along the runway to allow them to observe and report on the condition and position of the landing gear. However, they were unable to see the position of the landing gear due to the darkness.

The pilot manoeuvred the aircraft in an unsuccessful attempt to free the landing gear. Further cycling of the landing gear and a manual landing gear extension were attempted without success. Emergency services, notified by air traffic services, attended and were able to provide enough illumination of the aircraft for observers to report that the nose gear was not extended, the left main gear was at an angle and the right gear appeared to be in the down position.

After consultation with a pilot on the ground and briefing the passengers, the pilot made an approach to runway 28 with the gear selected up and the flaps down. While on final approach the pilot unlatched the emergency exit and cabin door. The emergency exit, which consisted of the pilot's side window and associated frame, detached from the aircraft.

The aircraft made a smooth touchdown and slid on the runway. The pilot selected the mixture control to idle cut off and the fuel, magnetos, alternators and battery to off. The passengers exited through the cabin door and over the right wing and moved away from the aircraft. There was no fire, but the propellers and the underbelly of the aircraft were damaged.

Inspection of the aircraft by engineering personnel revealed that the rod end on the outer end of the left main gear inboard push-pull tube had separated, effectively disconnecting the left main landing gear assembly from the actuating mechanism.

Laboratory examination by the ATSB revealed that rod end separation had occurred under predominantly tensile forces after fracturing through one side of the eye section. Detailed examination of the fracture surfaces found characteristic evidence of fatigue cracking, originating from the outer corners of an integral lubrication port within the eye body. The examination found no evidence of any material or manufacturing defects having contributed to the failure.

There were no reports of previous landing gear problems.

The electro-mechanical landing gear system utilised a gearbox, driven by an electric motor, to turn two bellcranks that extended or retracted a push-pull tube to each landing gear assembly. A landing gear hand-crank provided an alternative manual means to drive the gearbox.

In the sub-section titled `EMERGENCY LANDING PROCEDURES', the Cessna 310R pilot's operating handbook (POH) specified unlatching of the cabin door prior to flare-out. However, there was no reference to in-flight unlatching of the emergency exit in any of the emergency checklists.

1. The pilot was not recently experienced on the occurrence helicopter type.
2. The pilot in command allowed the helicopter to move laterally during the lift-off to the hover.
3. The pilot in command did not raise the helicopter to a hover height sufficient to prevent contact with the mobile platform.

Post-occurrence technical examination of the helicopter did not reveal any evidence of an airframe, engine or system fault that may have contributed to the accident. In addition, examination of the mobile platform did not reveal any evidence of it having moved throughout the rollover sequence.

The circumstances of the accident are consistent with the phenomenon known as dynamic rollover. Scrape marks from the helicopter's right skid were found on the lip along the right side of the mobile platform. That indicated that the helicopter was not raised to a height sufficient to clear the platform in the event of lateral movement. There was no wind reported at the time of the occurrence that could have contributed to the lateral movement. The pilot had extensive flying experience and normally flew a mix of different types, including a mix of European and North American types. Due to that experience, the investigation considered that confusion with respect to correct flight control input to control yaw was unlikely.

If the pilot had prevented the lateral movement of the helicopter during the lift off to the hover, and had raised the helicopter to a hover height sufficient to clear the platform, dynamic rollover would most probably have not occurred. Therefore, the investigation considered that the design or use of the mobile platform was not a factor in the occurrence.

The pilot had not flown a Bell 47 type helicopter during the preceding three months, and had not previously flown a turbine-powered Bell 47. It is likely that the pilot's lack of recency in the helicopter type, combined with his not having flown a turbine-powered Bell 47 previously, contributed to his:

• not making sufficient flight control input to correct the right lateral movement during the lift-off to the hover
• not raising the helicopter to a hover height sufficient to prevent contact with the platform.

The pilot's injuries were consistent with him being struck by the main transmission assembly as it separated from its mount as a result of the dynamics associated with main rotor ground contact.

History of the flight

At approximately 0830 EST, the pilot of the Bell 47G-4A turbine-powered (Soloy) helicopter, registered VH-MTX, was conducting a lift-off to the hover from a mobile helicopter landing site (HLS)¹ at Caboolture aerodrome, when the helicopter rolled onto its right side. Weather conditions at the time of the occurrence were reported to be '…little or no wind, warm and humid, some cloud but clearing.' The helicopter was substantially damaged and there was no post-occurrence fire. The pilot, who occupied the left seat², was fatally injured and the passenger, who occupied the right seat, sustained minor injuries.

Pilot in command

The pilot held an Airline Transport Pilot (Helicopter) Licence, a Commercial Pilot (Aeroplane) Licence, a Command Multi Engine Instrument Rating (CMEIR) (Helicopter), a CMEIR (Aeroplane), and a Grade 1 Instructor (Helicopter) Rating. According to his pilot flying logbooks, he had accumulated approximately 8,293 hours total flying experience, of which approximately 7,180 hours was on helicopters, including 14.8 hours on the Bell 47G helicopter type. He had flown 2.0 hours in the last 30 days and 15.7 hours in the preceding 90 days. His last flight prior to the occurrence flight was nine days previously in a Bell 206 type helicopter. He had last flown a Bell 47G type helicopter on 19 December 2002, including taking off from and landing back on the mobile platform involved in the occurrence. He had not previously flown a turbine-powered Bell 47G helicopter.

The pilot was endorsed to fly Bell 47G type helicopters in accordance with Section 40.3.0 of the Civil Aviation Orders (CAO). In accordance with paragraph 3.3 of those orders, he was permitted to fly turbine-powered Bell 47G helicopters without further endorsement. The operator also reported that, in accordance with paragraph 3A.4 of CAO 40.3.0, prior to the occurrence flight, the pilot was offered refresher training in the turbine-powered Bell 47G type, however the pilot declined that offer.

The pilot met the recency requirements of Civil Aviation Regulation 5.178, having completed CMEIR (Helicopter) and Instructor (Helicopter) Rating renewals on 13 May 2002, in a Bell 412 helicopter type.

Due to the nature of his employment, the pilot had flown 10 aircraft types in the preceding 12 months, of which 7 were helicopters. Those helicopter types included a mix of European and North American types, a mix of single engine and multi-engine types³, and a mix of turbine and piston engine types. Most of the pilot's helicopter flying experience was in turbine engine types. The pilot was experienced in operating helicopters from mobile platforms.

Passenger

The passenger was employed by the same organisation as the pilot in command. He was an experienced helicopter pilot, who reported that he had accumulated approximately 4,700 hours total flying experience, of which approximately 4,500 hours was on helicopters. He reported that he was not in current flying experience. According to witnesses, he did not have his hands or feet near the flying controls during the occurrence.

Medical information

The pilot's aviation medical certificate was valid and carried a restriction for him to have available reading vision correction. A review of his medical records, investigation interviews, results of the post-mortem examination and toxicological testing, found no evidence of pre-existing medical conditions or the presence of any substance that may have influenced his performance.

Survival information

Four-point restraint harnesses with inertia reel shoulder straps were fitted at the pilot and right side passenger positions. Post-occurrence technical examination revealed that they were firmly secured to their mounts, and the inertia reels appeared to operate normally. The pilot occupied the left seat and remained restrained during the rollover sequence. He sustained severe impact injuries to the rear of the upper torso and lacerations to the back of the head.

Helicopter information

Type: Bell 47 helicopter
Model: 47G-4A (Soloy)
Registration: VH-MTX
Serial Number: 7765
Year of manufacture 1971
Engine: Rolls Royce 250-C18
Total time in service: Approximately 5,570 hours
Maintenance release: Number 09425 issued 14/03/03 at 5,549.7 hours

The helicopter had been imported from Japan in October 2002, receiving an Australian Certificate of Airworthiness in December 2002. The helicopter was maintained in accordance with the manufacturer's and CASA approved documents and schedules, and had flown approximately 20 hours since its last 100-hourly inspection.

Damage to the helicopter

The helicopter came to rest on its right side, with its right skid landing gear resting on the right rear corner of the platform. The canopy bubble had shattered. The main rotor, mast and transmission assembly had detached from the helicopter as a single unit and was located next to the helicopter. The 'white' main rotor blade⁴ was lying across the cockpit. Evidence, in the form of multiple main and tail rotor impact marks, was observed in the ground with an area of burned grass beneath the helicopter's engine exhaust. One of the tail rotor blades had detached and was located approximately 30 metres from the helicopter. The tip of the 'red' main rotor blade was found approximately 150 metres to the east of the helicopter. Both arms of the main rotor stabiliser bar had fractured and were found within 20 metres of the helicopter. All major components were located and identified at the site. Post-occurrence technical examination of the helicopter did not reveal any evidence of an airframe, engine or system fault that may have contributed to the accident.

Mobile platform information

The mobile platform was a flat-based metal construction on wheels, measuring 365 x 300 cm and was approximately 22 cm above the ground. A 5 cm high metal lip ran along both sides of the platform. There were no metal lips at the front or rear of the platform. The front wheels were positioned forward of the base and were approximately 27 cm in diameter. The rear wheel axles were in line with the platform base, allowing the 38 cm-diameter wheels to sit approximately 18 cm above the base. There were two prominent worn strips along the base approximately 38 cm from the side lips, which indicated the usual positioning of the helicopter skids on the platform. At the time of the occurrence, the platform was connected to a small tractor. The tractor brake was engaged and there was no evidence of the tractor having moved throughout the accident sequence.

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 prevailing weather conditions), the aircraft can land at, or take-off from, the place in safety.

GEN 2.2 of the Aeronautical Information Publication, defines an aerodrome as:

A defined area of land or water (including any buildings, installations and equipment) intended to be used either wholly or in part for the arrival, departure and movement of aircraft.

Guidelines for the establishment and use of HLS are at Civil Aviation Advisory Publication (CAAP) 92-2 (1), which defined an HLS as:

…a place that may be used as an aerodrome for the purposes of landing or taking off of helicopters.

The definition of '…place…' in CAAP 92-2 (1) included '…on a structure…'.

There is no CASA regulation or guidance concerning the design or use of mobile platforms. Additionally, international standards and recommended practices contained in Annex 14 to the International Civil Aviation Organization (ICAO) Convention on International Civil Aviation, Aerodromes, Volume II - Heliports, do not refer to mobile platforms. A search of international regulatory authorities found no documentary guidance regarding design or use of mobile platforms.

The investigation found that mobile platforms of varying design are used throughout the Australian civil helicopter industry. Some of those designs incorporated side lips and others had no protrusions above the platform surface. Most of the platforms were tractor-towed.

Damage to the mobile platform

Examination of the mobile platform revealed two fresh scratch marks on the right lip approximately 50-85 cm from the rear of the platform. Also evident were fresh gouges on the rear edge of the platform, adjacent to the right wheel axle, which indicated that the helicopter had been moving rearwards as it contacted the lip. Examination of the platform did not reveal any evidence of it having moved throughout the rollover sequence.

Organisational information

Pilot employees from the same organisation as the occurrence pilot reported that they each received funding for up to 40 flying hours each financial year to maintain recency on selected aircraft types. They also reported that, due to the nature of their employment, some pilots had a requirement to maintain recency on a number of types simultaneously.

Dynamic rollover

The phenomenon known as dynamic rollover was described in helicopter textbooks, training manuals and industry and safety publications. It was included at:

• items 10.7 and 12.1 of the CASA Day (VFR) Syllabus - Helicopters, Issue 3, January 1999, which applied from Student through to Commercial Helicopter Licence standard
• item 2.1.14 of the Air Transport Pilot (Helicopter) Licence - Aeronautical Knowledge Syllabus, Issue 3, January 1999.

Dynamic rollover has been defined as:

The occurrence of a rolling motion, while any part of the landing gear is acting as a pivot that causes the aircraft to exceed a critical angle and roll over⁵.

Another definition states that:

Put simply, dynamic rollover is the result of the helicopter developing excess angular momentum about the skid in contact⁶...

Dynamic rollover typically occurs when a critical rollover angle is exceeded. That angle is dependent upon control limits and in most helicopters is in the order of 15 degrees. Accidents attributed to dynamic rollover have occurred previously on a number of surfaces, including open flat grassed surfaces.

1. The aircraft was being operated in weather conditions conducive to engine carburettor icing.
2. The aircraft maximum take-off weight limitation was exceeded for the flight.
3. The aircraft loaded moment envelope limitation was exceeded for the flight.
4. The aircraft departed controlled flight at a height above the ground from which aerodynamic stall recovery would have been unlikely.

Loss of engine power under normal loading conditions causes the aircraft nose to pitch downward for aerodynamic stall recovery because of the aircraft designed forward centre of gravity (C of G). Loading in a tail-heavy direction or rearward C of G condition has a most serious effect upon longitudinal stability, affecting the aircraft's ability to readily recover from stalls and spins. As the C of G moves rearward, a less stable condition occurs, which decreases the ability of the aircraft to right itself after manoeuvring or after disturbances by gusts. The aircraft as loaded had not exceeded the rearward C of G limitation. Although the rearmost C of G limit had not been exceeded, the location of the C of G just 25.4 mm ahead of that limit, meant that the aircraft exhibited a rearward C of G condition.

The aircraft exceeded both the aircraft manufacturer's Maximum Take-off Weight (MTOW) limitations and the aircraft loaded moment envelope. Exceeding the aircraft MTOW limitation may adversely affect flight characteristics. CAAP advisory number Number 235-1(1), advised pilots against using standard weights and recommended weighing occupants and baggage in order to prevent exceeding those limitations.

The examination of the engine and carburettor revealed no evidence of a preimpact failure or anomaly. The aircraft was most likely being operated with the carburettor heat set to the OFF position as indicated by the position of the heat lever at the air box on the engine. With the aircraft operating in weather conditions conducive to carburettor icing, it may have begun losing power. The onset of carburettor icing may have been insidious, as the pilot may not have noted the deterioration in engine RPM. As the aircraft was not equipped with a carburettor ice detection system, the pilot was not afforded any warning of the potential for carburettor icing. Without this warning, if the engine performance deteriorated, the pilot most likely would not have been able to apply carburettor heat in time for it to take effect sufficiently to regain full power. With the rearward C of G condition present, the pilot may not have been able to pitch the nose of the aircraft downward as required for aerodynamic stall recovery. However, in any case, the stall was most likely unrecoverable because of the low height above the ground. Weather conditions encountered on previous flights to and from the island may not have been sufficient to produce carburettor icing or may have been masked by the constant speed propeller.

The possibility also exists that wind shear and turbulence in the area, in combination with the adverse flight characteristics resulting from exceeding the aircraft's loaded moment envelope limitation, could have degraded the aircraft's controllability and resulted in the aircraft's departure from controlled flight.

While several possibilities exist as discussed in this analysis, the investigation could not conclusively determine the reason for the excessive nose-up pitch and departure from controlled flight.

The pilot of the Cessna 172G aircraft was conducting a series of charter flights between the Trefoil Island Aircraft Landing Area (ALA) and the Smithton, Tasmania aerodrome. Witnesses stated that the aircraft, with the pilot and three passengers on board, took off from the island ALA runway 28 on a west-south-westerly track at approximately 1745 hours EsuT, on the third return flight of the afternoon. Witnesses reported that the aircraft turned to the left on a southerly heading while climbing, followed by a left turn to the east. They reported that following the turn to the east, and after it had overflown the buildings on the island at approximately 200 feet above ground level, the nose of the aircraft pitched up abruptly to an angle of 30-40 degrees. According to the witnesses, following the nose-up pitch, the aircraft rolled abruptly to the left, lost altitude and descended from their line of sight. The witnesses heard the impact of the aircraft and ran to render assistance. The aircraft was destroyed by impact forces and all four occupants received fatal injuries.

Wreckage information

The wreckage of the aircraft was oriented on a heading of 191 degrees magnetic, indicating that it had rotated through about 270 degrees during the descent. The aircraft impacted the ground wings level, with a nose-down angle of approximately 39 degrees, on a downward sloping hill of approximately the same angle. There were no indications that the aircraft was in a spin at the time of impact. The cabin roof had separated at the rear attachment and both wing struts had separated. The forward cabin area had collapsed, with the tail section and the tail cone buckled and bent partially forward. Wreckage evidence indicated a high rate of vertical deceleration, in excess of 24 g (acceleration due to earth gravity, international standard value being 9.80665 metres per second squared, assumed at standard sea level), with indications of little forward airspeed.

The propeller/crankshaft assembly had separated behind the radius of the crankshaft flange. The fracture surface displayed evidence of a unidirectional bending overload failure, indicating low engine RPM at the time of the fracture. Examination of the propeller spinner and propeller blades confirmed low engine RPM at impact. The carburettor heat lever at the air box on the engine was noted to be in the OFF position. The aircraft was fitted with an elevator trim that allowed the pilot to minimise load forces on the elevator, depending on the position of the centre of gravity (C of G), airspeed and power settings. The aircraft's elevator trim system was found in the slightly nose down from the TAKE-OFF TRIM or neutral position. The seats and seat rails incurred substantial damage, but the pilot's seat end stop was still located intact on the seat rail. There were no indications of a bird strike on the aircraft.

Aircraft information

A 100 hourly inspection was completed on 17 February 2003 at 9,663.6 hours total time airframe (TTAF) with no major anomalies noted. A 50 hourly engine inspection was completed on 12 March 2003 at 9,713.6 hours TTAF and 1,085.2 hours engine time since overhaul (TSO), with no anomalies noted. At the time of the accident, the aircraft had accumulated 9,718.4 hours TTAF. The maintenance release listed no outstanding discrepancies for the aircraft and was current and valid.

Nothing was found during the investigation to suggest a mechanical failure of any part of the aircraft that could have contributed to the accident.

Engine information and examination

Supplemental Type Certificate number SA807CE was incorporated in 1977 with the installation of a 180 horsepower Lycoming model O-360-A1A engine and a Hartzell constant speed propeller, replacing the 145 horsepower Continental model O-300C engine and fixed pitch propeller. At the time of the accident, the engine, serial number L21971-36A, had accumulated 1,090 hours TSO. A technical disassembly and inspection of the engine and carburettor was completed at an independent maintenance facility under Australian Transport Safety Bureau (ATSB) supervision. The disassembly and examination did not reveal any evidence of pre-impact internal component failure or anomaly.

Pilot information

The pilot held a valid commercial pilot (aeroplane) licence and Class 1 medical certificate at the time of the accident. The pilot's last flight review was completed on 6 January 2003. Post-mortem and toxicological examination did not identify any factor that may have impaired the pilot's ability to operate the aircraft safely.

Meteorological Information

Documents recovered at the accident site included an Airservices Australia on-line weather forecast briefing for the area and for King Island and the Smithton aerodrome. The forecast was dated 14 March 2003 and the time noted was 1046 hours. Wind listed on the briefing for the 2,000 ft level was forecast as variable at 15 kts. The forecast also noted a south-westerly stream with a slow moving trough with drizzle and locally broken low cloud. King Island was located 107 km to the north of Trefoil Island. The King Island meteorological report noted light drizzle, south-south-westerly wind at 15 knots and a temperature/dewpoint spread of 5 degrees C.

A series of wind generators was located on the Tasmanian mainland at Cape Grim, approximately 5 km to the southwest of Trefoil Island. This facility periodically monitored and logged weather conditions. Documented information obtained from that facility indicated that the weather conditions at 1700 hours were: air temperature 15.4 degrees C; dew point 11.4 degrees C; relative humidity 77 percent; and wind from the southwest at 27 kts with gusts to 30 kts. Information documented at 1800 hours recorded: the air temperature 15.1 degrees C; dew point 10.6 degrees C; relative humidity 74 percent; and wind from the southwest at 28 kts, gusting to 29 kts. Relative humidity recorded from 1500 hours to 1700 hours (the estimated time of the first two return flights) was recorded as 77-79 percent. The wind recorded at that time was from the southwest and varied from 29 to 30 kts with the temperature/dew point spread from 4.0 to 5.6 degrees C. Witnesses stated that the accident occurred at 1750 hours. Last light on the day was about 2011 hours.

A pilot familiar with the area reported that a south-westerly wind often caused orographic lifting (when air is forced upwards by a barrier of mountains or hills) moving heavily laden moist air into the flight path of an aircraft departing from the island.

Carburettor and engine induction system icing

A search of the ATSB occurrence database indicated a total of eight carburettor or engine induction system icing related accidents since May 1994. These accidents resulted in two fatalities. One aircraft was severely damaged and three aircraft were destroyed. An article on the US Federal Aviation Administration (FAA) website, reprinted from Vintage Airplane Magazine and dated November 1994 stated: `According to the National Transportation Safety Board, carburettor ice was involved in over 360 accidents in the past five years. These figures do not include the unreported off-airport landings and incidents caused by icing. The results were 40 deaths, 160 injuries, 47 aircraft destroyed and 313 aircraft severely damaged.' Several of these accidents noted suspected carburettor icing at high power settings. Float-type carburettors, such at that used in the occurrence engine, are most susceptible to this event. Evidence of carburettor icing is highly perishable and dissipates rapidly.

When carburettor ice forms, it can obstruct the smooth flow of the air/fuel mixture, which results in a reduction of engine RPM, power, and an associated loss of airspeed and altitude. FAA Advisory Circular AC 20-113 provided information pertinent to aircraft engine induction system icing. It noted:

`c. Fuel Vaporization Ice- This icing condition usually occurs in conjunction with throttle icing. It is most prevalent with conventional float type carburettors, and to a lesser degree with pressure carburettors when the air/fuel mixture reaches a freezing temperature as a result of the cooling of the mixture during the expansion process that takes place between the carburettor and the engine manifold.'

The circular also noted that vaporisation icing may occur, when a relative humidity of 50 percent or higher is present, at temperatures from 0 degrees C to as high as 37.7 degrees C. It also stated that in general, when the temperature/dewpoint spread reaches 6.6 degrees C or less and a relative humidity of 50 percent or higher, there is a potential for icing.

The values from the weather observations for 1800 hours at Cape Grim were plotted on a carburettor icing probability chart. The temperature/dewpoint spread was 4.5 degrees C. The plot was located in the area of the chart labelled `serious icing- any power setting'.

Mitigating the effect of carburettor icing involves pilot action to apply full carburettor heat (the ON position), which initially causes a further loss of power (perhaps as much as 15 percent). The air heated by the exhaust is directed into the engine induction system, which results in a richer fuel/air mixture and additional power loss. A delay of 30 seconds up to several minutes may be expected until normal engine power returns. The circular recommended the use of carburettor heat briefly (particularly with float-type carburettors), immediately before take-off if the relative humidity was above 50 percent and the temperature below 21 degrees C, to remove any ice which may have accumulated during taxi and pre-flight engine checks.

The engine manufacturer recommended that carburettor heat should not be used for take-off as it was not necessary and it may cause detonation and possible engine damage. The aircraft manufacturer recommended a check of the system before take-off, and that carburettor heat be placed ON in the event of an engine failure, other than immediately following take-off. The Operations Manual noted that an unexplained loss in engine speed could be caused by carburettor icing or air intake filter ice and cautioned to watch for signs of icing and apply carburettor heat as required. The section titled `Engine Failure After Take-off (under 700 feet)' did not mention the use of carburettor heat.

One characteristic of the onset of carburettor, or induction system icing, on an engine fitted with a fixed pitch propeller, is the gradual deterioration of the engine RPM. Aircraft engines equipped with constant speed propellers, such as the accident aircraft, compensate for the gradual RPM deterioration by decreasing propeller pitch to maintain a given RPM.

The previous owner of the aircraft reported experiencing an engine power loss (with the 180 horsepower engine fitted), while on approach to land several years earlier. That event was believed to have been due to carburettor icing, as no mechanical anomaly was discovered. The previous owner further reported that the onset of the power loss was immediate, with little time to react.

Optional equipment for the Cessna 172 model aircraft included a Carburettor Ice Detector system. This system utilised an optical probe in the carburettor throat, which is so sensitive that it can detect `frost' up to five minutes before ice begins to form, giving the pilot time to take corrective action. Examination of the carburettor revealed that an optical probe was not fitted.

Aircraft performance

Reports from witnesses indicated that the aircraft took off from the island ALA runway 28 and that take-off performance was apparently acceptable. According to the Aircraft Flight Manual, the maximum permissible crosswind component for take-off and landing was 15 kts. The Operations Manual stated, `Pilots will not take-off or land a Company aircraft when the crosswind component exceeds that specified in the relevant Aircraft Flight Manual.' Using the weather information noted previously, the crosswind component during take-off was calculated by the ATSB to be in excess of 15 kts.

Aircraft fuel

The aircraft fuel selector was found in the BOTH position (both tanks feeding the supply line to the engine) as required for take-off. Damage to the aircraft fuel tanks precluded establishing the exact fuel state of the aircraft at the time of impact. There was a strong smell of fuel in the area of the crash site. A fuel sample was removed from the right wing tank and sent for analysis by a National Association of Testing Authority approved laboratory. The laboratory confirmed that the fuel sample was Avgas 100, which was the correct grade and specification for the engine and no anomalies were noted.

An examination of load sheets used on previous flights to the off-shore islands was completed. These sheets confirmed that the pilot had previously adhered to the Operations Manual policy of maintaining a maximum of 120 L total fuel for flights to off-shore islands of 40 minutes or less, to avoid aircraft structural stress during ground operations. The flight from Smithton to Trefoil Island was approximately 12 minutes. The investigation team generated a flight plan using this fuel information, weather data, witness statements and fuel consumption estimates for the aircraft. The ATSB calculated that the aircraft had approximately 90 L of fuel at the time of take-off on the accident flight.

Aircraft loading

The Operations Manual stated that `Pilots shall prepare a passenger list/manifest and leave it for retention at the aerodrome of departure on all Charter Flights. Pilots will also prepare and leave passenger list/manifest prior to departing the off-shore islands.' The manual specified that the passenger/manifest sheets were to be left `inside the tractor shed' on Trefoil Island. It also stated that if no scales for weighing were available, portable scales were to be carried for use by the company pilots. It further stated that company pilots were to ensure that the aircraft was loaded strictly within the weight and balance limitations.

No passenger/manifest sheet for the accident flight was recovered from the accident site, the island ALA area, or from the operator's Smithton facility. A witness stated that the pilot did not leave the immediate vicinity of the aircraft and did not leave any documents behind prior to the accident flight. No portable weighing scales were recovered from the accident site. When interviewed, passengers from previous flights to and from the off-shore islands reported that they were not weighed, and that the pilot had rarely asked for their body weights. These same passengers reported that the pilot personally loaded all baggage into the aircraft baggage compartment, but did not weigh it. Witnesses who observed the pilot loading the baggage compartment prior to the accident flight also reported that he personally loaded the baggage, but did not weigh it.

The FAA Type Certificate Data Sheet (TCDS) for the aircraft noted the maximum permissible baggage compartment load limitation was 120 pounds (54 kg). Numerous items, such as tools and personal equipment, were located in the immediate area of the wreckage. When weighed these items totalled 108.2 kg. Of that amount, 19.7 kg included items not normally kept in the baggage compartment, but in the main cabin area. That indicated a total baggage compartment load of 88.5 kg at the time of take-off, 34.5 kg in excess of the maximum weight limitation for the aircraft.

The TCDS further noted that the Maximum Take-off Weight (MTOW) for the aircraft was 2,300 pounds (1,043 kg). The MTOW is the maximum allowable weight at the start of the take-off run. The fuel estimated to be on board at the time of take-off was 90 L. ATSB and aircraft manufacturer's calculations indicated an aircraft take-off weight of 1,123 kg (2,475.7 pounds), signifying a take-off weight in excess of the maximum limitation by 79.6 kg. The calculations also indicated that the MTOW would have been exceeded even with all fuel removed.

Advisory material

The Australian Civil Aviation Safety Authority, Civil Aviation Advisory Publication (CAAP) Number 235-1(1), Standard Passenger and Baggage Weights, recommended the following:

`11. Because the probability of overloading a small aircraft is high if standard weights are used, the use of standard weights in aircraft with less than seven seats is inadvisable. Load calculations for these aircraft should be made using actual weights arrived at by weighing all occupants and baggage.'

Aircraft balance and centre of gravity

Aircraft balance refers to the location of the C of G, along the longitudinal and lateral axis. In order to assure predictable aircraft control, the aircraft manufacturer established limitations along the longitudinal axis at fuselage stations measured in inches, in relation to a reference point or datum (located at the forward face of the engine compartment firewall). The C of G limitation for operation of the aircraft in the normal category at maximum Take-off Weight of 2,300 pounds (1,043 kg) was a forward limit of 38.5 inches (977.9 mm) aft of the datum and rearward limit of 47.3 inches (1201 mm) aft of the datum. Aircraft longitudinal C of G was calculated by dividing the total moment of the empty aircraft and on-board items (weight multiplied by the fuselage station) by the total weight of the empty aircraft and items. The location of the C of G at the estimated take-off weight of 1,123 kg (2,475.7 pounds) was approximately 46.3 inches (1176 mm) aft of the datum.

US FAA AC 91-23A Pilot's Weight and Balance Handbook (superseded) stated, `C.G. limits may be expressed graphically in the aircraft weight and balance reports by means of an index envelope. The envelope defines the forward and aft limits and also the maximum weight limit in terms of index units.' The index envelope for the Cessna 172G was referred to as the loaded aircraft moment envelope. Flight with a total moment outside of this envelope was not recommended. When plotted, the calculated loaded aircraft moment at take-off was outside the C of G moment envelope, exceeding the aircraft manufacturer's recommendation.

The aircraft manufacturer advised that with an aircraft loading condition as plotted, if rising terrain and strong winds combined to create significant vertical shear, the risk of loss of control of the aircraft would be increased, even if anticipated by the pilot. The manufacturer further reported that the nose pitch-up and subsequent departure from controlled flight as witnessed were consistent with an aircraft that was flown exceeding the loaded aircraft moment envelope limitation. The estimated aerodynamic stall speed and other aircraft performance figures at the calculated aircraft take-off weight were not available from the manufacturer, as the aircraft was being operated outside the certified moment envelope.