Collision with terrain

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

This Safety Notice also stated:

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

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

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

CONTROL OF AIRCRAFT

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

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

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

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

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

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

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

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

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

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

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.