During a daily inspection three days prior to the accident, the pilot noted a fuel leak in the area where the fuel filter bowl fits on top of the fuel pump. A new gasket was made, from a thinner material than the original, and a silicon sealer was therefore applied as an added precaution. The aircraft subsequently operated without fuel problems, until the engine failed completely following a particular takeoff. The pilot attempted to land straight ahead, but the airspeed was less than the optimum and the aircraft touched down heavily, collapsing the right hand gear leg. The investigation determined that the fuel filter outlet was blocked by the silicon sealer that had been used on the gasket. The pilot believed that the silicon that he used was fuel resistant. It is not certain whether the silicon in fact did react with the fuel and degrade or whether an excessive amount was used resulting in the excess breaking away inside the filter bowl and blocking the outlet. The failed landing gear component was inspected and it was determined that it was manufactured from a material as specified in the plans and had failed due to overload.

The aircraft departed Gayndah at 0632 hours on a survey flight. This was the first flight of the day. The survey was terminated near a position about four kilometres north of Mt Goonaneman and the aircraft landed at Gayndah at 1138 hours. The crew assisted the oncoming crew to refuel and prepare the aircraft for the next flight. The aircraft was seen departing Gayndah shortly after 1200 hours. At about 1235 hours witnesses saw a column of smoke in the vicinity of Mt Goonaneman and notified Biggenden Police. The accident site was located by the police about 40 minutes later. The survey flights were being conducted at 80 metres above ground level and traversed the survey area on a flight path which was orientated in a north-easterly/south-westerly direction. The tracks being flown were about 40 kilometres long and 200 metres apart. The accident location indicates that the aircraft was on the first survey line adjacent to the position at which the earlier flight had finished, and was travelling in a north- easterly direction towards the mountain when the accident occurred. Witness observations were not conclusive in establishing the final flight path, however, some of the hearing reports were consistent in describing an unusual noise from the aircraft immediately before the accident. The engine noise was heard by one witness to cease and then recommence twice in quick succession. This could have been caused by a serious engine malfunction or changes made to engine controls by the pilot. Another witness reported seeing the aircraft make what he described as a steep climb and a sharp "U" turn near the mountain. This was not a normal manoeuvre for the aircraft to make and this witness may have observed the aircraft during the accident sequence. Unfortunately the time of this observation could not be established in relation to the time of the accident. The examination of the wreckage was hampered by the extensive destruction resulting from the post impact fire. Dust that was found in the fuel injectors from the left engine could have caused an engine malfunction but this could not be positively established. The right engine could not be fully examined due to fire damage. The propellers from both engines indicated that substantial power was being delivered at the time of impact. The nature of the ground impact was indicative of very low forward speed, high vertical speed, in a steep nose down, right wing low attitude. The pilot was very experienced, and it is considered to be most unlikely that he would allow such a situation to develop, or that he would attempt any unusual manoeuvre at low altitude near the mountain. There is no evidence to suggest that aircraft may have failed structurally, or that the pilot would not have been able to cope with any situation, other than that which had the potential to cause a sudden loss of consciousness. Pathology disclosed evidence of a heart condition in the pilot which could have caused sudden death or incapacity. The pilot's last medical examination, conducted in February 1988, had included an ECG. Had the pilot suffered incapacitation, the crewman, seated in the cabin behind the pilot would have had insufficient time, due to the low operating altitude, to correct the situation.

A witness, who lives alongside the airstrip, observed the aircraft flying in a north-south direction, about 300 feet above ground level and about 300 metres to the east of his position. The witness took particular notice of the actions of the aircraft and watched as it porpoised three times before diving steeply into the ground. Inspection of the wreckage revealed no defect which may have contributed to the accident. Weather at the time was fine, wind was light and variable, and the aircraft was being flown over flat pastureland. Investigation revealed that the pilot suffered from epilepsy, had heart disease, and that this was his first flight for over one year.

Circumstances:

The pilot reported that he joined downwind for runway 34 and completed downwind checks. He stated that normal landing gear extension occurred when he selected the undercarriage down, and a normal gear down indication was obtained (three green lights). The pilot stated that on late final approach the gear warning horn sounded when he reduced the throttle settings, and he then noticed there were no green lights. He said that he immediately went around and established the aircraft in the circuit at 1000 feet. The pilot stated that he extended and retracted the landing gear several times, and on each occasion the gear operated normally, and normal indications were obtained. A second approach was then commenced, and the pilot said that the lights were monitored all the way to touchdown. He reported that the landing was normal, but that shortly after touchdown the right wing began to sink. The pilot stated that he believed the right hand gear had failed, and that he attempted to recover the situation by applying full power and using aileron to level the wings with a view to going around, and proceeding to an aerodrome better equipped to handle an emergency landing. At this point, the left hand gear and the nose gear collapsed, and the aircraft slid to a stop on the gravel runway. Marks on the runway showed that the aircraft had landed with the landing gear extended and that the gear retracted soon after touchdown. Subsequent examination of the aircraft found no evidence of any pre-existing defect which may have contributed to the development of the accident. The landing gear switch was found in the down position, and the gear actuating mechanism was found to be in an intermediate position. Damage to the mechanism appeared to be consistent with the gear being driven down with the weight of the aircraft on it. This could occur if the gear was inadvertently selected up during the landing and then down again. The pilot is adamant, however, that this did not occur. The cause of the landing gear retracting during the landing roll could not be determined.

Significant Factors:

The following factors were considered relevant to the development of the accident:

The landing gear retracted during the landing roll. The cause of the unscheduled gear retraction could not be positively identified.

The pilot's custom was to connect an earthing lead to the aircraft whilst it was parked on the pad. The earthing lead consisted of about five metres of flexible wire with a jack type connector fitted to one end and an alligator type clip on the other. The jack mated with a socket fitted in the aircraft just behind the rear cabin door about half way up the fuselage on the left hand side. The pilot's preflight inspection was interrupted by the arrival of his passengers. After seeing them aboard and stowing their gear he failed to disconnect the earthing lead entered the helicopter started up and lifted off. As the tension came on the earthing cable it could not pull the jack out of the socket because of the sideways load on the socket and so pulled the alligator clip off the ground earthing point. The cable and clip trailed back along the side of the aircraft as it went forwards through translation and in so doing the clip and cable entered the tail rotor arc and were severed. Only minor damage in the form of a gouge in the end of one tail rotor blade occurred. The pilot heard the noise of the cable being cut but thought it to be the passengers moving gear around in the back of the helicopter. He continued with the flight and noticed the tail rotor damage at the next stop (Garden Island 35 mins) saw the severed cable still hanging from the fuselage jack and then flew the aircraft to Jandakot (another 8 mins) for inspection. At Jandakot the tail rotor gearbox was found to be making a large quantity of metal and the gearbox and rotor system were removed for overhaul.

Circumstances:

While aerial spraying rice crops the pilot detected a change in the engine note accompanied by a vibration. He thought it sounded like a failure of one cylinder and decided to return to the airstrip from which he was conducting the operation. Some 15 seconds later the engine lost power and it was obvious that the aircraft would be unable to reach the airstrip. The pilot then dumped the load and continued straight ahead for a landing into a cultivated paddock. Ten seconds later the engine failed completely. With the aircraft now being unable to reach the selected paddock the pilot made a heavy landing just short of it in an attempt to bounce the aircraft over an irrigation channel into the adjoining paddock. The aircraft struck the far bank of the channel resulting in substantial damage. The engine had completed approximately 170 hours since overhaul and examination revealed that it had suffered a broken crankshaft. The reason for the failure was not positively determined but overseas reports indicate that other low time engines which are fitted to the same type of aircraft have experienced similar failures. This type of aircraft is fitted with a larger diameter propeller than others using the same engine. The larger propeller may result in a harmonic vibration being set up at a particular engine manifold pressure and rpm setting which ultimately leads to crankshaft failure. An experienced USA engine overhaul company suggests that to prevent this vibration from developing the operating engine speed for Ag Cat aircraft should be increased from 2000 rpm to 2100 rpm. This suggestion has been bought to the attention of the Civil Aviation Authority. This accident was not the subject of an on-scene investigation.

Significant Factors:

The following factors were considered relevant to the development of the accident:

1. The engine failed due to a broken crankshaft.

2. The reason for the crankshaft failure could not be positively determined.

Circumstances:

The aircraft returned to the Bankstown circuit after a period of dual instruction in the training area. The instructor, who was flying the aircraft, landed on runway 11 centre. After touchdown, the right main gear leg separated from the wing; the aircraft swerved to the right and came to rest on the grassed area between runways centre and right. The investigation revealed the front attachment bolt of the right main gear leg had failed due to long term fatigue cracking. Corrosion of all attachment bolts was found which indicated moisture had entered the bolt holes. Degrading of bolt pre-load occurred as a result of corrosion beneath the bolt heads. The remaining two attachment bolts held the leg in place for approximately nine take-off and landing cycles before the final failure which led to the gear leg separating from the wing mounting. The organisation responsible for maintenance of the aircraft indicated periodic inspections were conducted in accordance with Civil Aviation Order 100.5.1. This order requires a visual inspection of the main landing gear components for condition and security. However, the aircraft manufacturer specifies in its periodic inspection schedule that main landing gear bolts are to be inspected for condition, torque and security, every 100 hours. Existing Civil Aviation Orders do not require maintenance organisations to comply with aircraft manufacturer's inspection schedules. Had the maintenance organisation complied with the manufacturer's inspection schedule, the failure of the front attachment bolt might have been prevented. This accident was not the subject of an on-scene investigation.

Significant Factors:

The following factors were considered relevant to the development of the accident:

1. The front attachment bolt failed due to long term fatigue cracking.

2. The aircraft was not inspected in accordance with the aircraft manufacturer's period inspection schedule.

3. Civil Aviation Authority Orders do not require maintenance organizations to comply with aircraft manufacturer's schedules.

Recommendations:

It is recommended that the Civil Aviation Authority consider amending the relevant Orders to ensure aircraft maintenance organisations comply with the recommended schedule of maintenance laid down by the aircraft manufacturer.

The aircraft was a single-seat, single-engined, ultralight monoplane. The mid-wing was supported by an over wing brace strut. It was being flown by an experienced ultralight pilot after having been purchased secondhand by another person. Following a normal takeoff the aircraft was observed to carry out a manoeuvre consistent with a deliberate stall and recovery. It was then seen to enter a level turn to the left, in which the angle of bank gradually increased to about 90 degrees, despite the aircraft being limited to an bank angle of 60 degrees. The turn continued through at least 360 degrees, during which the aircraft began to lose height. A witness said he then saw the aircraft roll out of the turn, still losing height, before it suddenly appeared to lose control and entered a steep spiral dive. The aircraft was seen to impact the ground heavily. An examination revealed that the welding of the attachment bracket, for the wing brace struts, had failed. This had allowed both wings to fold upwards and twist rearwards, resulting in an immediate loss of control. A laboratory evaluation of the failed weld indicated an extensive lack of fusion between the weld deposit and the bracket.

The pilot had failed to return from a flight in the local area. A search was organised, and the wreckage of the aircraft was located the following morning. It had crashed into a swampy area and the pilot, although still alive, was seriously injured and unconscious. Although he subsequently survived his injuries he has no recall whatever of the accident sequence. Inspection of the wreckage indicated that the aircraft had hit the swamp in a flat attitude at a high vertical velocity but with little forward speed. The propeller was undamaged and appeared to have been stopped at the time of impact. A subsequent check of the engine revealed a faulty magneto. The precise nature of the fault was not determined but the engine would not start or run until an alternative ignition system was fitted. The engine then ran perfectly and no other faults were found with it. This particular engine was fitted with a single ignition system, as permitted by existing regulations.

Circumstances:

The aircraft was observed to depart from a farm paddock, climb to approximately 300 feet above ground level and accelerate to cruising speed. Shortly afterwards the engine began to labour, there was a clattering noise a loud metallic bang and then silence. The aircraft was observed to pitch upside down and descend vertically until it collided with the ground. An inspection of the wreckage indicated that the labouring noise was caused when the hydraulic pump actuating cable became entangled in the propeller. The clattering noise was caused by the propeller blades striking the main rotor blades and the loud metallic bang was caused when the main rotor blades struck the rudder assembly. The hydraulic pump was part of a main rotor drive system used to increase main rotor rotational speed prior to take off. All noise ceased when the propeller made solid contact with the main rotor blades and the engine stalled. The main rotor strike on the rudder assembly caused the fibreglass rudder to break up in-flight and pieces of the rudder were found a significant distance downwind from the wreckage. A study of rotorcraft aerodynamics indicates that a sudden application of forward control input is required to create the circumstances which will significantly reverse the airflow through the rotor and cause the main rotor blade to make contact with the propeller and the rudder assembly. The pilot had completed a rotorcraft pilot's course and taken delivery of his aircraft the previous month. Shortly after the aircraft was delivered the pilot made a number of modifications to the hydraulic, fuel and the electrical systems. The pilot did not check any of his modifications with the manufacturer. The pilot had a contract which required the use of the rotorcraft, and he was in a hurry to complete preparations for the contract. During the modifications, the pilot re-routed an actuating cable running between a hand operated lever, mounted on the control stick, and a hydraulic pump mounted at the rear of the engine and immediately adjacent to the propeller hub. The manufacturer had routed and secured the cable so that there was no unnecessary slack. The re-routing increased the amount of slack in the cable by more than 200 mm, and this was sufficient under the right circumstances, to allow the cable to become entangled with the propeller. Marks and damage to the aircraft and its components indicate that when the aircraft reached its cruising speed, the slack cable at the rear of the aircraft was blown backwards until it was caught by a bolt shank which was part of the propeller hub. The cable became entangled in the propeller which in turn pulled the control stick back towards the pilot. The pilot's natural reaction would have been to apply force to the control stick to push it forward against the rearward force applied by the cable. As the propeller continued to turn with the cable attached, all slack was taken up and the cable failed at both the forward and rear attachment points. The rearwards force on the control stick would have ceased suddenly and the pilot's countering force would have pushed the control stick forward setting up the circumstances necessary to cause the main rotor blade to strike the propeller and rudder. The entire sequence would have taken no more than 2 to 3 seconds.

Significant Factors:

The following factors were considered relevant to the development of the accident:

1. The pilot made a number of modifications to his aircraft without checking whether or not they affected the safety of the aircraft. This probably occurred because the pilot was anxious to complete preparations for an impending contract.

2. The pilot failed to appreciate the significance of leaving additional slack in the actuating cable when he re-routed it.

3. The hydraulic actuating cable became entangled in the propeller and the ensuing unexpected control forces caused to pilot to lose control of the gyrocopter.

Recommendations:

1. It is recommended that the Sports Rotorcraft Association incorporate in their design standards, the requirement that only controls necessary for safe in-flight operations be attached to the control stick.