The crew of the Boeing 737 reported that when the speed brake was selected, during descent into Sydney with the autopilot engaged, the aircraft rolled slightly to the right. The autopilot was disengaged, and the speed brake was again selected with the same result. The speed brake was restowed and the flight continued and landed without further incident.

The operator reported that inspection of the aircraft, on 15 February 2001, revealed that the left-wing number three flight spoiler "UP" cable (P/No. WSA2-3) had failed at a pulley in the left wheel well at Wing Buttock Line (WBL) 73.00. The failure was due to corrosion as evidenced by rust deposits at the failure location. During rectification, all other left wing spoiler cables were replaced due to evidence of minor corrosion. Following repair, the aircraft was returned to service.

The operator reported that, after a previous spoiler cable failure in 1997 due to corrosion, an Engineering Release (ER) had been issued to require the inspection of all spoiler cables at the next Phase 20 check and subsequent 2C check with cable replacement at the next 4C check. Replacement at the 4C check terminated the inspection requirements of the ER.

As a result of the cable failure on 15 February 2001, the ER was revised to require inspection of the cables on an ongoing basis with cable replacement at every 4C check interval to preclude recurrence.

Subsequently, on 28 February 2001, the incident aircraft underwent overnight maintenance at Melbourne. During the maintenance inspection, the left-wing spoiler cables, that had previously been changed at Sydney on 15 February 2001, were found to be mis-routed. The operator's investigation revealed that the maintenance engineers involved in the original rectification had travelled from Brisbane to Sydney that day and had worked a period in excess of 24 hours with minimal breaks. Excessive hours worked and fatigue of the maintenance engineers was considered to have contributed to the misrouting of the cables and the failure to detect the misrouting during a duplicate inspection of the spoiler control system.

A Boeing 747 (B747) had departed Avalon, Victoria on a track that passed over Melbourne and then to the northeast. The crew had been issued with a requirement to initially maintain flight level (FL)200. The departures controller had imposed the limitation in accordance with standard operating procedures that required the "cap" to be placed on all aircraft that had planned to a higher flight level. Airspace above FL200 was under the jurisdiction of a sector controller. In addition, the horizontal boundary between the two controllers' airspace below FL200, was 30NM from Melbourne airport.

The departures controller had a Piper Navajo aircraft tracking ahead of a sequence of three jet aircraft departing from Melbourne airport. That situation required that the jet aircraft be radar vectored around the slower aircraft before they could be placed on their flight planned tracks. The first of the jet aircraft was a Boeing 737 (B737) for Brisbane, which departed approximately the same time as the B747.

The tracks of the aircraft were such that the B737 was initially to the left of the B747, but at approximately 30NM northeast of Melbourne they crossed and thereafter diverged.

The sector controller had noticed that the B747 would probably reach FL200 while still in departures airspace and, as a consequence, be forced to maintain FL200. In order to provide the crew of the B747 with an unrestricted climb profile, he coordinated with the departures controller to authorise the crew of the B747 to climb to FL370.

At that time, the B737 was below and approximately 30NM ahead of the B747. However, the ground speed of the B747 was approximately 70 knots faster than the B737, and the B737 had a greater rate of climb.

As the B737 approached the horizontal airspace boundary, the departures contoller handed over the aircraft to the sector controller while it was passing FL170. The sector controller then approved the crew to climb to FL370, as he believed there was sufficient distance between the aircraft to maintain separation. A short time later, the B747 was also handed over to the sector controller as it was passing FL200 and approximately 8 NM behind the B737. At that moment, The Australian Advanced Air Traffic Control System (TAAATS) Short Term Conflict Alert (STCA) activated and the controllers immediately attempted to prevent an infringement of separation standards. However, the crew of the B747 was on his frequency and asked the departures controller to maintain that aircraft at a lower level. The departures controller issued an instruction for the crew of the B737 to maintain FL190 but did not receive a reply, because that crew was on the sector frequency as instructed.

At that moment, the crew of the B747 made radio contact with the sector controller who immediately issued an instruction for them to turn the aircraft. There was a delay as the crew questioned the instruction, but they commenced the manoeuvre when the controller issued the instruction a second time using the word "immediately". He then instructed the crew of the B737 to level out, which they did.

The aircraft passed approximately 2.5NM apart while the vertical separation standard of 1,000ft did not exist. The required radar separation standard was 5 NM.

Shortly after take-off, the crew of the Boeing 747 aircraft received a number four engine fire warning advisory. The crew actioned the non-normal checklist, secured the number four engine, and discharged two engine fire bottles before the fire indication extinguished. The crew then jettisoned fuel and completed an uneventful landing at the departure airport.

Initial inspection by operator maintenance personnel revealed evidence of charring and discolouration around the engine combustion area. Maintenance personnel removed the engine from the aircraft and the engine was then test run on an engine test cell. The test run revealed a fuel leak from the left and right fuel manifold connectors. The connectors were confirmed as being secured with safety lockwire, but subsequent torque checks revealed a lower value of torque than that required by the engine manufacturer's maintenance manual. The fuel manifold connectors were re-torqued and another test run completed. The resulting test run at high power settings revealed no discrepancy. A decision was made by the operator to have the engine overhauled before being put back into service.

During cruise, the Boeing 767 (B767) crew noticed that the fuel consumption was 230 kg per hour more than normal. After the cargo had been unloaded at the destination, the crew had the cargo re-weighed. The actual cargo weight was more than 3,400 kg greater than the weight stated on the manifest.

The operator reported that its investigation revealed that four pallets had been carried from Honolulu to Sydney and that all were heavier than the weights recorded on the manifest. The discrepancies were 2,277 kg, 887 kg, 220 kg, and 10 kg.

The two pallets with the greatest weight discrepancies had been transported from mainland USA in a freight aircraft operated by a US operator, and had been transferred at Honolulu to the B767 for onforwarding to Sydney. The weights stated on the pallet tags for the internal US leg were in pounds, but the B767 operator required the weights to be in kilograms. The agent who handled the freight for the US operator at Honolulu converted the weights from pounds to kilograms and pencilled in the converted weights on the deadload sheet for the Honolulu-to-Sydney flight.

When the agent who handled freight at Honolulu for the B767 operator received the pallet weights, she did not check the figures against the loadsheet issued by Load Control. Consequently, she did not realise that the weights stated on the loadsheet had already been converted to kilograms, and applied the conversion a second time. Also, as the agent for the US operator was confident that she had passed the correct weights to the B767 agent, she did not recheck to ensure that the B767 agent had received the correct weight information.

The 220 kg and 10 kg weight discrepancies affecting the other two pallets were probably the result of weighing or recording errors.

The B767 operator has since amended procedures for its agent at Honolulu. Under the new procedures, only the supervisor will carry out weight unit conversions and liaise with other operators' agents. This should minimise the risk of erroneous information being passed between agents and help to ensure that all required checks are carried out.

The B767 operator is monitoring the effectiveness of the new procedures.

A Boeing 717 aircraft was in a left turn holding pattern, descending through flight level 230, when the right engine shut down. The flight crew actioned the emergency procedures and attempted, unsuccessfully, to restart the engine. They notified air traffic control of the problem, then requested and received a vectored straight-in approach and landing.

Following the event, the operator's maintenance personnel conducted troubleshooting of the right engine. Several fault codes were noted in the computer memory, which related to Channel A of the electronic engine controller (EEC). A maintenance records check found that the right engine fuel metering unit (FMU) had been replaced approximately 50 flight hours prior to the event. At the time of the event, there were no maintenance manual requirements for an EEC stored faults check following an engine run after replacement of the FMU. Maintenance personnel noted, then cleared, the fault codes from the computer memory and the engine was successfully test run. They then chose to remove both the right engine FMU and the EEC for further testing. The EEC unit was sent to the engine manufacturer for bench testing and operating on a test bed engine.

Component testing

The FMU manufacturer's testing found no faults in the unit. Initial testing of the EEC by the engine manufacturer could not duplicate, on the test bed engine, the dual channel failure and subsequent shutdown. Analysis of the fault codes recorded by the operator's technicians following the event confirmed that several FMU electrical related fault codes were pre-existing on Channel A of the EEC at the time of the occurrence. When the engine manufacturer repeated the testing with the recorded fault codes entered into Channel A of the EEC, and simulated loss of Channel B, they successfully repeated the dual channel failure and resulting engine shutdown.

Electronic engine controller

The electronic engine controller was a two-channel (Channels A and B) electronic unit with system redundancy. It controlled, among other items, engine start sequencing, power requirements, operating temperature, turbine speeds, fuel flow, engine monitoring, and automatic relight. It contained fault detection, storage, and readout capabilities, all stored on an electrically erasable/programmable read-only memory (EEPROM) located on a computer board assembly. The EEPROM provided a history for troubleshooting purposes of any fault event within the EEC or associated control systems by logging a fault code of the event. Those fault codes were then stored until intentionally cleared during maintenance action. The distinct two-channels in the unit ensured that should one channel fail, the other would assume control and monitoring of the engine. Testing by the engine manufacturer revealed that repeated random access memory (RAM) parity errors in Channel B of the EEC resulted in repeated multiple resets of the channel. Those repeated resets had proven to result in a loss of Channel B functionality.

Engine rotation during restart attempts

Following the event, the flight crew stated that they could not obtain a windmilling engine N2 (gas generator RPM) value of 14% as required in the emergency procedures for restart of the engine. They stated that the maximum engine N2 witnessed was 8%.

The engine manufacturer recommended a 14% N2 value (approximately 2,380-RPM) at fuel introduction during engine starting procedures. That allowed a cooler start and prevented engine deterioration. The value of 8% N2 for the ALL ENGINE FLAMEOUT emergency windmilling procedure was based on the minimum engine-driven fuel pump pressure to open the engine pressurising valve. The airframe manufacturer reported that windmilling flight tests were successfully demonstrated at airspeeds as low as 240 knots with N2 windmilling rotor speeds as low as 8%.

The airframe manufacturer estimated that at a stable condition of 10,000 ft altitude, and 250 knots indicated airspeed, the occurrence engine should have exceeded 10% N2 before engine start switch engagement. Their review of the digital flight data recorder (DFDR) revealed no evidence of N2 increase as would be seen with engine starter engagement. The airframe manufacturer reported that their understanding was that the anomalies experienced during the event would have prevented the starter air valve opening during the restart attempts.

Flight data recorder

Examination of the recording indicated that the aircraft arrived at the destination and then departed on another flight to the north. The reported ground runs had not been recorded, as required by Civil Aviation Order (CAO) 20.18 Section 6 paragraph 6.6.

Australian CAO 20.18 Section 6 paragraph 6.6 stated, "The operator of an aircraft which is required by this section to be equipped with recorders shall take action to ensure that during ground maintenance periods the recorders are not activated unless the maintenance is associated with the flight data recording equipment or with the aircraft engines."

Australian CAO 20.18 Section 6 paragraph 6.3 stated, "Where an aircraft is required to be so equipped by this section, the flight data recorder system shall be operated continuously from the moment when the aircraft commences to taxi under its own power for the purpose of flight until the conclusion of taxiing after landing."

The intent of the Australian legislation was that when an aircraft commenced taxiing under its own power for the purpose of flight, the flight data recorder (FDR) would record until the aircraft was parked at the conclusion of the flight. That action ensured a continuous record of aircraft operation was maintained for the duration of the flight.

Subsequent Enquiries made to Boeing Long Beach Division, the aircraft manufacturer, revealed that when the aircraft park brake was set, the FDR would cease recording. Boeing Long Beach Division stated that the FDR would begin recording by two methods. The first "normal" recording mode activated when either fuel shutoff switch was set to run and the park brake was released. The second "maintenance" recording mode was activated by accessing a FDR RUN command via the Multifunction Control Display Unit.

Aircraft Australian certification

The Boeing 717-200 was issued a Type Acceptance Certificate in accordance with Civil Aviation Regulation, (CAR) 21.29A which allowed Type Certificate acceptance for imported aircraft certified by the National Airworthiness Authority (NAA) of a recognised country, in this case the United States of America. The Boeing 717-200 aircraft may not comply with the appropriate Australian Civil Aviation Regulations and associated Orders.

The Sikorsky S76C helicopter with two crew and ten passengers on-board was in a hover with the flight crew completing before takeoff checklist items. The pilot reported that while trimming the engines, a "pop" was heard. He then noted that the number-one engine gas generator outlet temperature (T4) was in excess of 1000 degrees C. The helicopter was then landed uneventfully. The crew reported that the only cockpit indication of imminent failure was the almost simultaneous illumination of the number-one engine chip (magnetic particle) detector advisory. There was no reported engine fire.

Examination of the helicopter revealed minor shrapnel damage to the engine exhaust extension and engine cowling. The engine was removed from the helicopter and sent to the engine manufacturer for technical disassembly inspection.

Engine manufacturer examination

The Arriel engine was comprised of five modules. The gas generator or high-pressure section was known as Module Three (MO3). The manufacturer's final report noted a separation of turbine blade number six of the gas generator second stage disc. The blade had separated above the `fir tree' attachment point but below the blade platform, and had punctured the second stage nozzle guide vane turbine ring. One adjacent blade (number seven) in the direction of turbine wheel rotation was also noted as cracked.

Metallurgical examination by the manufacturer attributed the blade failure to a low-cycle fatigue cracking mechanism. Dimensional inspections failed to reveal any sign of non-conformity that could have led to the development of the abnormal loads. However, the manufacturer stated that platform/disc interferences were a potential source. The manufacturer concluded that abnormal loading was the major contributing factor in the failure, given the reported absence of anomalous material features or evidence of high-temperature operation.

Engine background

At the time of the occurrence, the Turbomeca Arriel engine model 1S1, serial number 15522, had accumulated 4,737.4 hours in service and 4,471 cycles since new. It had accumulated 1,740.0 hours in service and 1,615 cycles since overhaul. Following overhaul, the engine had been installed on March 11, 1999.

Turbine disc life limit

Part number 0292253250 turbine disc was life limited to 10,000 cycles. The incident turbine disc had accumulated 1,615 cycles since new.

The pilot of a Bell Long Ranger 206L-1 was returning to base following an agricultural crop-spraying task. While transiting a ridgeline of the Connors Mountain Range, the helicopter collided with wires and impacted the ground in a densely wooded area about 200 metres beyond the wires.

The pilot received fatal injuries, and the helicopter was destroyed in a post-crash fire. Witnesses had observed a helicopter approaching the ridgeline at a very low height, and reported that, shortly afterwards, a pall of black smoke was visible.

The helicopter had struck two three-strand lightweight high-tensile steel wires of a powerline supplying a repeater site. A wire strike protection system (WSPS) had not been fitted to the helicopter. The wires were aligned on 060 degrees magnetic, with a maximum height of 31.5 metres for the upper wire and 30.1 metres for the lower wire. The position of the wires was not annotated on the relevant Visual Terminal Charts and they did not have high visibility devices attached. Company employees said that it was usual for the pilot to fly at a low height when transiting to and from the work location.

Examination of the wreckage indicated that the helicopter had struck the ground with a vertical downward force and at a low forward speed on a heading of 030 degrees magnetic. The main rotor blades were severely fractured by contact with the surrounding trees. The tail boom was severed just forward of the horizontal stabiliser and the tail rotor gearbox and blades were intact. Fire destroyed the helicopter cabin and heavily damaged the engine compartment and upper transmission deck. Examination of the engine to determine pre-impact airworthiness was limited due to fire damage.

The white colour coded main rotor pitch control rod, broken sections of the windscreen, and the top section of the main transmission cowling were found along the flight path about 75 metres before the main wreckage. One main rotor blade severed the upper wire. The lower wire contacted the fuselage in the area of the forward canopy, progressed up to the fibreglass transmission cowl, and separated the top lip of the cowl. That wire, together with the separated section of cowl, then contacted the flight controls above the main rotor swashplate, causing static overload and separation of the white colour-coded main rotor pitch change rod. Directional control of the helicopter was lost following the separation of the control rod.

The investigation found no evidence to suggest that the helicopter was not fully serviceable at the time of the accident. It is likely that the oblique angle of approach to the wires limited the pilot's ability to detect them, and "contour flying" offered minimal reaction time for the pilot to avoid the wires had they been detected.

A review of the Australian Transport Safety Bureau (ATSB) database for the period January 1, 1995, to March 3, 2001, revealed six fatal agricultural helicopter accidents in Australia. These accidents represented seven fatalities and 42% of all fatal helicopter wire-strike accidents during the period.

Previous safety action

The Australian Transport Safety Bureau (then known as the Bureau of Air Safety Investigation) made a recommendation in 1995 to the Civil Aviation Safety Authority (then known as the Civil Aviation Authority) following a similar fatal accident related to a wire strike of a helicopter. The recommendation, R19950120, stated:

The Bureau of Air Safety Investigation (BASI) recommends that the Civil Aviation Authority (CAA):

(1) Require the fitment of approved WSPS kits for all helicopters engaged in low flying activities for which a kit exists; and

(2) That only agricultural spray kits compatible with WSPS be approved for fitment to these helicopters.

The CAA response to the recommendation was as follows:

While WSPS may have been of benefit in this and similar accidents, the Authority believes that the fitment of WSPS should not be mandatory. However, the CAA is of the view that it should be strongly encouraged when suitable equipment is available.

The CAA in conjunction with BASI, is prepared to undertake an industry education program highlighting the hazards associated with low level helicopter operations as well as the advantages provided by the fitment of WSPS to appropriate helicopters.

On 15 May 1996, BASI classified the recommendation CLOSED- No further action, however, continued to monitor the number of wire strike accidents.