At approximately 1211 Eastern Standard Time on 2 April 2004, the pilot of a Bell Helicopter Company 47G Soloy helicopter, registered VH-UTY, was conducting fire-ant baiting operations at Nudgee, about 5 km north-west of Brisbane Airport. The operator's chief pilot occupied the right control position and was supervising the pilot.

Near the end of a baiting run, the chief pilot told the pilot that he wanted to demonstrate a procedural turn and asked the pilot to follow him through the manoeuvre. Both pilots reported that, during the turn, the helicopter began to yaw right. The chief pilot then said that he was taking control of the helicopter. He reduced engine power but was unable to arrest the right yaw. The helicopter continued to descend towards a canal and struck the water slightly nose down and banked to the right. Both occupants were injured in the impact but were able to exit from the helicopter unaided.

The pilot reported that he had completed two previous baiting operations in the helicopter during that day without incident.

A subsequent examination of the helicopter found that the tail rotor control pedals installed at the right control position operated in the reverse sense, compared with the tail rotor control pedals installed at the left control position. That meant that tail rotor control pedal inputs made by the chief pilot would have produced a yaw response opposite to that which would normally be expected.

The helicopter operator reported that the tail rotor control pedals for the right control position had been refitted to the helicopter before the accident flight.

In 1954, the Bell Aircraft Corporation, as it was then known, issued Service Bulletin (SB) 98. The SB required installation of a stop assembly (part number 47-722-165-1), under both control position footrests. The purpose of the stop assembly was to prevent the incorrect re-installation of the tail rotor control pedals. UTY was manufactured in 1966, and the stop assembly would have been incorporated as a standard build item during manufacture.

In October 1971, the then Australian Department of Civil Aviation issued Airworthiness Directive (AD) AD/Bell47/69 titled Tail Rotor Control Pedal Assembly Interference Bracket. That AD, which mandated the installation of the interference (stop) brackets to all Bell 47G series helicopters as introduced by Bell SB 98, was still current at the time of the accident.

Examination of the helicopter showed that only part of the tail rotor control pedal assembly bracket as specified in AD/Bell 47/69, remained fitted in the helicopter. The majority of the bracket had previously been removed. There was no evidence to indicate that the removal was as a result of wear or damage sustained in the accident. The maintenance organisation that certified for the last scheduled maintenance check advised that the bracket was in place, and that the co-pilot tail rotor control pedals were not fitted at that time.

The helicopter's maintenance documentation contained no record of the installation of the right tail rotor control pedals, or of the required independent inspection of the flight controls after the installation of the tail rotor control pedals.

The helicopter examination also found that the forward section of the tail rotor drive output shaft, from the main gearbox to just forward of the first bearing hanger assembly, had separated. The separated section was not found. Examination of the remaining broken section of the drive shaft indicated that it had separated due to overload forces that occurred during the accident impact sequence. There was no evidence found of any pre-existing fault in the shaft.

At approximately 1211 Eastern Standard Time on 2 April 2004, the pilot of a Bell Helicopter Company 47G Soloy helicopter, registered VH-UTY, was conducting fire-ant baiting operations at Nudgee, about 5 km north-west of Brisbane Airport. The operator’s chief pilot occupied the right control position and was supervising the pilot.

On 22 March 2004, at about 0435 Coordinated Universal Time (UTC), a Boeing Company 767-338ER, registered VH-OGB, and a Boeing Company 767-319, registered ZK-NCF, were involved in a serious incident in the Ujung Pandang Flight Information Region (FIR).

OGB was southbound and NCF was northbound on air route B473, between waypoint OLENG and waypoint OPABA. The crew of OGB advised Ujung Control that they were at OLENG at 0423 and were level at flight level (FL) 350. The crew later reported that they requested approval from Ujung Control to climb their aircraft to FL370. Ujung Control subsequently approved that request. During that climb, and passing approximately FL355, the crew of OGB received a traffic advisory on their aircraft's traffic alert and collision avoidance system (TCAS), followed by a resolution advisory, about NCF. The resolution advisory instructed the crew of OGB to descend their aircraft. The crew complied with the resolution advisory and descended their aircraft to FL350. The crew later reported that, during the descent manoeuvre, they saw NCF pass 600 ft above their aircraft.

NCF was northbound on air route B473, between waypoint OLENG and waypoint OPABA and was level at FL360. The crew later reported that they received a TCAS resolution advisory on OGB, travelling in the opposite direction. The crew of NCF climbed their aircraft in response to the resolution advisory. The crew later reported that they observed OGB on their TCAS, pass 400 ft below them.

The Australian Transport Safety Bureau (ATSB) was advised of the serious incident and commenced an investigation. The crew of OGB was interviewed and data from that aircraft's quick access recorder was analysed. As the incident occurred within Indonesian territory, the Indonesian National Transportation Safety Committee (NTSC) had the responsibility to conduct an investigation in accordance with Annex 13 to the Convention on International Civil Aviation.

On 26 March 2004, the NTSC informed the ATSB that they had commenced an investigation into the incident and the ATSB appointed an accredited representative to that investigation. The NTSC, being the investigation agency of the country in which the incident occurred, will be preparing the report and has control over the public release of any investigation findings.

As a result of another occurrence within an Indonesian FIR, which involved aircraft from the same Australian operator, that operator implemented new procedures to be followed when its aircraft are operating in Indonesian FIRs (see ATSB occurrence report number 200402411).

The Bureau will publish the NTSC report, when released by the NTSC, on the ATSB website.

The pilot in command of the Fairchild Industries SA-227 aircraft, registered VH-HPE, operating a scheduled Regular Public Transport flight, reported that excessive forward control column force had been required 'to trim the aircraft nose down' during departure from Sydney Airport.

The Australian Transport Safety Bureau did not conduct an on-scene investigation of this occurrence.

On Monday, 22 March 2004, the pilot in command of the Fairchild Industries SA-227 aircraft, registered VH-HPE, operating a scheduled Regular Public Transport flight, reported that excessive forward control column force had been required 'to trim the aircraft nose down' during departure from Sydney Airport. The pitch trim selector was switched to the copilot position, control was passed to the copilot, who was then able to trim the aircraft, and the flight continued to Taree, NSW.  After landing, an examination by the crew revealed that the pilot in command's (left side) control yoke pitch trim switch was operating in the reverse sense from normal operation.

Discussions were held between the flight crew and the operator's chief pilot and chief engineer and a decision was made to continue with the following two scheduled flights before the aircraft returned to a suitable maintenance facility for rectification.  A subsequent engineering examination revealed that the pilot in command's pitch trim switch had been installed upside-down and had to be removed and re-installed in the correct orientation (refer figure 1).  The pitch trim system was checked for correct operation and the aircraft was returned to service.

Figure 1:  Left side pitch trim control switch

In the days preceding the occurrence, the aircraft underwent scheduled maintenance at a contractor maintenance facility.  During maintenance, there was a requirement to replace the left side control column pivot bearings.  To access the bearings, it was necessary to remove the control yoke and the control yoke pitch trim switch by de-soldering the switch wiring and removing the switch from the yoke housing.  After the control column bearings were replaced, the control yoke was re-installed and the trim switch wiring was re-soldered to the respective terminals.  The trim switch was then re-installed into the control yoke and the engineers reported that they conducted a full installation and duplicate functional check of the pitch trim system and completed the documentation in the aircraft maintenance worksheets.

During the investigation, the aircraft maintenance engineers responsible for the switch installation and functional check indicated that they had completed the work and that the duplicate functional check was conducted with no apparent discrepancies.  The aircraft was then handed over to other maintenance engineers for the completion of further maintenance tasks.  The following day, the scheduled departure of the aircraft was delayed due to on-going maintenance rectifications.  None of these further maintenance tasks involved the aircraft pitch trim system.  Following the delay, the aircraft departed on the occurrence flight after the crew had conducted pre-flight checks, including a check of the pitch trim system cockpit indication for correct operation.  The aircraft maintenance engineers had been assigned the maintenance tasks away from their normal location on a weekend and the aircraft was required for scheduled operations on the Monday morning.

The maintenance contractor and the aircraft operator conducted separate investigations into the trim switch misalignment and concluded that the only plausible scenario leading to the misalignment was that the engineers responsible for the pitch trim switch installation had installed the switch incorrectly. The discrepancy had not been detected during the installation and duplicate functional checks or the flight crew's pre-flight checks.

The type certificate data sheet holder for the aircraft type reported that the aircraft Minimum Equipment List (MEL) provides no relief for flight with one pitch trim system inoperative and so the decision to continue the scheduled flights in this condition was contrary to the requirements of the operator's flight operations manual.

The inadvertent slip by the approach controller was the final action of a number of lapses or omissions that led the pilots to believe that the ILS was available, despite previous advice. The level of situational awareness of the pilots, the fact that the approach was conducted in VMC and the provision of EGPWS warnings were factors that assisted in minimising the failure of other defences in the aviation system.

A Boeing 737-8BK (737), registered VH-VOB, was en route from Melbourne, Victoria to Adelaide, South Australia on descent from flight level 240 to 3,000 ft. The copilot was flying the aircraft under the supervision of the pilot in command. The pilots were aware of a current notice to airmen (NOTAM) advising that the instrument landing system (ILS) was radiating intermittently on test and was not to be used for navigation. The weather at Adelaide was visual meteorological conditions (VMC) and they intended to conduct a visual approach to land on runway 23. During the turn onto final, the pilots observed that the aircraft's flight instruments indicated that the ILS was operating and after the approach controller instructed them to make an ILS approach, they elected to continue with the autopilot selected. About 20 seconds after ILS glide path (GP) capture by the aircraft's flight management system, the aircraft commenced to descend rapidly and its enhanced ground proximity warning system (EGPWS) announced a SINK RATE caution. The pilot in command immediately assumed control of the aircraft and arrested the rate of descent. During that manoeuvre the EGPWS announced a PULL UP warning.

It is normal operator practice for pilots to fly a visual approach using the ILS if possible. The pilots reported that they were in VMC, at about 3,000 ft, after capturing the GP. The pilot in command reported that because they were visual and there were no terrain concerns he used minimal control inputs during the recovery from the descent. Information from the aircraft's flight data recorder indicated that the maximum rate of descent was 6,100 feet per minute and that it had descended to a radio altitude of 1,180 ft above ground level (2,000 ft above mean sea level) before resuming the approach profile.

Technicians were scheduled to work on the runway 23 ILS and there were three NOTAMs issued for the facility. The pilots had the first two NOTAMs issued on the runway 23 ILS but they did not receive the third NOTAM, as they were en route at the time of issue. The initial NOTAM, advising that the glide path was not available, that the pilots had used to brief themselves, remained current. A precondition for the conduct of work on the ILS was the requirement for VMC. An ILS consists of a number of elements including a GP, a localiser (LLZ) and distance markers. It is identified by a two letter code prefixed with the letter I that is included in the LLZ signal (IAD for runway 23 at Adelaide). Following a request for the LLZ to be made available for a training flight, the technicians released the LLZ for operational use while continuing pre-calibration testing of the GP. Consequently, the LLZ (with the ILS identification code) was serviceable while the GP was operating intermittently and not available for operational use.

The controllers in the Adelaide Air Traffic Control tower had previously included information that the LLZ and the GP were not available on the computerised automatic terminal information system (CATIS) that is used to broadcast operational information to pilots. When the LLZ was returned for operational use, they abbreviated the advice to `localiser available', due to system constraints on the amount of additional information that could be included. The information that the GP was not available was not included in the CATIS. The majority of the additional information consisted of advice of restrictions due to aerodrome works. The CATIS was normally broadcast on the non-directional navigation beacon (NDB) and a very high frequency (VHF) radio transmitter. However, at the time of the occurrence the VHF transmitter was not available and the information was only available on the NDB. Despite listening to that information, the pilots missed the fact that the localiser was available due to the poor quality of the received audio. Consequently, when the pilots reported on first contact with the approach controller that they had received the CATIS they were unaware that only the localiser was available.

The approach controller was on his fourth shift in the current shift cycle. After arriving at work, he self briefed and became aware of the ILS maintenance and the non-availability of the GP. A number of control positions were concentrated to a single position and the controller was not initially required to take an operational role. He undertook other duties before returning about 90 minutes later to open the approach east position. His normal practice was to include a text note on the radar display, near the final approach aid or path, of any pertinent operational information, including restrictions or airspace limits. However, on this occasion, following the handover from the approach west controller, he forgot to make a note for his display.

About 40 seconds before the aircraft captured the GP, the controller advised the pilots, `you should get visual shortly, but you're cleared for the 23 ILS approach'. The pilot in command acknowledged and read back that clearance. The controller later reported that at the time, the fact that the glide path was not available had slipped his mind and he reverted to his normal radio telephony phraseology for aircraft on final.

1. Most VSV arm pin retaining flare diameters recorded during the engine disassembly were less than the minimum dimensional limit specified in the engine manufacturer's manual.
2. Two 6th stage VSV arm pins were missing from the synchronizing ring at the 10 o'clock and 12 o'clock positions.
3. Four other vane arm pins were loose in the synchronizing ring.
4. The 6th stage HPC blade failed due to high cyclic stress of the blade in turbulent airflow created by off-schedule stator vane angle.
5. Maintenance actions in accordance with prescribed procedures for VSV arm pin flaring and a published service bulletin, specifically designed to address VSV synchronizing ring pin security via a tack weld process, failed to prevent pin loss.

The documented history of both engines showed that all engine operating parameters matched and were within the expected normal ranges until the loss of thrust event for the right engine, which was observed by the crew and recorded on the Digital Flight Data Recorder.

The damage to the engine was consistent with the failure of a 6th stage HPC blade. This blade, when liberated from its platform, generated debris that caused the trailing edge ricochet damage evident in the 5th stage blades, then travelled downstream in the gas path creating a cascade effect of damage as it collided with other structures.

The absence of integrity in the VSV ring arm pins at the 10 o'clock and 12 o'clock attachment points, due to incorrect or inadequate pin flaring technique and subsequent tack weld breakage, probably allowed the vanes at those positions to move independently of the ring and remain at, or adopt, angles other than those commanded by the engine thrust management computer. This movement was most probably within a range equivalent to the slot machined in the unison ring for the vane arm.

Off-schedule VSV operation can cause higher than normal vibratory stresses in the HPC 6th stage rotor blades and subsequent fatigue fracture of the root attachment. A single vane arm disengaged from the unison ring at any circumferential location, from any of the four variable vane stages, can produce a severe, once per revolution, vibratory excitation within the HPC. These turbulent pulses in turn induced the cyclic stresses on the blades that ultimately led to the blade failure in this occurrence.

The processes of flaring the vane arm pins and then tack welding to provide retention of the pins appears, either to not have been accomplished appropriately by the contracted engine repair facility, or the processes themselves were unsuccessful in meeting their intended purpose. Both of these aspects should be reviewed and addressed by the manufacturer.

Sequence of events

On 28 February 2004, at 1248 western standard time (WST), the crew of a Boeing Company 767-33A, registered V8-RBG, departed Perth, Western Australia, for Brunei on a scheduled passenger flight. While on climb approaching 11,000 ft, the crew heard a loud bang and observed a high exhaust gas temperature (EGT) accompanied by loss of thrust from the right engine. The right EGT was then observed to be over limits¹ at 662 C. The right thrust lever was then closed and all recall items were actioned. The crew declared a MAYDAY² and Rescue and Fire Fighting Services were placed on local standby at Perth Airport. The crew then returned the aircraft to Perth where they conducted a single engine, overweight landing.

Engineers examined the engine and found that the high pressure (N2) section of the engine could not be rotated. There was also evidence of molten metal debris in the exit screens of the 2.95 bleed air valve and some metal spray in the exhaust duct.

The operator carried out an engine change and shipped the removed engine to its engine heavy maintenance contractor in Taiwan. The engine was then disassembled under the supervision of the Aviation Safety Council of Taiwan on behalf of the ATSB.

The ATSB analyzed the flight information from the aircraft's digital flight data recorder (DFDR) and found that the data was consistent with an in flight loss of engine thrust due to an engine failure. Prior to this failure, no anomalies were observed in the performance of the right engine and parameters for both engines were in close agreement.

Engine description

The Pratt & Whitney PW4000 is a third generation, high bypass ratio, commercial turbofan engine (see Figure 1). PW4000 engines have two spools with separate primary and fan duct exhaust systems. The engine has a compression ratio of approximately 30 to 1 and a fan-air to primary-air bypass ratio of approximately 5 to 1.

The low pressure spool (N1) consists of a single stage fan, four-stage low pressure compressor (LPC) and a four-stage low pressure turbine (LPT) on a common shaft.

The high pressure spool (N2) consists of eleven compressor stages driven by a two-stage turbine. The first four stages of the high pressure compressor (HPC) incorporate variable stator vanes (VSV), which are positioned automatically by the full authority digital electronic control.

It utilizes VSV and bleed valves to maintain acceptable stall³ margins throughout the operating envelope of the engine. The VSV are commanded via unison rings and vane arms to change their angle of attack and therefore the angle of attack of the inlet air relative to the compressor blades of the compressor stages at varying engine thrust settings.

Figure 1: PW4000 engine.⁴

The engines also employ Active Clearance Control, which controls the clearances between the turbine blade tips and the turbine case to improve operating efficiency in cruise flight.

Engine disassembly and inspection

Pratt & Whitney engine model PW4056, serial number P724374 was received into the heavy maintenance contractor facility under the supervision of the Aviation Safety Council of Taiwan. The low pressure fan was examined for bird strike damage, with no evidence of such an event found. Upon disassembly, extensive mechanical and thermal damage was noted throughout the engine downstream of the 6th stage compressor. This included severe debris induced airfoil damage to the high pressure compressor (HPC) stages 6 through 15. All 5th stage blades exhibited evidence of blade rubbing at the trailing edge tips, as well as numerous minor debris impact dents to the airfoil trailing edge. The 5th stage VSV also displayed evidence of debris damage to the trailing edges of some vanes. The 5th stage circumferential rub strip had minor abrasion damage; however the 6th and 7th stage rub strips were heavily damaged and were beyond repair limits. The HPC drum also displayed evidence of scraping damage to the knife-edge seals at the 10th and 11th stages.

A boroscope inspection (BSI) was then carried out to inspect the engine gas path. That inspection revealed coating loss and metal spray on the T1 (first stage) nozzle guide vane (NGV). Molten metal was also found blocking several BSI ports.

Associated with this was severe impact damage and molten metal spatter to the stator sections of the HPC, particularly the 5th through 14th rotor stages. With the HPC stator and case halves removed, the entire HPC blade rows were examined from front to rear. One HPC blade at the 6th stage was found to be fractured (see Figure 2). The blade attachment for this blade was observed to be intact and still retained within its dove tail slot in the disc. The broken blade attachment was removed and forwarded to a metallurgical laboratory for further analysis.

Figure 2: 6th stage HPC blade missing.

Further examination of the HPC turbine found burning and erosion of the leading-edge tips and also tip rubbing on the T1 turbine blades, and impact damage from debris to the T2 (second stage) NGV with cracking to the concave face of the airfoil. Molten metal spray was also evident between the diffuser case and outer combustion chamber.

Of note during the engine inspection was the absence of arm attachment pins at the 10 o'clock and 12 o'clock positions for the (VSV) vane arms in the 6th compressor stage synchronizing ring. The remnants of tack-weld material were observed on the synchronizing ring at both locations where the pins were missing.

A pronounced looseness of the remaining four arm attachment pins between these two clock positions on the ring (see Figure 3) was also observed.

Figure 3: VSV synchronising ring and vane arms.

A levelling pin flare diameter check was carried out on the VSV ring assembly during the engine inspection and the measurements recorded for the investigation. It was found, with the exception of one pin, that all pin flare diameters were below the acceptable minimum limit stipulated in the engine manufacturer's table of limits. This flare diameter measurement check, also stipulated in the engine manufacturer's manual⁵, was to be accomplished prior to the tack welding of the pins to the unison ring. The overhaul facility informed the ATSB that there was no stipulation by the manufacturer to check pin flare diameter after the tack weld process had been completed.

Metallurgical examination of failed blade

The blade platform was examined using macro observation, photographic, fractographic and chemical composition techniques⁶.

There was evidence of galling to the attachment element faying surfaces of the blade platform. The surfaces were coated with an anti-galling compound during assembly and this compound appeared to have been degraded during the engine's service life. The blade exhibited two crack origin points; one on the pressure side of the airfoil section measuring 3 mm in height from the platform and the other on the opposite surface measuring 2 mm in height from the platform (see Figure 4).

The fracture sites also exhibited distinct beach marking and were typical of a fatigue type crack propagated under high cyclic stress. The striations present indicated that the cracks, which initiated from either side of the airfoil, were as a result of high alternating vibratory stress.

Figure 4: Crack origins (two places).

Engine history

The engine P724374 had completed 14,914 hours and 3,320 cycles since last overhaul at an approved facility in Singapore. During that shop visit, the VSV synchronizing ring vane arm pins were re-flared after pin replacement. The PW4000 engine family had a history of the pins working loose and allowing vanes to cause turbulent engine airflow when liberated from the synchronizing ring, which drove them. This turbulence was a result of the stator vane being allowed to move away from the scheduled vane angle commanded by the engine EEC. There were also instances recorded by the engine manufacturer of this off-schedule movement resulting in HPC compressor blade stress and in several instances, actual failure of the blade in an identical fashion to the failure observed in this investigation.

The engine manufacturer had published several service bulletins (SBs) requiring repetitive inspections in the interim until a more permanent solution could be applied. These SBs required the synchronizing ring rivet pins to be inspected at regular intervals. During the SB accomplishment process on the occurrence engine, all 23 vane arm pins were removed and replaced with new items.

Pratt & Whitney Service Bulletin PW4ENG 72-432 R3 then called up a tack weld of the pins to secure them in place on the synchronizing ring. This process was re-certified as completed on the occurrence engine after the pin replacement and completion of the flaring. The overhaul facility stated that the pin flaring was accomplished in accordance with the manufacturer's required tolerances. Compliance with this weld process was believed by the manufacturer to remedy the problem of pin wear and looseness, and be terminating action on the repetitive inspection requirements. However, the pins were lost even though they had undergone this tack weld process after their replacement (see Figure 5).

Figure 5: Broken tack weld remnants on synchronising ring.

¹ Type Certificate Data Sheet E24NE specifies a maximum continuous limit of 629o C. The FIM specifies the maximum operating temperature as 650 deg C. At 662 deg C, the temperature was 12 deg C above the maximum operating temperature for a period of 22 seconds. The temperature may have been greater due to EGT Probe blockage and Turbine Cooling Air restriction by molten titanium from the HPC meltdown.
² MAYDAY is the international broadcast for urgent assistance.
³ Stall occurs in a turbine engine when the airflow around the compressor blades degenerates and becomes turbulent, destroying the smooth flow of air through the compressor. It may affect one stage or a group of stages.
⁴ Cutaway source: Pratt & Whitney.
⁵ Pratt & Whitney PW4000 (94"/100") CIR manual (PN 51A357).
⁶ Chung Shan Institute of Science and Technology, Aeronautical Systems Research Division, Aero Materials Section, Taiwan R.O.C Report Number 930107.