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Factual Information

Summary

History of the flight

Approximately eight minutes into a regular passenger transport flight from Melbourne to Sydney, while the Boeing 767 aircraft was climbing through flight level 160, the crew and passengers heard a loud bang and felt severe vibration throughout the airframe. Engine indication and crew alerting system (EICAS) messages on the flight deck indicated the left (number-one) engine had no N1 turbine rotation and an elevated exhaust gas temperature. After discontinuing the climb and advising air traffic services (ATS), the flight crew actioned the 'engine fire, severe damage and separation' checklist and advised the cabin crew and passengers of the engine failure and the intention to return to Melbourne. Several aircraft crewmembers that were passengers aboard the flight advised the flight crew (via the cabin services manager) that the left engine had lost a fan blade and that it had perforated the engine cowling. The flight crew made a PAN radio call to ATS and requested emergency services be placed on local stand-by. After configuring the aircraft for a single-engine approach and landing, some adjustment of the airspeed was required to minimise the level of vibration from the failed engine. The aircraft landed safely on Melbourne airport runway 27, eighteen minutes after the engine had failed and twenty-six minutes after departure.

After exiting the runway, the aircraft was stopped and airport rescue and fire-fighting services carried out a safety inspection before allowing the aircraft to taxi to the terminal buildings using thrust from its remaining serviceable engine. Following passenger disembarkation, the flight crew conducted an operational debriefing with the cabin crew.

Injuries to persons

Injuries Crew Passengers Others Total
Fatal



Serious



Minor



None 10 194 Nil 204

Damage to the aircraft

Damage to the aircraft was limited to the left engine assembly and nacelle. While multiple punctures of the engine nose cowling indicated the liberation of debris from the confines of the intake area, none of this debris had struck the wing, fuselage or tailplane of the aircraft.

Failure of the Pratt & Whitney JT9D-7R4 engine (serial number P-716610) fitted to the aircraft was attributed directly to the fracture and release of the outer half of a single low-pressure compressor (fan) blade (part number 5001341-22, serial number ND9278).

Liberation of the blade segment caused appreciable damage to the remaining fan blades and extensive damage to the intake linings. Ancillary damage to the engine included distortion of the fan casing, loss of the fan speed (N1) sensor and the overload failure of several nose-cowl bolts. Although the primary impact of the released blade with the fan casing had resulted in the segment being contained, the subsequent forward movement of the blade allowed it to impact the nose-cowling with sufficient energy to puncture the cowl wall and escape the engine intake. The initial impact with the cowl occurred at the two-o'clock position (looking forward), with the blade segment passing through the cowl with a tangential trajectory, exiting at around the three-o'clock position. From the impact point and angle, it was evident that the blade segment had been ejected downward and beneath the aircraft. Other debris liberated through the nose cowl or fairings included the N1 sensor and one of the nose cowl lip bolts. Both components were located adjacent to the initial blade impact point and thus were likely to have been subject to a very large reactive force as the blade segment struck the fan case. Figures one to four illustrate the trajectory followed by the released blade segment and the fan case components that perforated the engine cowling.

Aircraft information

Manufacturer Boeing Co.
Model 767-238
Serial number 23896
Registration VH-EAQ
Year of manufacture 1987
Certificate of airworthiness Issue date: 27 August 1987
Certificate of registration Issue date: 27 August 1987

Engine information

The subject engine (serial number P-716610) had been installed on VH-EAQ since October 2001 and had operated for 319 hours and through 200 cycles while fitted to the aircraft. Pratt & Whitney first purchased the engine for leasing in 1998 and, since that time, it had been installed on several different aircraft from different airlines. At the time of failure, the engine had operated for a total of 26,138 hours and through approximately 8,900 cycles.

The failed fan blade (part number 5001341-22, serial number ND9278) was fitted to the engine in August 1998. Before this, the blade had been held as a stock component since its repair and refurbishment in 1991. Work done on the blade at that time included two elevated-temperature straightening operations, where the blade was heated to 650 degrees Celsius and the aerofoil shape re-formed. The manufacturer's records indicated a subsequent blade service life of 7,187 hours and 2,083 cycles. The total time and cycles accumulated by the blade since manufacture was unknown.

Blade inspection

Various non-destructive inspections had been carried out on the blade since overhaul, including eddy current inspections after the thermal straightening operations and periodic visual inspections of the blade while in operational service. Prior to installation on VH-EAQ, the engine underwent a foreign object damage inspection (conducted every 200 cycles) and an eddy current inspection of the leading edge (conducted by the operator every 350 hours). No further inspections had been performed or were required at the time of failure. The requirements and frequency of these on-wing inspections were specified in the aircraft manufacturer's maintenance manual (B767-72-31-02/601) and in Pratt & Whitney service bulletin SB 72-255. At the time of the failure, these documents contained no requirement to carry out a periodic eddy current inspection of the blade trailing edges while the engine was in-service. SB 72-255 stated that 'Eddy current inspection may be used as an option at the operator's discretion'.

After the 1991 refurbishment work, the manufacturer's records indicated that the failed blade was inspected to the engine manual requirements using a single-pass eddy current technique. The eddy current procedure was specified as having the capability to detect crack-like defects as shallow as 0.25mm (0.010") along the blade edges. No defects were detected as a result of this procedure and the blade was subsequently accepted for service.

Cabin aspects

The cabin services manager (CSM) reported the initial engine failure event as "like hitting a brick wall; obviously not turbulence". The CSM described a noisy, high level vibration throughout the cabin, causing some unsteadiness to the crew standing in the cabin service areas. After the vibration had abated, the crew commenced securing the cabin and awaited instruction from the flight deck. Several aircraft crewmembers travelling as passengers reported damage to the left engine nacelle to members of the cabin crew. The CSM passed those observations on to the flight crew. The CSM reported no adverse passenger reactions during the event or during the subsequent return to Melbourne.

Flight recorder

The aircraft was fitted with an L3 Communications (LORAL) model FA2100 solid-state flight data recorder (SSFDR). An excerpt of the data from the recorder containing information from the previous flight and the incident flight was obtained by the ATSB. That data was analysed by the ATSB and used to prepare a summary of events and actions during the incident flight.

The FDR information indicated that the left engine failed at 00:19:59UTC (11:19:59 Eastern Summer Time) and was characterised by a sudden increase in the engine broadband vibration and a decrease in the engine pressure ratio (EPR). At that time, the aircraft was climbing through an altitude of 16,134 feet and maintaining 311 knots airspeed. Both left and right engines were operating at an N1 speed of approximately 94 percent. Vibration levels peaked around two seconds following the initial event and the engine exhaust gas temperature (EGT) peaked at 633 degrees C, six seconds after.

Within the next fourteen seconds, the flight crew had retarded the left engine thrust lever, disengaged the auto-throttle and move the left engine fuel cut-off lever to the OFF position. The left engine fire switch was pulled at 00:21:47, however neither fire bottle was discharged. All actions taken were as documented in the 'Engine fire, severe damage or separation' section of the B767-238 quick reference handbook.

Comparison of the engine broadband vibration levels found no specific differences between the incident flight (before the failure) and the previous flight. Examination of the graphically presented information showed that at approximately twenty seconds before the major vibration transient associated with the fan blade release, a smaller transient occurred in the base vibration levels (figure 5). Short-term escalations in engine vibration levels are anomalous and often indicative of transient events such as compressor aerofoil stalls and surges or foreign object ingestion.

Tests and research

The ATSB examined the released blade segment, assisted by authorised representatives from Pratt & Whitney.

Liberation of the fan blade segment occurred as a direct result of fatigue cracking developing within the trailing edge of the blade aerofoil section. A single transverse high-cycle fatigue crack had developed from a 0.6mm deep pre-existing defect at the blade trailing edge, approximately 290 millimetres above the root face. Multiple surface arrest marks indicated to the growth of the cracking over multiple flight cycles. Final tensile overload of the remaining cross-section released the outer blade section after the fatigue crack had grown to a length of approximately 85 millimetres.

The characteristics of the defect at the fatigue origin identified it as a crack-like feature formed under localised tensile loads. Heat tinting of the defect surfaces indicated the exposure of the region to the elevated temperatures associated with the blade overhaul. The implication from this was that the defect was either present before the overhaul or was produced by the overhaul operations. The defect location was within an area of repair blending at the blade trailing edge. While the blending had reduced the chord-wise width of the blade to one millimetre below the specified minimum limit, it was not considered to have significantly contributed to the development of fatigue cracking from the trailing edge defect. Non-destructive testing procedures carried out following the blade re-work had failed to detect the trailing edge defect before the blade was re-introduced into service within engine P-716610.

A copy of Technical Analysis report number 9/02 detailing the examination of the failed blade is available from the bureau on request.

 
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