During the early stages of a regular passenger transport flight between Melbourne and Sydney, Boeing 767 aircraft, registration VH-EAQ, sustained the failure of the left engine, necessitating a return to Melbourne airport. The turn-back and subsequent single-engine landing was uneventful.
Failure of the left engine resulted from the liberation of the outer-most half of a single first-stage compressor (fan) blade. The fan casing had contained the initial blade impact, however subsequent forward movement of the segment allowed it to strike and penetrate the engine intake (nose) cowling. Other small components had also penetrated the inboard fan case cowl. Inspections found no evidence that any of the released debris had damaged the aircraft structure outside of the engine nacelle.
Laboratory examination of the section of failed blade remaining within the fan rotor found that the fracture had occurred from high-cycle fatigue cracking that initiated from a pre-existing defect at the blade trailing edge. This crack-like defect showed evidence of having formed during or before the blade was last overhauled in 1991 and had remained undetected during post-overhaul non-destructive testing inspections. The manufacturer fitted the blade to the subject engine in 1998 and it remained in service until the failure, accumulating 7,187 hours and 2,083 cycles since overhaul.
The engine manufacturer attributed the failure to detect the original fan blade defect to procedural deficiencies and operator error during inspection. In response to previous blade failures, the manufacturer amended the engine manuals to incorporate a number of improvements aimed at increasing the probability of defect detection. The failed blade in this occurrence had been inspected before the engine manual changes. The engine manufacturer has also issued an all-operator communication recommending that any blades last inspected before the engine manual changes should be re-inspected to the latest requirements.
(Damage confined to number one engine and cowling. Loss of single fan blade produced multiple punctures of nose cowling and smaller preforations in fan case cowling. Extensive impact damage to remaining fan blades and cowl linings.)
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
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.
|Year of manufacture||1987|
|Certificate of airworthiness||Issue date: 27 August 1987|
|Certificate of registration||Issue date: 27 August 1987|
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.
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.
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.
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.
Failure of the left engine from VH-EAQ occurred as a result of the fracture and liberation of approximately two-thirds of the length of a single first-stage low-pressure compressor (fan) blade. The loss of the blade section produced a significant imbalance in the fan rotor, which manifested as severe vibration of the airframe and produced heavy tip rub on the remaining blades against the fan case lining. The flight crew's subsequent observations of a high exhaust gas temperature indicated the development of anomalous combustion conditions within the engine because of the airflow interruptions produced by the fan failure.
Rotor kinematic laws predict that the partial loss of a fan blade will result in the fragment striking the fan case, before folding flat and moving forward from the plane of rotation with a helical motion. This motion will continue until the fragment either perforates the intake cowling forward of the fan case, exits the front of the intake cowling, or is drawn back into and re-ingested by the fan rotor. In this case, damage to the intake (nose) cowling indicated the loss of the blade section soon after liberation, with comparatively little other damage to the remaining blades. Measurements of the damage to the intake cowling indicated the blade exited the cowling with a forward and downward trajectory, sufficient to take it away from the aircraft without impacting any other part of the structure.
The radial forces transferred to the fan case by the initial impact of the blade segment were sufficient to break away several of the nose cowl bolts, one of which punctured the inboard fan case cowl. The fan speed (N1) sensor was also lost in a similar manner. Neither of these components had damaged the aircraft after exiting the engine nacelle.
Laboratory examination of the blade fracture surface confirmed the presence of a pre-existing trailing edge defect, from which high-cycle fatigue cracking initiated and propagated. The examination identified the defect was either produced by, or was present before the last major blade refurbishment operation incorporating an elevated temperature straightening operation. The size and nature of the defect was such that it should have been detectable by non-destructive means following the blade refurbishment operations. The engine manufacturer stated that the eddy-current method specified for this inspection had the capability to detect defects well below the size of the actual defect present. In this regard therefore, error by the inspecting technician was the most likely factor contributing to the failure to detect the defect.
Pratt & Whitney service bulletin SB 72-255 was the core document that specified the requirements for the periodic in-service inspection of the engine low-pressure compressor blades. The objective of that service bulletin was to provide opportunities for the early discovery and repair of foreign object damage, thereby reducing the potential for foreign object damage induced blade failures. The bulletin required the visual inspection of the fan blade assemblies and the blend repair of all leading edge damage found, with eddy current inspection included as an option at the operator's discretion. In this case however, visual inspection alone would have likely proved unsuccessful in detecting the defect at the origin of fatigue cracking, due to the small defect size and the absence of any associated foreign object damage.
In the current engine operator's case, an eddy current technique was used to complement the visual inspection, however this was a limited survey and did not extend to the examination of the blade trailing edges. It was not known whether any of the previous engine operators had used an eddy current inspection as part of their compliance with SB 72-255.
The US Federal Aviation Administration has published a draft advisory circular that provides a mechanism for the assessment of the continued airworthiness of powerplants and auxiliary power units on transport category aircraft (AC39-XX). The advisory circular describes the Continued Airworthiness Assessment Methodologies (CAAM) and uses them to identify unsafe conditions, before prescribing corrective actions in accordance with the Federal Aviation Regulations (FAR) part 39.
The CAAM recognise uncontained engine blade failures as an 'historically unsafe condition' and as such, require that an appropriate response be determined and carried out. In this case, the unsafe condition may be more specifically defined as the presence of undetected defects within the blade trailing edges, from which uncontained failure may result. Following from this, aircraft exposed to this unsafe condition are defined as those aircraft fitted with engines carrying defective fan blades. In the context of this occurrence, the assessment of the level of exposure (ie. number of aircraft) can only be determined by the inspection of each blade currently in service, to determine the presence or otherwise of the defect/s. Minimisation of the risk is achieved by the subsequent removal of all blades found to contain defects.
This action is proposed by the engine manufacturer and is detailed within section 4 of this report (Safety Action).
- A small crack-like defect remained within the trailing edge of a first-stage low-pressure compressor (fan) blade after the component had undergone overhaul operations in 1991.
- Post-overhaul non-destructive inspection procedures conducted on the blade edges failed to detect the defect and the blade was placed in storage until 1998, when it was installed into engine P-716610.
- Because of its small size, the trailing edge defect was not detectable by the in-service visual inspections required by the manufacturer and carried out by the various engine operators (SB 72-255).
- An optional eddy-current inspection of the blade by the last operator was capable of detecting the defect, but was not performed (nor required to be performed) along the blade trailing edge.
- Fatigue cracking initiated and propagated from the trailing edge defect in response to vibratory and centrifugal operating loads.
- Fracture of the fan blade occurred after growth of the fatigue cracking to critical size.
- The left engine of the aircraft failed after the fracture and liberation of the blade segment during climb to cruising altitude.
Local safety action
In response to other JT9D blade failures from trailing edge cracks, the engine manufacturer implemented improvements to the eddy-current techniques used to inspect the blades after overhaul. The improvements included the addition of a new eddy-current inspection requirement after fan blade patch repairs, changes to the probe scanning methods and the use of a chart-recorder device to produce a 'hard-copy' of the test results for post-inspection review. The blade fracture on engine P-716610 occurred on a component that was repaired and inspected before the technique improvements and relevant engine manual changes were made.
The engine manufacturer is revising the relevant engine manuals to include the use of an enhanced 3-pass eddy current inspection procedure in lieu of the single pass procedure. This procedure provides a greater degree of confidence for the detection of small defects. The issue of these engine manual revisions is planned for early 2003.
An all-operator communication (AOW) was issued by the engine manufacturer on 19 July 2002, providing for a temporary revision of the engine manuals to incorporate the 3-pass eddy current inspection procedure, prior to the full revision of the engine manuals as mentioned above. The AOW also recommended the re-inspection of all fan blade leading and trailing edges at next overhaul, using the 3-pass eddy current procedure.
After the engine failure, the aircraft operator implemented a once-off fleet-wide inspection of the fan blade trailing edges using an eddy-current technique. No other similar defects were found. The operator has also implemented an engineering instruction requiring that all new or lease engines introduced into the fleet will automatically be flagged as requiring a blade trailing edge eddy-current inspection.
|Date:||27 November 2001||Investigation status:||Completed|
|Time:||1122 hours ESuT|
|Location:||56 km NE Melbourne, Aero.|
|Release date:||02 October 2002||Occurrence category:||Incident|
|Report status:||Final||Highest injury level:||None|
|Aircraft manufacturer||The Boeing Company|
|Type of operation||Air Transport High Capacity|
|Damage to aircraft||Minor|
|Departure point||Melbourne, VIC|
|Departure time||0014 hours ESuT|