Description

On 28 May 2012, a Boeing 777-300ER powered by GE90 engines and being operated by Air Canada on a scheduled passenger flight from Toronto to Tokyo Narita experienced sudden failure of the right engine on the initial climb after take off in day Visual Meteorological Conditions (VMC). There were no indications of any associated engine fire and so the failed engine was secured, an emergency declared to ATC and an uneventful return to land made after fuel jettison to reduce landing weight to the maximum permitted. Debris ejected from the exhaust of the failed engine fell onto and damaged several vehicles on the ground but no injuries from this falling debris were reported.

Investigation

An Investigation was carried out by the Canadian TSB. DFDR, Cockpit Voice Recorder (CVR) and Quick Access Recorder download of data covering the period prior to, during and after the failure was successfully downloaded.

It was noted that the sudden failure had occurred as the aircraft was passing 1590 feet agl in manually controlled flight with the EICAS engine fail warning appearing six seconds later when the N2 speed fell below idle. The aircraft had been initially levelled at 3000 feet and the AP engaged so that the required Quick Reference Handbook (QRH) response could be accomplished and the Auxiliary Power Unit started. An emergency was declared to ATC and the climb then continued on the remaining engine, initially to 7000 feet and later to 12000 feet in order to jettison the 86.6 tonnes of fuel needed to get the ELW within the MLW. The return to Toronto was uneventful and shortly after touch down, partial thrust reverser was selected on the operating engine to assist deceleration. The aircraft came to a stop on the runway and the Rescue and Fire Fighting Services attended and chocked the main landing gear wheels to allow brake release to facilitate cooling before the aircraft was taxied to the parking gate. The total time airborne was 86 minutes. Since a review of the flight crew response showed that it had been in accordance with relevant company procedures for dealing with an engine failure, attention was focused on the engine and the reason for its failure.

An initial inspection of the aircraft after the engine failure flight found that the carbon fibre underside of the right wing flaperon had been damaged by debris ejected from the failed engine. The engine exhaust cone was found to contain pooled engine oil and small fragments of both high-pressure turbine (HPT) and low-pressure turbine (LPT) material. “Most of the damage to the LPT was located towards the outer periphery” and it was concluded that “this damage was consistent with excessive heat and impact from debris originating from forward of the LPT stage.” All engine damage was contained and there was no damage to the engine cowlings, although evidence of heat damage to the HPT module was found.

The debris ejected from the engine exhaust when the failure occurred was located and retrieved. It was found to correspond to similar fragments found in the engine exhaust cone. A subsequent examination of the failed engine off-wing led to the conclusion that the engine core had seized as a result of the ingestion of several HPT stage 2 nozzle fragments jammed in different locations, which had collectively prevented the core from rotating freely. It was confirmed that there was no indication of a bird strike or any component failure forward of the HPT module.

It was found that the service and maintenance history of the engine indicated that it had been maintained in accordance with current regulations and engine manufacturer procedures including regular inspections under SB 72-0401 which had been issued by GE in 2010 following in-service problems with the high-pressure turbine (HPT) stage 1 shrouds on some GE90 engines. The engine which failed had been identified as one of the suspect engines listed in this SB and therefore subject to repetitive borescope inspections (BSI).

The origin of the problem with these shrouds was already suspected - and had subsequently been confirmed - to be a result of a particular feature of the manufacturing process in which their cooling holes had been drilled with a newly introduced high-intensity laser which “had produced a variation in the shape of the cooling holes in the shrouds, which accelerated the deterioration of the shrouds and reduced their structural integrity”. It was found during teardown inspection of the failed engine that specific features of the normal gas flow in the area of shroud number 33 had led to a more rapid deterioration than anticipated. Substantial damage to this shroud and to its corresponding hanger was found, as well as damage to both number 26 and number 27 shrouds. It was concluded that “the extent of the damage suggests that it was likely present but misevaluated or overlooked during the previous BSI”.

GE determined that the origin of initial shroud distress on engines with similar failures had been shroud 33 and suspected that “as the hole in (this) shroud reaches a critical size, the fuel-air mixture is ingested behind the shroud and is introduced to a larger volume of air, thus creating a super-heated zone that leads to rapid distress of the stage 1 shroud hanger and stage 2 nozzle outer band” with release of the stage 2 outer band causing downstream damage.

The Investigation found as follows in respect of Causes and Contributing Factors

1. During shroud production, a change to a higher-intensity laser resulted in a variation in the shape and size of the shroud cooling holes. Over a period of time in service, these cooling holes eroded, which resulted in both degraded shroud cooling and a super-heated zone. This, in turn, increased the rate of erosion until the shroud integrity was reduced to the point of failure.
2. Damage to high-pressure turbine shrouds and hangers, which was likely present during the last borescope inspection, went undetected prior to the occurrence. As a result, the engine was not removed from service.
3. The number 2 engine shut down during the initial climb-out due to a failure of the high-pressure turbine stage 1 shroud.

Safety Action taken in response to the ongoing Investigation was notified as follows:

• Air Canada took action to re-inspect all engines subject to SB 72-0401 in order to establish a new ‘baseline’. A total of fifteen in-service engines were found with shroud deterioration which would qualify for re-inspection in accordance with the criteria in the SB and three engines displayed sufficient shroud deterioration to be identified as needing a reduced inspection interval. Two other engines were found to have sufficient shroud damage to require removal from service.
• Air Canada performed its own risk management exercise in order to evaluate the condition of the remaining engines affected by SB 72-0401 and establish appropriate prioritisation for their removal and repair. This assessment included the way in which engine pairing was permitted as a means to reduce the probability of a double engine failure. After completing this work and evaluating the results, it was decided to reduce the number of cycles between BSIs by 50% or more from the schedule contained in SB 72-0401.
• The FAA issued an AD 2013-17-07 in respect of borescope inspections GE90 engines fitted with a specific P/N range of Stage 1 HPT Stator Shrouds which was effective from 18 October 2013. This AD required the use of specified GE90-100 SBs as guidance for the inspection procedures and to determine if any shroud hole dimensions made the inspected engine unacceptable for continued service.

The Final Report of the Investigation was authorised for release by the Board on 6 November 2013 and officially released on 13 December 2013.

Further Reading

• Engine Failure During Takeoff - Multi-Engine Transport Category Jet Aircraft
• Engine Failure at Takeoff (OGHFA SE)
• Uncontained Engine Failure
• Engine Failure: Guidance for Controllers
• Damage Tolerance
• Certification of Aircraft, Design and Production