Following take-off from Los Angeles International Airport (LAX), the crew of the B747-438 aircraft noticed a severe airframe jolt while conducting a climbing left turn. The cockpit instruments indicated that the number 1 engine exhaust gas temperature was rising through 900 degrees C. Passengers also reported flames emanating from the number 1 engine tailpipe.
The crew shut down the number 1 engine and returned the aircraft to LAX for a one- engine inoperative landing.
An initial investigation carried out by the operator's maintenance personnel revealed that there had been an apparent failure within the engine's high-pressure compressor (HPC) assembly. The engine was removed from the aircraft and transported back to the operator's engine maintenance facility in Australia, where a more detailed examination was carried out. An Australian Transport Safety Bureau (ATSB) metallurgist was present for that examination.
The engine was a Rolls Royce RB211-524G2-T-19/15 turbofan engine. The designation `T', in the engine model number, indicated that the engine had been manufactured with a core engine from the larger Rolls Royce `Trent' engine series. The inclusion of the `Trent' core had enabled the engine to be more fuel-efficient and operate at a lower exhaust gas temperature.
The `Trent' core engine was split into numbered modules. The three modules of interest to the investigation were: Module 33, the intermediate case module; Module 41, the high-pressure system module; and Module 51, the intermediate and low-pressure turbine assembly module.
The engine had a nominal overhaul life of 30,000 hours or 4,000 cycles. At the time of the failure the engine was well within its overhaul life, having been in operation for a total of 13,922 hours and 1,395 cycles. It had not undergone any major maintenance.
The ATSB Technical Analysis report on the engine failure, (see Appendix A), indicated that the engine failure had resulted from the liberation of a single blade from the first-stage HPC rotor in Module 41. The blade release had resulted in extensive damage to the engine. The friction from the liberated blade impacting the surrounding blades on the HPC rotor resulted in a titanium fire within the compressor assembly.
A close inspection of the remains of the liberated blade root stub showed evidence of fatigue cracking and loss of the forward trailing edge corner of the blade dovetail root block. None of the other blades within the first-stage HPC assembly showed any visible evidence of cracking when inspected with the unaided eye. Minor collateral damage was also evident to components in Module 51 and Module 33 resulting from the blade failure.
The manufacturer was aware of three similar failures of the HPC blades in the RB211-524G/H-T series of turbofan engines. In those failures, cracking in the blade root area was believed to have resulted from uneven friction on the blade root bedding surfaces due to a breakdown in blade root lubricant. The manufacturer further indicated that damage in the blade root area that led to local stress concentrations, such as scores and sharp edges, might also have contributed to those blade failures.
The evidence from the ATSB investigation indicated that the failure mode in this incident was the same as the three other known failures in RB211-524G/H-T turbofan engines.