Incident
On 11 June 2004, an MD Helicopter (Hughes) 369D (registration C-GWPQ, serial number 700755) was engaged in a lifting operation with a 900-pound sling load when a loud bang occurred, accompanied by a partial engine power loss. The pilot executed a forced landing. During the landing, the helicopter struck the ground and rolled onto its right side, with the main rotor blades still turning. The engine continued to operate on the ground and was subsequently shut down by the pilot. There was no post-impact fire. The pilot later experienced accident-related health issues.
Aircraft and Engine Details
The helicopter was powered by a Rolls Royce Allison 250-C20B modular engine (serial number CAE-836346). The compressor assembly (serial number CAC-35032) had accumulated 7,808.1 hours total time since new and 866.5 hours time since overhaul, with 90.5 hours installed on this airframe. Previously, the compressor had been installed on a Bell 206B helicopter equipped with a bleed air centrifugal particle separator. The accident helicopter was fitted with an FDC/aerofilter particle separator. During the latest engine installation, unremarkable damage was noted on a compressor first-stage rotor blade leading edge, but was considered acceptable by the aircraft maintenance engineer.
Investigation
The Transportation Safety Board of Canada investigated the occurrence. Examination of the compressor revealed high-cycle fatigue fracture on one of the 2nd stage rotor blades. This blade separation was identified as the lead event in the compressor failure. The failure resulted in the shedding of all 2nd and 3rd stage rotor blades and major penetration of the case halves in the 2nd stage rotor path. Analysis of the stator vanes showed no indications of premature failure; vane failures were secondary, resulting from the initial rotor blade failure. This pattern was consistent with other compressor failures of this engine type where stator vane failure was not the initial cause.
Foreign object damage (FOD) was considered as a possible precursor. Evidence of a recent impact mark and a particle of aluminum metal transfer on a 1st stage rotor blade was observed, but it could have occurred after ground impact. An oxidized mark on a front support vane indicated it had been present for some time. A piece of aluminum rivet found among debris could not be positively linked. The investigation concluded that FOD could not be ruled out, but its influence as a fatigue initiator was inconclusive due to the absence of fractographic evidence near the fatigue initiation zone.
Findings
The TSB Engineering Laboratory Report LP 082/2004 documented the compressor failure. The finding as to causes and contributing factors stated that the compressor failure resulted from the separation of a 2nd stage rotor blade due to high cycle fatigue, which had initiated at or near the leading edge of the blade. Post-fracture mechanical damage in the origin area prevented determination of the cause of fatigue initiation. Another finding noted indications of foreign object damage, but its significance as a precursor to the fracture initiation was inconclusive.