On 5 May 2008, a Boeing 747-300 being operated by Saudi Arabian Airlines on a non revenue positioning flight from Jeddah to Jakarta for maintenance input experienced an uncontained failure of no 1 engine shortly after take off and subsequently made an uneventful return to land. The damage to the failed engine was externally visible with part of the engine clearly missing.
An Investigation was begun by the General Authority of Civil Aviation (GACA) of the Kingdom of Saudi Arabia, but was delegated to the United States National Transportation Safety Board (USA) (NTSB) on 11 April 2009.
The flight crew reported that they had reduced the No. 1 engine thrust soon after takeoff after seeing fluctuations in its EGT and N1 indications. However, the fluctuations had continued and about a minute later, at about 1100 feet agl, the same engine’s low oil pressure warning illuminated and it was seen that the oil quantity indicator read zero. The engine was shut down, fuel dumped and a return to Jeddah made. Post-flight inspection of the airplane found that the aft end of the No. 1 engine was missing and that the aircraft left wing and flaps had been damaged by impact penetration of debris from the failed engine. The liberated engine parts were recovered from a position about 2½ miles from the upwind end of the take off runway.
It was concluded that the low pressure turbine (LPT) stage 3 disc had separated at the forward spacer arm of the GE CF6-50 engine and that all components aft of that separation had been ejected.
Examination of the fracture surfaces of the failed disc established that it had separated circumferentially near the fillet between the spacer arm and the disk rim. The fracture surface had been damaged by post-fracture smearing but “interpretable areas showed damage consistent with high amplitude/high cycle fatigue from multiple initiation sites, primarily on the inner diameter”. The evidence suggested that “once initiated, the cracks propagated rapidly through the spacer arm thickness and joined to form a single circumferential crack, resulting in disc separation”.
According to GE, when a high level of high pressure turbine rotor unbalance occurs in the engine involved, the resulting synchronous vibration forces can interact with the engine’s LPT through a common bearing support and excite a bladed-disc vibration mode within the engine operating range. A borescope inspection of the High Pressure Turbine (HPT) of the engine found that three stage 1 blades over a nine-blade sector were partially missing, the missing material being equivalent to about 1.8 fan blades. This loss of material was, according to GE, enough to cause HPT rotor unbalance which would be large enough to excite the bladed-disc vibration mode in the LPT Stage 3 disc.
The Probable Cause of the failure was determined as:
Failure of the low pressure turbine stage 3 disc due to a design that is vulnerable to high pressure turbine unbalance-induced synchronous vibration that cannot be detected in flight and the subsequent uncontained engine failure.
The Final Report ENG08IA030 was adopted on 19 December 2012. No Safety Recommendations were made.