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The failure of the number-2 engine was attributed to the progressive release of a single blade from the first-stage high-pressure compressor disc, and the subsequent titanium metal fire within the compressor assembly.

Liberation of the blade from the compressor disc was due to the cracking and subsequent failure of the blade's dovetail root corners. The extent of galling damage to the dovetail root surfaces of the retained blades strongly suggests that the fatigue cracking mechanism initiated under the influence of stresses induced by uneven dovetail root bedding. Evidence established by the engine manufacturer regarding the initiation of blade dovetail root cracking is consistent with this conclusion.

The procedures recommended by the service bulletin RB211-72-E181, introduced to improve dovetail root bedding, had not been incorporated as the engine had not become due for refurbishment or overhaul since the bulletin was issued.

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