The flight load sequence of a combat aircraft was modified by eliminating different levels of small amplitude load excursions to derive several different test load sequences. The fatigue crack growth behavior under all these spectrum load sequence was predicted in a single edge notched tension specimen of an airframe grade La16 aluminum alloy. Crack growth behavior was predicted using a fatigue crack growth law derived from the constant amplitude fatigueudcrack growth tests, incorporating crack closure effects. It was observed that full scale fatigue testing time can be reduced significantly by using of one of the derived test load spectrum without compromising, in general, on the fatigue damage caused by eliminated load cycles.
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