In order to accelerate full scale fatigue testing, the original flight load spectrum of a combat aircraft was modified by eliminating small amplitude load excursions lower than a prefixed filter value. Different levels of filter value were used to derive several different test loadudspectrum having lesser number of fatigue cycles than the flight load spectrum. The fatigue crack growth behavior under all these spectrum load sequences was predicted in two different airframe materials viz., D16 aluminum and Ti6A14V alloys using a fatigue crack growth model derived from constant amplitude fatigue crack growth tests, which incorporated crack closure effects. Results show that the basis of spectrum modification should not only be the elimination of small amplitude load cycles but also on the fatigue crack growth behavior of the structural materials used in airframe construction.
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