Abstract
Accounting for damage tolerance (DT) is crucial during the design process of aerospace composite structures. Typically, a DT design allowable limits the permitted strain level. Calculating this design allowable requires an assessment of the expected damage, the damage detectability, and the residual strength. Current state-of-the-art methods rely on empirical data, offering little flexibility and constraining the design space for a structural optimization. For a tailored calculation of the design allowable for arbitrary laminates and materials, we present an analytical analysis chain, composed from existing methods for the assessment of accidental damage (impact), the damage detectability, and the residual strength. We employ the assembled process in three steps. The determination of a DT design point for a given laminate, the calculation of a laminate-specific allowable, and the obtainment of a ply-share specific and general design allowables through minimization procedures. The highest allowables are achieved by stacking [−45,90,45]n-blocks in the outermost plies while moving the 0∘-plies to the laminate center. Compared to a standard legacy quad configuration, an allowable increase between approx. 30% and 50% was identified. Applied in a structural optimization procedure for a composite wing, this corresponds a mass reduction of up to 5%.
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