I'm getting negative strength margins in the design of a composite wing spar cap and would like to understand the failure mechanism better. The spar cap component is modeled as a 2-Stack Unstiffened component: the top stack has 4-80 unidirectional plies (representing the cap) while the bottom stack has 8-12 plies of fabric with close to quasi-isotropic layup (representing skin plies). Both are represented as effective laminates. The failure mode is Composite Strength, Tsai-Wu Interaction in the Top Stack. The optimization has moved to the maximum thickness, even though hand calculations indicate that only a moderate thickness should be required.

Is there an easy way to perform a laminate analysis on this effective laminate to the design-to loads for this component in order to determine ply-by-ply stresses, strains and margins and figure out why it is failing? Would it be possible if there was a discrete laminate for this configuration? If so, how can I generate a discrete laminate for this one component?