In an aircraft structure with fastener holes, stress concentrations that occur near the holes may dominate the design consideration. Softening strips have been tested to reduce these stress concentrations in laminated composites [1, 2]. In the region containing the hole, the high-modulus composite material is replaced with low-modulus, high-failure-strain composite. For example, in a [0, ± 45, 90] graphite/epoxy laminate, a softening strip can be made by replacing the 0° plies with 0° glass/epoxy composite plies. Experiments have shown that a softening strip can increase the static strength over the plain laminate by as much as 20% [2]. This paper reports an analytical study of softening strips in various laminates. The maximum stress criterion was used to predict the first ply failure. We present the calculated results first for the plain, baseline laminates, then for laminates with softening strips.