Aircraft composite structures are mostly joined by mechanical fasteners like bolts, pins or screws. However, the effect of the presence of holes in the remaining strength of the composite structures is still being studied extensively. In this work, epoxy/glass laminates with drilled holes of different sizes were tensile tested and from these results, the residual strength was plotted. Strength vs. hole’s diameter at different fiber orientation was obtained. The fracture path and failure mechanism were identified by fractographic examination. The Point Stress Criterion (PSC) was used, in order to establish the stress intensification due to the presence of a drilled hole. A numerical model by Finite Element Method was carried out to verify the experimental results and the analytic failure predictions. A reduction of 50% in laminate strength was observed when diameter-width ratio was 0.12. The principal fracture mechanism observed in composite laminates was interface breakup. FEM results and analytic results by PSC show accuracy of 90% for predicting the damage in drilled composites.