Fracture Mechanics Analysis and Strength Prediction of Carbon Fiber Composite Laminate with a Delamination

摘要

The use of advanced carbon fibre-reinforced composites in aircraft primary structures has been steadily increasing over past two decades due to their high specific strength and stiffness, and their tailoring as per the need. The composite panels used in primary structures of aircraft are liable to be buckling during its service periods. It is observed that structures can withstand substantial amount of loads after they have buckled. Therefore, an approach to efficiently design the postbuckled composite structures is required to be developed. The designers of the next generation of aircraft are looking into the aspect of postbuckling composite structures to achieve substantial improvements in aircraft structural efficiency. In this work, the postbuckling response and growth of circular delamination in flat and curved composite plates are investigated for different delamination sizes and their locations through the laminate thickness. The prediction of delamination initiation and growth is carried out using the strain energy release rates obtained from the finite element analysis and comparing them to B-K's mixed-mode fracture criterion. The failure load is thus predicted. Predicted results for onset of delamination growth compared well with experimental results. Its variation with different delamination sizes and their locations across panel thickness was also investigated. It is observed that the failure loads are influenced by the delamination sizes depending on their locations across the laminate thickness. The different delamination sizes at H/3 laminate thickness did not have significant effects on the variations of compressive strengths of the delaminated composite panel. But, the compressive strengths of the panels having different delamination sizes at H/2 laminate thickness are more than that at H/3 and increase linearly with increase in delamination sizes.