Investigation into the failure of aircraft stringer–skin composite structure under bending load

The progressive failure in the top-hat stiffened composite panel under four-point bending damage was analyzed. The cohesive element was applied to simulate the debonding between the fuselage skin and stringer. In addition, the quadratic stress criterion and B–K criterion were adopted to predict initiation and propagation of the debondings. The propagation process, failure mode and ultimate load of the stiffened composite panel under four-point bending were predicted and compared with the test results. Specimens with co-bonded and co-cured processes were considered in this study. In regard to the co-bonded specimens, there existed only delamination between skin and stringer. The error between the predicted and experimental results of debonding load was about 2–8.8%. In terms of the co-cured specimens, the debonding between skin and stringer and the intra-layer cracks both can be observed. © 2023, Shanghai Jiao Tong University.