Abstract: Carbon fiber reinforced polymer (CFRP) is gradually used in aircraft main load-bearing structure due to its good mechanical properties. Delamination is one of the most common damage forms of composite laminates, and it is also the focus of damage tolerance design and analysis of aircraft composite structures. For mode II delamination, the existing research mainly focuses on unidirectional laminates, while multidirectional laminates are widely used in engineering practice, and there is a lack of in-depth understanding of the influence of ply angle on mode II delamination. Therefore, experimental and numerical investigations were conducted in this paper. Firstly, three different interfaces (0°/5°, 45°/−45° and 90°/90°) were designed for composite laminates made of T700/QY9511 carbon fiber/bismaleimide prepregs, in order to reduce the coupling effect and ensure the mode II dominant delamination propagation. The mode II delamination test was carried out using end-notched flexure device and the fracture toughness was measured. The results show that the ply angle has a significant effect on the fracture toughness and delamination failure behavior. The control variable method is used for establishing the key parameter model of cohesive zone model. On this basis, the simulation of mode II delamination propagation behavior of multidirectional laminates is realized. The predicted load-displacement curve response is in good agreement with the experimental results, which shows the effectiveness of the finite element model. The finite element results show that the energy dissipated by the matrix damage increases with the increase of interfacial angle. In order to reveal the relationship between the energy dissipated by matrix damage and the size of damage area around the crack tip, the damage area around the crack tip of specimens with different interfaces was simulated by using user-defined subroutine.