Abstract: The failure mechanism of the composite fuselage curved panel under circumferential bending load was studied by four-point bending test and finite element analysis (FEA). A set of strengthening and connecting fixture was designed to avoid undesirable failure. A finite element model was also established, in which cohesive element was used to simulate the interface between the hat stringer and skin. Quads criterion and Hashin criterion were used as failure criteria of the interface and the laminate respectively. The results obtained by experiment and FEA are in good agreement. It can be concluded the crack initiates from the R-zone of the bonding area between the hat stringer and skin, due to local buckling of the skin at the bottom of the hat stringer. Then it extends and causes debonding of the hat stringers. With the extension of skin buckling and stringer debonding, the skin shows global instability and loses its bearing capacity, which eventually leads to the failure of the frame due to excessive load. According to the initial damage mode, the full around bonding process between the hat stringer and skin is adopted to improve the circumferential stability of the curved composite panel. The bending test result shows that the bending loads of the initial buckling and overall failure increase 21.9%and 16.8%, respectively.