Composite aircraft structures inevitably sustain low-velocity blunt-body impacts, resulting in internal damage that is difficult to detect. Existing research primarily focuses on damage analysis of aircraft panel structures under ideal boundary conditions, which do not fully correspond to the actual boundaries of panel structures. This study examines the AC531/CCF800H carbon fiber epoxy resin-based composite skin panel and introduces a design methodology for the elastic support boundary of large curved composite aircraft skin panels. Elastic-support boundary: = 435 KN/m, = 280 KN*m/rad reproduces the measured stress error within 10.44% and displacement error within 8.18%. Peak impact force rises linearly with energy (Fpeak = 1.464 E + 52,406 N, R2 = 0.997) and drops 26% when the angle increases from 0° to 40°. This study delivers a predictive tool that directly supports damage-tolerance compliance for future composite fuselages exposed to ground-handling operations.
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