Skin-stringer interface failure investigation of stringer-stiffened curved composite panels under hail ice impact

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Large scale aircraft composite panels can be damaged by hail ice in high speed ductile impacts (without penetration by sharp shape) and predicting the damage modes and failure mechanisms under these complex boundaries and loading conditions poses a unique challenge. Thus, the dynamic behavior and failure of stringer-stiffened curved carbon/epoxy composite panels impacted by hail ice at two classic locations (mid-bay of impact type I and mid-flange of impact type II) with different velocities have been investigated in this research via experimental and numerical methods. The results show that panel damage from ice sphere impacts is a stress wave dominated dynamic response and the initial delamination of the panel always occurs at the skin-stringer interface when the ice sphere reaches a minimum threshold impact velocity irrespective of the loading site. The finite element methodology employing rate sensitive ice models and composite laminates with failure is capable of accurately predicting, the delamination site, the total delamination area versus impact velocity, and the separation of the skin-stringer bonded joints which may be located away from the impact site at areas not typically examined. The numerical simulation has been validated at two different impact sites and provides insight into stress wave propagation despite non-uniform thickness and stiffness variations of the stringer flange bonded to the composite skin. These experimental and finite element methodologies can be applied to improve the design of new airframes, composite joints, and improve the accuracy of residual strength evaluation of existing aircraft panels under hail ice impacts.