Search Results (2)

Aircraft structures must be capable of performing their function throughout their design life while meeting safety objectives. Such structures may contain defects and/or damages that can occur for several reasons. Therefore, aircraft structures are inspected regularly and repaired if necessary. The concept of combining an inspection plan with knowledge of damage threats, damage growth rates, and residual strength is referred to as “damage-tolerant design” in the field of aircraft design. In the present study, we fabricated a composite panel with a cutout (which is generally found in the bottom skin of the wing) using a resin infusion process and studied the damage tolerance of a co-cured skin-stringer composite panel. The composite panel was subjected to low-velocity impact damage, and the extent of damage was studied based on non-destructive inspection techniques such as ultrasonic inspection. Fixtures were designed and fabricated to load the composite panel under static and fatigue loads. Finally, the panel was tested under tensile and fatigue loads (mini TWIST). Deformations and strains obtained from FE simulations were compared and verified against test data. Results show that the impact damages considered in this study did not alter the load path in the composite panel. Damage did not occur under the application of one block (10% life) of spectrum fatigue loads. The damage tolerance of the stiffened skin composite panel was demonstrated through test and analysis. Full article

Aeronautical stiffened panels composed of thin shells and beams are prone to deformation or buckling due to the combined loading, functional boundary conditions and interface forces between joined parts in the assembly processes. In this paper, a mechanical prediction model of the multi-component panel is presented to investigate the deformation propagation, which has a significant effect on the fatigue life of built-up structures. Governing equations of Kirchhoff–Love shell are established, of which displacement expressions are transformed into Fourier series expansions of several introduced potential functions by applying the Galerkin approach. This paper presents an intermediate quantity, concentrated force at the joining interface, to describe mechanical interactions between the coupled components. Based on the Euler–Bernoulli beam theory, unknown intermediate quantity is calculated by solving a 3D stringer deformation equation with static boundary conditions specified on joining points. Compared with the finite element simulation and integrated model, the proposed method can substantially reduce grid number without jeopardizing the prediction accuracy. Practical experiment of the aircraft panel assembly is also performed to obtain the measured data. Maximum deviation between the experimental and predicted clearance values is 0.193 mm, which is enough to meet the requirement for predicting dimensional variations of the aircraft panel assembly. Full article