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The metallic airplane structure fuselage design is characterized by skin, frames, stiffeners, and attachments. In most airplanes, the attachments between these components are made by rivets. The influence of the attachments in the panel behavior under diagonal tension can be verified in the metallic Wagner beam. For stiffened composite panels, like metallic Wagner beams, there is insufficient data about attachment design. In order to design and build lightweight composite structures, the analyst must consider different ways in which the skin is connected to the stiffeners and frames. Therefore, the objective of this paper is to investigate different conceptions of a real-reinforced composite panel used in the aeronautical industry. Experimental and numerical results for strains showed good agreement. The finite element model and the criteria used in the failure analysis are also presented. Comparisons between different panel configurations are made, and conclusions are drawn about attachment efficiency.

Finite element models are proposed to the micromechanical analysis of a representative volume of composite materials. A detailed description of the meshes, boundary conditions, and loadings are presented. An illustrative application is given to evaluate stress amplification factors within a representative volume of the unidirectional carbon fiber composite plate. The results are discussed and compared to the numerical findings.

As shown in the literature, there is plentiful information about sandwich panels. Two of the most common points under discussion are the failure modes and the efficiency of numerical simulations considering the stiffness and interlaminar stress. The failure modes in the literature are not always likely to happen in practice, and representing them becomes a challenging task. Regarding the numerical simulations, new assumptions and formulations appear in order to consider the shear stress in the honeycomb CORE and to minimize processing time in 3D models. Although new mathematical solutions emerge, in some cases they are unpractical for engineering applications and must be evaluated and compared with test results in order to verify their consistency. Therefore, experimental results are necessary to validate theories to comply with the failure modes observed in sandwich panels and to validate the finite element model. Also, the main focus of the literature is on the theoretical formulation and not in engineering applications. In this sense, the main contribution of this paper is to bring forward experimental results of aeronautical sandwich panels whose data are scarce and therefore contributes to the validation of new developments. In addition, the purpose of this work contributes to the use of the finite element models with composite sandwich panels where the appropriate input for 2D (plate) and 3D (solid) elements is unclear. It should be pointed out that for failure investigation the first step is validating the finite element model. In this sense, a typical aircraft panel with experimental results is presented. The finite element model and the input parameters that are not mentioned in the classical literature are also presented. The experimental strain from specimen tested agreed well with the numerical simulations results.

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