A building block pyramid is designed used to evaluate a composite aircraft structure, combining tests and analyses, using stiffened panels, single-stringer specimens and coupons. Failure in these structures can be caused by postbuckling-induced skin-stringer separation, which is complex, involving matrix cracks, fibre bridges, and delamination migrations. Standardized tests covering postbuckling-induced skin-stringer separation are lacking. First, a material characterization is performed on the coupon level. These properties are applied to a stiffened panel model to identify critical postbuckling regions. Then a single-stringer specimen is designed that combines the material complexities of the coupon level and the geometrical complexities of the panel by mimicking the postbuckling shape. Specifically a seven-point bending configuration to study bending-induced separation and a four-point twisting configuration for skin twisting-induced separation. The guidelines for modelling and testing can assist in standardising these test methods. Using composite materials in aircraft structures can reduce weight compared to conventional metals. However, utilizing more of the material's load-carrying capabilities can further reduce weight.@en

A methodology is presented to investigate and improve the strength and damage tolerance of stiffened composite panels used in aerospace structures subjected to postbuckling deformation. These structural panels have the capability to operate in the postbuckling field, but the possible interaction between the postbuckling deformation and the damage initiation and propagation is yet to be fully understood. The developed methodology considers single-stringer specimens representative of stiffened panels to analyze skin-stringer separation. In this paper single-stringer specimens are studied in a four-point twisting configuration in order to investigate the region of maximum twisting, where the separation between the skin and the stiffener can initiate. A new test set-up is presented that recreates the four-point layout that can trigger separation due to twisting. The applied methodology shows that it is possible to mimic the out-of-plane buckling deformation of a large panel and study this numerically and experimentally through a single-stringer specimen.@en

Aeronautical composite stiffened structures have the capability to carry loads deep into postbuckling, yet they are typically designed to operate below the buckling load to avoid potential issues with durability and structural integrity. Large out-of-plane postbuckling deformation of the skin can result in the opening of the skin-stringer interfaces, especially in the presence of defects, such as impact damage. To ensure that skin-stringer separation does not propagate in an unstable mode that can cause a complete collapse of the structure, a deeper understanding of the interaction between the postbuckling deformation and the development of damage is required. The present study represents a first step towards a methodology based on analysis and experiments to assess and improve the strength, life, and damage tolerance of stiffened composite structures subjected to postbuckling deformations. Two regions were identified in a four-stringer panel in which skin-stringer separation can occur, namely the region of maximum deflection and the region of maximum twisting. Both regions have been studied using a finite element model of a representative single-stringer specimen. For the region of maximum deflection, a seven-point bending configuration was used, in which five supports and two loading points induce buckling waves to the specimen. The region of maximum twisting was studied using an edge crack torsion configuration, with two supports and two loading points. These two configurations were studied by changing the positions of the supports and the loading points. An optimization procedure was carried out to minimize the error between the out-of-plane deformation of the representative single-stringer specimen and the corresponding region of the fourstringer panel. @en

To design aeronautical composite multi-stringer panels that can safely operate in a postbuckled state, it is important to identify the parameters that can influence the different modes in which skin-stringer separation might occur. A methodology is under development to study the interaction between the skin-stringer separation and the postbuckling deformation using the building block approach and single-stringer specimens. In particular, the methodology can identify whether the skin-stringer separation occurs due to bending or twisting, so that these two possible modes can be studied separately. For bending, a simple criterion that can predict the location of initiation is presented. This procedure has the potential to reduce the overall development cost and allows the investigation of the design parameters that influence the skin-stringer separation. @en