The study presents results of experimental examination of a model representing a fragment of an aircraft wing structure with the skin made of a glass fibre/epoxy composite. For such a system, the deformation pattern has been found and the representative equilibrium path determined. The finite element method has been used to develop the corresponding numerical model, the correctness of which has been then verified by comparing the obtained results with the course of the relevant experiment. Conformity of the results allowed one to determine usefulness of the applied methods in the assessment of mechanical properties of modified solutions involving integral skin stiffening elements.

Article citation info: (*) Tekst artykułu w polskiej wersji językowej dostępny w elektronicznym wydaniu kwartalnika na stronie www.ein.org.pl IntroductionThe dynamic evolution of solutions used by designers of aircraft load-bearing structures, initiated in the 1920s, resulted in establishment of aircraft operation standards based on assumptions which would be unacceptable for any other type of engineering structures. One of such assumptions is the admissibility of post-buckling deformations of semi-monocoque skin structures provided that the loss of stability is of local and linear-elastic nature [1,2]. The rule applies mainly to skins made of isotropic materials, because in case of glass, carbon, and aramid composites which are used in the aircraft industry for a relatively short period of time, destruction processes occurring in the course of their prolonged operation in post-buckling deformation conditions are still the subject of numerous studies [6-8, 10, 13, 19].Although composites become more and more popular in the aircraft technology, aluminium alloys still remain the materials most universally used by the industry for their well-known mechanical properties and high reliability. Application of these materials for aircraft skins was initially connected with some distrust as designers traditionally strove after elimination of post-buckling deformations in the first place. In such cases, the need to increase skin thickness emerged resulting in inevitable increase of the overall mass of the structure. For some time, the problem had been solved by using corrugated sheet metal for fuselage skins. This technology was commonly applied only in the beginning of 1930s, e.g. in designs of such aeroplane manufacturers as Ford and Junkers. With improvement of aircraft performance parameters, such structures became more and more troublesome because of the related aerodynamic problems, as a result of which it became necessary to coma back to smooth skin materials and admit the possibility of occurrence of local post-buckling deformations [11,14].While in case of skin fragments without geometrical singularities, the above-mentioned phenomenon does not result in any decrease of their operating durability stability, presence of cut-outs of any type becomes a source of problems. Aircraft structures, by virtue of for what they are intended and in what conditions are to be operated, are characterised with existence of a large number of cut-outs with various dimensions. Such discontinuities may occur within segments of skin of a semi-monocoque structure limited by skeleton components (e.g. passenger cabin windows or small inspection openings). They can also represent major discontinuities of the structure with larger dimensions, e.g. doors, cargo loading hatches, or large access openings ( Fig. 1), which make segmentation of stringers necessary.From the point of view of strength properties and operating durability of a structure, presence of cut-outs is a very unfavourable circumstance. They reduce overall rigidity of the str...

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The aim of the study was to determine the impact of the use of isogrid stiffeners on the stress and displacement distribution of a thin-walled cylindrical shell made of layered composites subjected to torsion. It also strives to define criteria for assessing the results of non-linear numerical analysis of models of the examined structures by comparing them with the results of the model experiment. The study contains the results of experimental research using models made of glass–epoxy composites and the results of numerical analyses in non-linear terms. The experiment was carried out using a special test stand. The research involved two types of considered structures. The results of the research allowed to create the concept of an adequate numerical model in terms of the finite element method, allowing to determine the distribution of stress and strain in the components of the studied structures. Simultaneously, the obtained conformity between the results of non-linear numerical analyses and the experiment allows to consider the results of analyses of the modified model in order to determine the properties of different stiffening variants as reliable. The presented research allows to determine the nature of the deformation of composite thin-walled structures in which local loss of stability of the covering is acceptable in the area of post-critical loads.