PSI - Issue 24

Alessandro Castriota et al. / Procedia Structural Integrity 24 (2019) 279–288 A. Castriota et al. / Structural Integrity Procedia 00 (2019) 000 – 000

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experiencing great growth in recent years, due to the high ratio between mechanical strength and weight. However, this fact introduces new questions on the behavior that structures in composite material could have in the presence of damage. In particular, given that the use of composite materials involves thin structures of an aircraft such as the fuselage and wings, it is very important to study how the buckling behavior varies in the presence of damage (Mallick (1999)). Since the thin structures are subject to the phenomenon of buckling due to the in-service compression loads, they are commonly reinforced with stringers and ribs in order to stiffen them and to sustain higher loads before structural instability occurs. The possibility to withstand compressive loads even after going into instability (post-buckling behavior) is complicated by the local behaviour of the composite material at higher level of deformation, which involves different kind of damages, such as debonding, delamination, matrix crack propagation (Megson (1993)). The study and analysis of the buckling behavior of aeronautical structures is a widespread practice, which often undertaken during the design of a new aircraft or during the service to analyze particular conditions For example, Zimmermann et al. (2006) evaluated the buckling and post-buckling of curved panels stiffened with stringers. In other cases, it is important to evaluate the buckling behavior of the component in the presence of some forms of damage already present on the panel before it is loaded. Sepe et al. (2016) investigated the influence of different areas with buckling and post-buckling of a panel stiffened with omega stringers. In 2016, Feng and al. studied the effect of the position of impact on the buckling and the post-buckling behavior of the panel. Instead, Wang et al. (2015) investigated the numerical and experimental behavior at post buckling of panels reinforced with impact damage. In other works, such as those done by Wilckens et al. (2013) and by Boni et al. (2012), they analyzed the buckling behavior of composite panels but in the absence of damage. Finally, Riccio et al. (2017) and Sellitto et al. (2019) analyzed the progress of damage due to other causes. In this work, a CFRP panel with longitudinal and transverse reinforcements has been studied to evaluate how the presence of a relevant damage might influence the buckling behavior. At this aim, the results of two numerical models of the intact and of the damaged component were compared. The damaged panel is characterized by a cut in the center that interrupts the continuity of the central stringer. The numerical results were validated using the data of an experimental test carried out on the damaged panel, extending this validation also to the case of the intact panel.

Nomenclature CFRP Carbon Fibre Reinforced Polymer P Reference load

2. Geometry and material

The component is a large CFRP panel representative of a structural part of an aircraft. The panel is constituted by a CFRP skin longitudinally reinforced with 5 co-bonded stringers glued to it and transversely with 2 ribs, fixed to the panel on the same side of the stringers with rivets in Ti6Al4V to simulate a rib pitch. The panel has dimensions of 762x914.4 mm and the ends were inserted in two plates used for applying the compression load. The damaged panel presents a transversal cut having dimensions of 152.4x6.35 mm in the center of the panel that eliminated all the load capacity of the central stringer and of the surrounding skin (Fig. 1). Skin, stringers and ribs were realized in laminated CFRP with different stacking sequence, as detailed in the following:  Skin: [+45/90/-45/-45/+45/90/0/-45/45/0]s plus 4 tabs ([0/0]s; the first 2 tabs are between the initial plies +45 and 90, the other 2 follow the symmetry;  Stringers: cap ([+45/90/0/0/-45/0/0/-45/0/0/90/+45]s); web ([+ 45/90/0/0/-45/0/0/-45/0/0/90/+45]);  Ribs: [(±45) (0/90) (±45) (0/90) (±45) (0/90)]s.