PSI - Issue 60

Amardeepa KCS et al. / Procedia Structural Integrity 60 (2024) 60–74 Amardeepa KCS/ StructuralIntegrity Procedia 00 (2023) 000–000

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2 showed that the L angle is safe for carrying internal fuel pressure of 12.5 PSI. The study also showed that the strain and displacement values distribution in the standalone FE model are comparable with that of the global wing box model within the range of (8-10) % and (4-5) % difference, respectively. Keywords: Bond Energy Method; Composite wing; Failure prediction; Fuel pressure; L angle; Wing box; 1. INTRODUCTION One of the critical challenges in the structural design of composite aerospace structures is the design of components subjected to out-of-plane loading. The load must often be transferred in an out-of-plane direction when subjected to internal fuel pressure in the fuel tank region. When designing a composite wing structure, it is crucial to carefully design L shaped joints or L angles carrying out-of-plane loading to ensure structural integrity and safety. When subjected to out-of-plane loading, the composite laminate's performance is dominated mainly by the strength and stiffness of the resin material, the weakest phase in the laminate. The stacking sequence of the composite laminate influences the stress distribution in the resin region. The strength and stiffness properties of the resin region govern the ultimate load-carrying capacity of the L joint. Generally, the types of failure detected in composite joints are fiber breakage, microcracking, and delamination when subjected to in-plane loading. The most common source of failure of composite laminates/joints is delamination caused by the separation of composite layers, constraining cracks to grow in resin-rich regions between the adjacent plies, described as an interface crack. Unbalanced properties of reinforcement, matrix, and radial geometry of the curved part create weakness in a through-the-thickness direction that leads to delamination failure. Composite joints may pose a severe risk if the load transfer and damage mechanisms are improperly analyzed or understood. The composite structure experiences various types of stress, including tension, compression, and shear. These joints exhibit complex stress distribution and failure modes. A better understanding of material properties and superior design/analysis can overcome these unexpected failures that occur in inaccessible and invisible regions. Thus, finite element analysis is carried out on composite wings using a new design /validated FE analysis procedure. Figure 1 (a)&(b) shows the composite wing on the assembly jig and stringer co-cured with the bottom skin in the previous/originally designed composite wing of an aircraft, wherein the ribs are co-cured to the bottom skin and fastened onto the top skin through mechanical countersunk fasteners as shown in Figure 2 (a)&(b), however, in the modified /new design of the composite wing, the interspar box is modified by separating the IS ribs from the bottom skin and both front and rear spars and mechanically fastened to the top and bottom skin as shown in Figure 2 (c). © 2024 The Authors. Published by Elsevier B.V. This is an open access article under the CC BY-NC-ND license (https://creativecommons.org/licenses/by-nc-nd/4.0) Peer-review under responsibility of the ICONS 2023 Organizers Amardeepa KCS/ StructuralIntegrity Procedia 00 (2023) 000–000 showed that the L angle is safe for carrying internal fuel pressure of 12.5 PSI. The study also showed that the strain and displacement values distribution in the standalone FE model are comparable with that of the global wing box model within the range of (8-10) % and (4-5) % difference, respectively. Keywords: Bond Energy Method; Composite wing; Failure prediction; Fuel pressure; L angle; Wing box; 1. INTRODUCTION One of the critical challenges in the structural design of composite aerospace structures is the design of components subjected to out-of-plane loading. The load must often be transferred in an out-of-plane direction when subjected to internal fuel pressure in the fuel tank region. When designing a composite wing structure, it is crucial to carefully design L shaped joints or L angles carrying out-of-plane loading to ensure structural integrity and safety. When subjected to out-of-plane loading, the composite laminate's performance is dominated mainly by the strength and stiffness of the resin material, the weakest phase in the laminate. The stacking sequence of the composite laminate influences the stress distribution in the resin region. The strength and stiffness properties of the resin region govern the ultimate load-carrying capacity of the L joint. Generally, the types of failure detected in composite joints are fiber breakage, microcracking, and delamination when subjected to in-plane loading. The most common source of failure of composite laminates/joints is delamination caused by the separation of composite layers, constraining cracks to grow in resin-rich regions between the adjacent plies, described as an interface crack. Unbalanced properties of reinforcement, matrix, and radial geometry of the curved part create weakness in a through-the-thickness direction that leads to delamination failure. Composite joints may pose a severe risk if the load transfer and damage mechanisms are improperly analyzed or understood. The composite structure experiences various types of stress, including tension, compression, and shear. These joints exhibit complex stress distribution and failure modes. A better understanding of material properties and superior design/analysis can overcome these unexpected failures that occur in inaccessible and invisible regions. Thus, finite element analysis is carried out on composite wings using a new design /validated FE analysis procedure. Figure 1 (a)&(b) shows the composite wing on the assembly jig and stringer co-cured with the bottom skin in the previous/originally designed composite wing of an aircraft, wherein the ribs are co-cured to the bottom skin and fastened onto the top skin through mechanical countersunk fasteners as shown in Figure 2 (a)&(b), however, in the modified /new design of the composite wing, the interspar box is modified by separating the IS ribs from the bottom skin and both front and rear spars and mechanically fastened to the top and bottom skin as shown in Figure 2 (c).

(a) Composite wing on the assembly jig

(b)Stingers co-cured with the bottom skin

Fig. 1. Composite Wing

(a) Composite wing on the assembly jig

(b)Stingers co-cured with the bottom skin

Fig. 1. Composite Wing

(a) Typical interspar rib

(b) Rib co-cured to the bottom skin

(c ) New design (IS rib mechanically fastened to the spars and skins)

Fig. 2 Interspar rib details

(a) Typical interspar rib

(b) Rib co-cured to the bottom skin

(c ) New design (IS rib mechanically fastened to the spars and skins)

Fig. 2 Interspar rib details