PSI - Issue 60

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Procedia Structural Integrity 60 (2024) 1–2

3 rd International Conference on Structural Integrity 2023 (ICONS 2023) 3 rd International Conference on Structural Integrity 2023 (ICONS 2023) 3 rd International Conference on Structural Integrity 2023 (ICONS 2023) 3 rd International Conference on Structural Integrity 2023 (ICONS 2023)

Editorial – ICONS 2023 Editorial – ICONS 2023 Editorial – ICONS 2023 Editorial – ICONS 2023

2452-3216 © 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 10.1016/j.prostr.2024.05.025 Structural Integrity is crucial from the viewpoint of safety, longevity and economics of structures and components deployed in aerospace, civil engineering, automotive, energy systems, and many more. This is a complex domain, involving design, materials selection, materials processing, component level assembly, in-service loading, periodical inspection, repair and overhaul of critical components to assure safety of operations over the intended life time with high degree of reliability and confidence. Understanding the behavior of structures and materials under different loading and environmental conditions to determine their resistance to the deformation and failure prediction, in terms of strength and life is the central theme of structural integrity assessment. Recognizing the above and in consideration of interplay between design, manufacturing and operation of safety critical systems, the ICONS conference series was initiated by Indira Gandhi Centre for Atomic Research (IGCAR), Kalpakkam and Society for Failure Analysis, Chennai Chapter in 2014; the second edition was held in 2018 at IIT-Madras. ICONS series conference has been providing an appropriate platform for the exchange of knowledge and ideas in these rich and diverse fields of research and development. The third edition of the conference, ICONS 2023 held at Four Points by Sheraton, Mamallapuram, Tamil Nadu during Aug 2023 attracted quality interaction between industry, academia and R&D institutions, with a very good response from several leading industries, academic and R&D institutions. Major research topics covered in the conference included Mechanical Behavior of Materials, Fatigue and Fracture Mechanics, High Strain Rate Testing, Small Specimens Test Methods, Fretting, Fracture Mechanics based Design, Computational Mechanics, Micromechanics of Structural Integrity is crucial from the viewpoint of safety, longevity and economics of structures and components deployed in aerospace, civil engineering, automotive, energy systems, and many more. This is a complex domain, involving design, materials selection, materials processing, component level assembly, in-service loading, periodical inspection, repair and overhaul of critical components to assure safety of operations over the intended life time with high degree of reliability and confidence. Understanding the behavior of structures and materials under different loading and environmental conditions to determine their resistance to the deformation and failure prediction, in terms of strength and life is the central theme of structural integrity assessment. Recognizing the above and in consideration of interplay between design, manufacturing and operation of safety critical systems, the ICONS conference series was initiated by Indira Gandhi Centre for Atomic Research (IGCAR), Kalpakkam and Society for Failure Analysis, Chennai Chapter in 2014; the second edition was held in 2018 at IIT-Madras. ICONS series conference has been providing an appropriate platform for the exchange of knowledge and ideas in these rich and diverse fields of research and development. The third edition of the conference, ICONS 2023 held at Four Points by Sheraton, Mamallapuram, Tamil Nadu during Aug 2023 attracted quality interaction between industry, academia and R&D institutions, with a very good response from several leading industries, academic and R&D institutions. Major research topics covered in the conference included Mechanical Behavior of Materials, Fatigue and Fracture Mechanics, High Strain Rate Testing, Small Specimens Test Methods, Fretting, Fracture Mechanics based Design, Computational Mechanics, Micromechanics of Structural Integrity is crucial from the viewpoint of safety, longevity and economics of structures and components deployed in aerospace, civil engineering, automotive, energy systems, and many more. This is a complex domain, involving design, materials selection, materials processing, component level assembly, in-service loading, periodical inspection, repair and overhaul of critical components to assure safety of operations over the intended life time with high degree of reliability and confidence. Understanding the behavior of structures and materials under different loading and environmental conditions to determine their resistance to the deformation and failure prediction, in terms of strength and life is the central theme of structural integrity assessment. Recognizing the above and in consideration of interplay between design, manufacturing and operation of safety critical systems, the ICONS conference series was initiated by Indira Gandhi Centre for Atomic Research (IGCAR), Kalpakkam and Society for Failure Analysis, Chennai Chapter in 2014; the second edition was held in 2018 at IIT-Madras. ICONS series conference has been providing an appropriate platform for the exchange of knowledge and ideas in these rich and diverse fields of research and development. The third edition of the conference, ICONS 2023 held at Four Points by Sheraton, Mamallapuram, Tamil Nadu during Aug 2023 attracted quality interaction between industry, academia and R&D institutions, with a very good response from several leading industries, academic and R&D institutions. Major research topics covered in the conference included Mechanical Behavior of Materials, Fatigue and Fracture Mechanics, High Strain Rate Testing, Small Specimens Test Methods, Fretting, Structural Integrity is crucial from the viewpoint of safety, longevity and economics of structures and components deployed in aerospace, civil engineering, automotive, energy systems, and many more. This is a complex domain, involving design, materials selection, materials processing, component level assembly, in-service loading, periodical inspection, repair and overhaul of critical components to assure safety of operations over the intended life time with high degree of reliability and confidence. Understanding the behavior of structures and materials under different loading and environmental conditions to determine their resistance to the deformation and failure prediction, in terms of strength and life is the central theme of structural integrity assessment. Recognizing the above and in consideration of interplay between design, manufacturing and operation of safety critical systems, the ICONS conference series was initiated by Indira Gandhi Centre for Atomic Research (IGCAR), Kalpakkam and Society for Failure Analysis, Chennai Chapter in 2014; the second edition was held in 2018 at IIT-Madras. ICONS series conference has been providing an appropriate platform for the exchange of knowledge and ideas in these rich and diverse fields of research and development. The third edition of the conference, ICONS 2023 held at Four Points by Sheraton, Mamallapuram, Tamil Nadu during Aug 2023 attracted quality interaction between industry, academia and R&D institutions, with a very good response from several leading industries, academic and R&D institutions. Major research topics covered in the conference included Mechanical Behavior of Materials, Fatigue and Fracture Mechanics, High Strain Rate Testing, Small Specimens Test Methods, Fretting, S A Krishnan 1* , A Nagesha 1,2 , Aniruddha Moitra 1,2 , M Vasudevan 1,2 , Raghu V Prakash 3 1 Indira Gandhi Centre for Atomic Research, Kalpakkam, Tamil Nadu, India – 603 102 2 Homi Bhabha National Institute, Mumbai, Maharashtra, India – 400 094 3 Indian Institute of Technology Madras, Chennai, India – 600 036 S A Krishnan 1* , A Nagesha 1,2 , Aniruddha Moitra 1,2 , M Vasudevan 1,2 , Raghu V Prakash 3 1 Indira Gandhi Centre for Atomic Research, Kalpakkam, Tamil Nadu, India – 603 102 2 Homi Bhabha National Institute, Mumbai, Maharashtra, India – 400 094 3 Indian Institute of Technology Madras, Chennai, India – 600 036 S A Krishnan 1* , A Nagesha 1,2 , Aniruddha Moitra 1,2 , M Vasudevan 1,2 , Raghu V Prakash 3 1 Indira Gandhi Centre for Atomic Research, Kalpakkam, Tamil Nadu, India – 603 102 2 Homi Bhabha National Institute, Mumbai, Maharashtra, India – 400 094 3 Indian Institute of Technology Madras, Chennai, India – 600 036 S A Krishnan 1* , A Nagesha 1,2 , Aniruddha Moitra 1,2 , M Vasudevan 1,2 , Raghu V Prakash 3 1 Indira Gandhi Centre for Atomic Research, Kalpakkam, Tamil Nadu, India – 603 102 2 Homi Bhabha National Institute, Mumbai, Maharashtra, India – 400 094 3 Indian Institute of Technology Madras, Chennai, India – 600 036 *Corresponding author. Tel.: +91 8778577893 E-mail address: sakrish@igcar.gov.in *Corresponding author. Tel.: +91 8778577893 E-mail address: sakrish@igcar.gov.in *Corresponding author. Tel.: +91 8778577893 E-mail address: sakrish@igcar.gov.in *Corresponding author. Tel.: +91 8778577893 E-mail address: sakrish@igcar.gov.in

2 for Atomic Research (IGCAR), Kalpakkam and Society for Failure Analysis, Chennai Chapter in 2014; the second edition was held in 2018 at IIT-Madras. ICONS series conference has been providing an appropriate platform for the exchange of knowledge and ideas in these rich and diverse fields of research and development. The third edition of the conference, ICONS 2023 held at Four Points by Sheraton, Mamallapuram, Tamil Nadu during Aug 2023 attracted quality interaction between industry, academia and R&D institutions, with a very good response from several leading industries, academic and R&D institutions. Major research topics covered in the conference included Mechanical Behavior of Materials, Fatigue and Fracture Mechanics, High Strain Rate Testing, Small Specimens Test Methods, Fretting, Fracture Mechanics based Design, Computational Mechanics, Micromechanics of Fracture and Damage, Steel and Concrete Structures, Composites, Structural S.A. Krishnan et al. / Procedia Structural Integrity 60 (2024) 1–2 Materials and Weldments, NDE and Structural Health Monitoring, Failure Analysis, Reliability and Structural Integrity Assessment, Fitness for Service and Remaining Life Assessment and Aerospace Applications. The ICONS 2023 was attended by nearly 250 delegates comprising of engineers, materials scientists, academicians, industry experts, plant managers and regulatory personnel. There were three plenary lectures, eleven Keynote lectures, nineteen invited lectures and eighty-eight contributory paper presentations apart from seventy poster presentations. Many industrial partners have supported the event and displayed their products during the conference. As a precursor to the ICONS 2023, two pre-conference workshops were organized: 1) Fatigue Life and Fracture Assessment of Steel and Concrete Structural Components at CSIR-SERC, Chennai, and 2) Design, Manufacture and Quality Assurance of Large Size Components at IGCAR, Kalpakkam. The hard efforts put in by the organizers have culminated in the success of ICONS 2023, making it a memorable one. The organizing team would like to place on record the support and guidance received from Dr. B. Venkataraman, Director, IGCAR and Dr. Divakar Ramachandran, Director, Metallurgy and Materials Group . Dr. Divakar’s personal involvement was instrumental in the timely release of this special issue. We thank the authors for contributing their novel and valuable research findings and the papers presented in this special issue have been selected through a rigorous peer review by experts in the relevant fields of specialization. We thank the rewiewers for their painstaking efforts in critically reviewing the manuscripts. Thanks are also due to Prof. Francesco Iacoviello, Editor, Procedia Structural Integrity, and Elsevier Publishers for agreeing to publish the accepted papers as a special issue. We also thank the Elsevier production team for their sincere efforts in the timely release of this special issue.

Available online at www.sciencedirect.com StructuralIntegrity Procedia 00 (2023) 000–000 Available online at www.sciencedirect.com ScienceDirect

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Procedia Structural Integrity 60 (2024) 60–74

Third International Conferences on Structural Integrity 2023 (ICONS 2023) Analytical Assessment of Composite L Angle Strength Under

Internal Fuel Pressure in Composite Wing Amardeepa KCS 1a* , Polagangu James 2b , Byji Varughese 3c 1 Principal Scientist, 2 Senior Principal Scientist, 3 Chief Scientist Advance Composites Division, CSIR, National Aerospace Laboratories, Bangalore, India .

Abstract The Interspar (IS) box stores the turbine fuel in the composite wing of an Aircraft. The filled fuel is under pressure that induces out-of-plane stresses on all the wing's structural members around the fuel tank's boundary. The wing is designed using carbon fiber reinforced plastic (CFRP) with the co-curing of the bottom skin and its stringers along with front and rear spars. A suitable number of cut-outs are made on the top and bottom sides of the IS ribs for the stringer to pass through them. The global finite element analysis showed that the stringer cut-out regions in the IS ribs are highly stressed, as they are subjected to out-of-plane load due to internal fuel pressure, which induces transverse tensile stress in the cut-out region. Resin is the weakest material in the composite laminate, but it has to resist the high stress component due to out-of-plane loads. The top and bottom flanges of the 'C' section IS ribs geometrically form an L angle section with a stringer cut-out made in it. At the L angle location, the IS ribs are critical from both the strength and stiffness point of view. Strength and stiffness at the interface of the fastener joint are essential factors in maintaining the composite wing's overall structural integrity under internal fuel pressure. The authors adopted the same standalone finite element (FE) modelling approach developed for the Bond Energy Method and carried out appropriate analytical studies to ascertain the safety and structural integrity of the composite wing under internal fuel pressure. This article presents the analytical study carried out on the standalone FE model of an L angle with the stringer cut-out introduced at its appropriate position. The static stress analysis is carried out by enforcing the global displacement on this standalone FE model and extracting three force components in CBAR beam elements representing the resin property in the transverse direction of the composite laminate. The stresses are calculated outside the analysis deck using Excel® spreadsheets. The maximum Normal Stress (σ n ) in the L angle flange is found to be 13.25 MPa, which is well within the resin or transverse tensile strength of the composite laminate 53 MPa, a nd the shear stress is found to be 11.10 MPa, which is less than the resin shear strength property of 23 MPa. Hence the study

* Corresponding author. fax: 91 80 25267352 E-mail address: amardeepa@nal.res.in

2452-3216© 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

2452-3216 © 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 10.1016/j.prostr.2024.05.031

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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

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2. LITERATURE REVIEW Lekhnitskii [1] introduced a beam theory-based calculation method to predict the transverse tension state and curved beam strength in an L-shaped beam with unidirectional composite plies under four-point bending. These loads were valid only for pure bending or edge loading because they cannot sustain the circumferential force without radial restraints. Baba [2] proposed a simplified post-processing approach to predict the folding-unfolding interlaminar stress components created at the symmetrically balanced curved composite laminate thickness under the combined action of shear forces, axial forces, and bending moments applied in the curvature plane. This approach is based on the extraction of forces and moments in the elements at the baseline curvature region. It provides a comparison between 3-D layered solid elements and layered shell elements. Thurnherr. C et al. [3] investigated the delamination failure generated in the curved laminates from applied bending moments, which induces tensile stresses in the radial direction using a higher-order beam model, which gives an idea about the mechanics of failure initiation for design purposes. In thin laminates, in-plane failure is initiated by the hoop stress (critical stress component). However, the transverse shear stress (critical stress component) initiates delamination cracks (de-bonding failure), which are predominant rather than in-plane failure, as the thickness increases in the laminates. As an extension of this theory, two clamped ends and one non-clamped end case were studied, making interlaminar shear stress and radial stress critical. J.S.Charrier et al. [4] investigated the out-of-plane tensile strength of an L shaped specimen with a constant radius-to thickness ratio (different thicknesses) and various carbon /epoxy laminates such as highly oriented, highly disoriented, quasi-isotropic, and unidirectional were used. At experimental failure load, maximum out-of-plane tensile stress is approximated by a 3D linear elastic finite model to predict the Curved Beam Strength (CBS). They concluded that the thick laminates were sufficient to avoid arms bending in four-point bending simultaneously; stacking sequence may not cause significant changes in observed strength. Hao et al. [5] investigated carbon-epoxy prepreg L-shaped specimens with 20, 40, and 60 plies. The delamination starts at once at various interfaces for the thinnest specimens, whereas for thicker specimens, delamination was started close by the center of the arc at the mid-thickness and then occurs in upper and lower interfaces. Polagangu James et al. [6] carried out a two-stage FE analysis for analytically determining the strength of composite co-cured T-joint against internal fuel pressure. In the first stage, an analytical study was carried out on a wing /wing box by applying an internal fuel pressure of 12.50 PSI to identify the high out of-plane stress region along the co-cured region of both bottom skin and spars. In the second stage, he created a 1D 2D FE model of a composite co-cured T-joint through an innovative, simplified approach developed based on a physical phenomenon inspired by nature named Bubbles in the Bermuda Triangle. The 1D elements represented the resin-bonded roving fragments filled in that triangle region precisely captured the stress distribution in that failure region. Polagangu James et al. [7] used a similar FE modelling approach and created adhesively bonded regions in various adhesively bonded composite (ABC) repaired joint configurations, explaining the structural failure mechanics. He also proposed a novel failure criterion using mechanical energy principles, named the Bond Energy Method. This modelling approach and the novel failure criterion precisely estimated the ultimate failure load of various third-party experiments. By using this approach, areas of bonded regions can be identified that are subjected to tensile, compression, and shear stresses, and the failure of bonded joints was explained through six modes of failures. Martin. R. H [8] studied delamination failure in a unidirectional curved composite laminate. In the curved region, at the location of the highest radial stress, delamination was supposed to be created, and this delamination caused the unstable failure of curved laminates. The interlaminar tension failure was predicted by strength-based failure criterion, and 2D FEA and a closed-form curved beam elasticity solution found excessive radial stress location. The present paper discusses details of the analytical study carried out on the standalone FE model of a composite L angle with a stringer cut-out. The structural integrity is analytically assessed when subjected to internal fuel pressure. The static stress analysis was carried out to understand the failure load of the L angle when subjected to out-of-plane loads through a novel FE modelling approach and failure criterion [6]. 3. FE ANALYSIS OF WING / WINGBOX The experimentally validated FE modelling approach and novel failure criteria [6,7] are the baseline methodology adopted in the present study for creating an FE model of an L angle joint and estimating the ultimate failure load or margin of safety values. Static stress and buckling analysis are carried out for critical air load cases generated as per

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FAR 23 standards. In addition, critical fuel pressure distribution and other inertia loads are considered for analysing the wing structure. Figure 3 shows the global wing analysis results with the distribution of Yamada Sun's failure index value for the internal fuel pressure of 12.5 PSI. The results show that an L angle region in IS ribs # 7,8,9 and 10 is subjected to high stress and failure index value around the stringer cut-out region. The wing design cannot be cleared with this high failure index (safe limiting value is <1.00) unless it is proven separately that this L angle region in IS ribs is experimentally or analytically safe. Before taking up experimental studies, demonstrating its safety through analysis is necessary as per certification requirements and regulations. Therefore, the present analytical study considered a section from the global wing model extended between stations #6 to 10 that represents the full global wing features in the wing box structure, as shown in Figure 4. The dimensional details of the wing box measure 1150 mm in length, 1148 mm in width at the root, and 950 mm in width at the tip. The depth is 340 mm at the root and 290 mm at the tip. This composite wing box is part of a global composite wing FE model considered for this study purpose. The study simplifies the bigger size problem into a smaller one by considering the typical region of the wing box encircled within the dotted line shown in Figure 1(a) and (b). The pre-and post-processing of the finite element model of the wing box is carried out using HyperMesh® V19 with MSC Nastran interface features, and various analysis solutions are obtained through MSC/NASATRAN® R2016 Solver. Figure 5(a)&(b) shows the load and boundary condition applied in the wing box to simulate balloon constraints [9,10], allowing the skins and spars to expand freely under applied internal fuel pressure without offering any nature of physical constraints seen in a blown balloon or football.

Fig. 3. High-stress regions at the rib location of the global wing model

Fig. 4. FE model of the wing box Station (6-10)

(b)

(a)

Fig. 5. Loads and Boundary conditions of the wing box

Figure 6 shows the deformed shape of the wing box when subjected to internal fuel pressure with balloon constraints. Figure 7 shows the net displacement comparison of the wing and wing box, which is found within a 4-5 % difference at the maximum displacement region. The failure index value distribution or high stress region in the L angle of typical IS ribs of the standalone wing box is shown in Figure 8. It also shows the same high stresses at the stringer cut-out region in the rib of the wing box, similar to that of the global wing. Therefore, it is found that the L angle flange of IS ribs is the most critical in the fuel tank region. Hence the standalone FE analysis is carried out on the L angle to assess and ensure the strength and safety of the composite wing structure for maintaining the overall structural integrity.

4. FINITE ELEMENT ANALYSIS OF STANDALONE L ANGLE Figure 3 and Figure 8 show the global wing and wing box analysis that an L angle region in rib # 7,8,9,10 is subjected

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to high stress and failure index value around the stringer cut-out region. In the next level of analytical study, the L angle from the same region is considered for continuing the standalone analytical study marked in the dotted rib region, as shown in Figure 9(a) at stringer number 8. Figure 9(b) shows the FE model of an L angle flange created by dividing the mesh into shell elements using 2D CQUAD4 and CTRIA3 elements, assigned with PCOMP property and defined by MAT8 material card of NASTRAN®, which have six degrees of freedom in displacements and rotations.

Fig. 6. Displacements of the wing box

Fig. 7. Net displacements of the wing box and wing

Fig. 8. High stress and high failure index value around the stringer cut-out region in the rib of wing/wing box

The dimensional details of the L angle considered in the present standalone study are shown in Figure 9(c). The composite material property is defined with the experimentally measured values E11=130 Gpa, E22=10 Gpa, υ=0.35, and G=5 Gpa is assigned to 2D CQUAD4 elements with different ply orientations [+45/-45/0/+45/-45/90] s as defined in the wing and the wing box. The composite laminate is divided into four separate layers as it contains the four orientations of plies, and these layers are offset by 0.45 mm. Each layer contains a nearly equal number of composite layers in sequence, counting from one of the faces of the laminate. One limitation of carrying out a two-dimensional FE analysis is that it cannot give the stress components in the third dimension. In experimentally validated analytical studies carried out by authors [6,7], the third directional stresses were captured by modelling 1D elements across the thickness of the composite laminate. A similar approach is also adopted in this present study to capture the third directional stress components acting normal to the outer/inner surfaces of the composite laminate at all points. Various stress components like shear and normal stresses acting in this area govern the failure of the L angle.

(b) Parts of the L angle with stringer cut-out

(c ) Dimension of the L angle

(a) A typical IS rib

Fig. 9. Detailed features of the Interspar rib

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The resin or the adhesive layer is modelled using 1D CBAR beam elements with the PBAR property card of NASTRAN. And the 1D CBAR beam element is assigned with isotropic material properties of E=10 GPa, G=5 GPa, and υ=0.35, which is the same as the resin material. The 1D resin or adhesive element is connected with 2D CQUAD4 shell elements through each layer, as shown in Figure 10(a) to (d). The cross-sectional area of these 1D CBAR elements is defined so that the cross-sectional area of 1D CBAR elements and 2D CQUAD4 /CTRIA3 elements connected to these elements is almost the same to obey the mass conservation principle [7]. This is achieved by defining half of the density value to the composite layers and the other half to these CBAR elements. The 2D elements in the curve region are divided into minutes of a standard analogous clock dial (ACD) [6] to represent the number of 1D elements possible in a 1/4 th area of ACD [6], as shown in Figure 10(e).

(a) and (b) 2D CQUD4 elements connected with 1D CBAR element, (c) and (d) 3D view of 2D and 1D elements

(e) Standard analogous clock dial representing the minutes at the curved region of an L angle

Fig. 10. 3D view of the L angle FE model; curved region divided into minutes of a clock

A typical illustration is shown in Figure 11(a) to calculate the cross-sectional area of the 1D CBAR elements that can be defined by choosing either a rectangular or circular cross-section shape conveniently. However, in this study, a rectangular cross-sectional shape is considered. The composite laminate of 12 plies (ply thickness is 0.15mm) is divided into 4 shell layers, each consisting of 3 composite plies with the sequence of [ (+45, -45, 0); (+45, -45, 90); (90, -45, +45); (0, -45, +45)]. A circular patterned 2D elements around the stringer cut-out region is modelled for the ease of capturing radial stress distribution precisely as shown in Figure 11(b). More than 10000 1D CBAR elements are modelled in the standalone FE model of the L angle and numbered in chronological order to quickly identify their precise location across the whole face and thickness direction of the composite laminate.

(a)

(b) Fig. 11. A typical illustration for calculating the 1D CBAR Element cross-section; Force components of the bar elements and zones

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This numbering system helped post-process the results with proper output data control. The entire surface area of the L angle is divided into various zones to identify the location in web, radius, and flange regions as shown in Figure 11(b). Zone 1 represents the flat laminate in the flange region, Zone 2 is the fillet or radius region, Zone 3 is the web region with regular CQUAD elements, and Zone 4 CQUAD elements aligned radially around the stringer cut-out region. The shear force components of the 1D CBAR elements are aligned in the plane of composite laminate either in the flat or fillet region, while the axial force component is placed normal to the composite laminate surface in all zones, as shown in Figure 11(b). The direction of these force components is symmetrical about its axis of symmetry, as shown in the same Figure. The 1D elements are assigned resin material properties, and 2D elements are assigned CFRP material properties for analytical studies. 4.1. Loads and Boundary Condition The global displacement values are extracted for the set of fasteners connecting the top skin to the IS rib of the wing box FE model, as shown in Figure 9. A maximum displacement value is chosen from the extracted values and applied as enforce displacements D1 and D2 (load) in the standalone FE model of L angle, as shown in Figure 12 where D1 = 0.6 mm and D2 = 0.71 mm are the displacement values extracted at the fasteners located on either side of the stringer cut-out as shown in Figure 13. The boundary conditions are applied as an enforced displacement at an appropriate location, as shown in Figure 14 to simulate the global behavior of the wing box and wing. All the 3 translations' degree of freedom (∆x, ∆y, and ∆z) is constrained using an SPC (single point constraints) load card, and 6 degrees of freedom (∆x, ∆y, ∆z, θx, θy, and θz) are constrained using an SPCD card; 3 translation and 3 rotations are applied on the local model in ordered to solve for the static analysis. The same FE analysis is carried out for different displacement factors to know the safety parameters and overall structural integrity, as shown in Table - 1.

Fig. 14. Boundaries of an L angle from the global wing box

Fig. 13. exploded view of a typical fastener location of L angle with maximum displacement observed

Fig. 12. Loads and boundary conditions applied in the FE model of L angle

Table 1. Displacements of fasteners for the L angle to check failure for different cases

No. Fasteners/ Node No.

1

2

3

4

5

6

7

8

9

[1]

[2]

[3]

[4]

[5]

[6]

[7]

[8]

[9]

[10]

6027185/ Displacement 1 (Z-axis)

0.66

1.4

2.1

2.2

2.25

2.3

2.4

2.5

2.6

6027184 / Displacement 2 (Z-axis)

0.71

1.5

2.3

2.4

2.45

2.5

2.6

2.7

2.8

4.2. Challenges Understanding the failure of composite structures even after carrying out mechanical tests becomes cumbersome due to the non-availability of appropriate analytical and numerical methods, modelling approaches, and suitable failure criteria. This work presents the numerical methods and FE modelling approach adopted for understanding the failure load, and the location of failure of an L angle flange of composite IS rib that experiences out-of-plane loading due to

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the application of internal fuel pressure. This FE modelling approach was validated by experiments using the test results of composite co-cured T-joints and adhesively bonded composite joints. Assessing the failure load of this L angle flange in a global wing or wing box requires an enormous amount of composite materials for fabrication, workforce, and time. With or without expensing all these resources, it still poses technical challenges to the structural designer to understand the failure load and location of failure of the wing or wing box when subjected to internal fuel pressure. The present modelling approach and novel failure criterion helped overcome such technical challenges in understanding the stress distribution, failure load, and failure location. Initially, extensive time has been spent on creating the FE model. The sectional properties of each 1D CBAR element became different from other elements due to dimensional changes in each zone. Thereby the size of the FE model became large. Subsequently, the mesh pattern is revised to have the same sectional properties of 1D elements over a large area in the web and flange regions. The radial element pattern shown in Figure 11(b) helped to have a minimum number of sectional properties for 1D elements, which drastically reduced the size of the FE model, as shown in Figure 10. The other challenge is enforcing boundary conditions in the standalone FE model of the L angle flange. After many iterations, it is decided to apply an enforced global displacement on the standalone FE model. Estimating failure load with the enforced displacements is possible only through analytical study; an experimental test is impossible to carry out with the enforced displacement. Conducting the experimental study until the failure of the L angle does not validate the design. The failure of the L angle with differently randomly applied displacements does not truly represent the actual failure load in the global wing structure. Therefore, the present study gained importance over the mechanical tests as the stress and strain distributions in the web of the L angle around the cut-out region are closely correlated with that of values obtained from the global wing/wing box. 5. Analysis The linear static analysis provides three force components of 1D CBAR elements in the output data file F06, which is the crucial input data for computing the various stress values using Eq. (1) to (3). The axial force and shear force components are acting along the plane 1 and 2 of 1D CBAR elements as shown in Figure 11. The force values are exported to an Excel spreadsheet for carrying out simple engineering and mathematical calculations/operations outside the FE analysis domain. The failure criterion given in Sec. 5.1 is applied for estimating the failure load and its location identification between layers (Figure 10(b)) of composite laminate.

5.1. Failure Criteria used for checking the strength of L angle against out-of-plane stresses at the radius region:

σ � = � �� + �� + �� +� � − 2 � � +� � − 2 � � + � � − 2 � � τ � =�P � � +P � �

Eq-(1)

Eq-(2)

τ = τ r /a n ≤ tt No failure n > tt Failure

Eq-(3)

τ ≤ ss No failure τ > ss Failure

Where, σ X = stress due to Axial force (along the element axis)

σ Y = stress due to P1 (along the tangent of the curve)

σ Z = Stress due to P2 (along the span) A F = Axial force of a CBAR element

σ n = Normal stress

P 1 = Tangential force component of a CBAR element tt = Transverse tensile strength of composite laminate, which is equivalent to resin tensile strength ss = Shear strength of composite laminate, which is equivalent to resin shear strength τ = Shear stress τ r = Resultant shear

P 2 = Longitudinal force component of a CBAR element

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5.2. Yamada-Sun's failure criterion used for checking the safety of composite plies against in-plane stresses:

2

2

   

   

S      L



   

1.0

Ys =

Eq-(3)



12

S

Where, σ 11 = Normal stress in lamina along fiber direction τ 12 = Shear stress in lamina S L = Allowable stress in lamina along fiber direction (X t = 520 N/mm

2 and X

c = 506 N/mm

2 )

S = Allowable in-plane shear stress in the lamina (s = 56.3 N/mm 2 ) The Axial Force (AF) acting along the axis of the 1D CBAR element and the transverse forces along Plane 1 shear and Plane 2 shear are resolved for the applied enforced displacement, as shown in Figure 15. As discussed in Section 4, the 1D CBAR elements are numbered in chronological order for one complete row of the FE model starting from 1 up to 1D element number 131 of 1D CBAR elements as shown in Figure 16 (a) for its easy identification and precise location. The standalone FE model is a symmetry model; hence a typical numbering pattern is illustrated for envisaging the numbering pattern as shown in Figure 16 (b)&(c) at zone 2. Similarly, the numbering pattern continues for 25 rows for layers (1-2), (2-3), and (3-4). As the data is enormous to discuss, Figures (17), (18)& (19) show a typical distribution of the forces along Axial Force (AF), Plane 1 shear, and Plane 2 shear for the resolved force components of points 1 to 4 for one complete row of elements based on the axial and shear force diagrams shown in Figure 15.

Fig 15 Typical ID CBAR element axial shear force components resolved to show the force equilibrium.

(a) Numbering of ID CBAR element for 1 Row of elements between layers 1-2

(b) zone 2

(c) 3D view of 1D elements at zone 2

Fig 16 (a) Numbering of ID CBAR element for 1 Row of elements between layers 1-2, (b) zone 2, (c) 3D view of 1D elements at zone 2

6. Results and Discussions Figure 20 shows the maximum displacement of the L angle is 1.23 mm along the (Z-axis) ), at the free edge, but at

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the point of application of enforced displacement (D1 and D2), values are the same as that of applied enforced displacement called in Table 1. The variation in the maximum displacement values was found to be only 4-5% when compared with that of a global wing (component), wing box (sub-component), and standalone L angle. The 2D layers are connected by 1D CBAR elements at zone 2 Figure 10(e) & Figure 11(b) with a rectangular cross-section property adopted for modelling by choice.

Fig. 17. Axial force for 1 st row of elements

Fig. 18. Shear force in Plane 1 for 1 st row of elements

Fig. 19. Shear force in Plane 2 for 1 st row of elements

Fig. 20. Max displacement along z

A 3D view of the 1D CBAR elements is shown in Figure 21(a)&(b) for zones 2,3 and 4. The Normal Stress (σ n ) at zone 2,3, and 4 are plotted by choosing the CBAR elements in the clockwise direction, as shown in Figure 21(a). There are 16 elements along the column and 25 elements in the row, and 4 points (P1, P2, P3, and P4) at each layer (layer 1 to Layer 4) shown in Figure 10(a) across the thickness direction in the curved region. The points are chosen along the column (22-41) in row 1 (153-172) in row 2, and so on up to row 25, as shown in Figure 16(b) and Figure 21(a). The normal stress distribution is plotted in Figure 22 at zone 2 for all 25 rows. The out-of-plane stress value that induces the tensile stress between each layer is 13.25 MPa at zone 2 and 20.90 MPa at zone 3 and 4, which is less than 53 MPa (the resin tensile strength or transverse tensile strength property of the composite laminate). The resultant shear stress due to Plane 1& 2 is 11.10 MPa, and axial stress for zone 2 is 7.92 MPa, plotted for one case (max displacement 0.7), as shown in Figure 23 & 24, respectively.

(a) 3D view of CBAR elements at zone 2

(b) 3D view of CBAR elements at Zone 3 and 4

Fig. 21. 3D view of 1D CBAR elements

11

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Fig. 22. Normal stress at Zone 2 for all 25 rows; (NS= σ n )

Fig. 23. Resultant Shear Stress due to (τP1 &τP2 ) @ zone 2 for 25 rows

Fig. 24. Axial stress at zone 2 for 25 rows

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19.300 mm

21.439mm

17.161mm

27.098 mm

27.098 mm

23 mm

(a) stress location between layers 1-2 (15 th row)

(b) stress location between layers 1-2 (16 th row)

(c) stress location between layers 1-2 (17 th row)

Fig. 25. Stress location at zone 2

Table 2. Summary of the enforced displacement values and induced stress values, along with their location

Zone 2

Zone 3 and 4

Max displacement applied

Failure spot Normal Stress

Failure spot Resultant shear stress for Plane 1 &2

Failure spot

Normal Stress

Failure spot

Failure spot Axial stress

Resultant shear stress for Plane 1 &2

Failure spot Axial stress

[1] 0.7 1.5 2.3 2.4 2.5 2.6 2.7 2.8

[2]

[3]

[4]

[5]

[6]

[7]

[8]

[9]

[10]

[11]

[12]

[13]

11.10 23.56 36.11 37.67 38.46 39.24 40.80 42.36 43.94

Fig. 25 (a) 7.92 Fig. 25 (a) 17.9 Fig. 25 (a) 28.0

Fig. 25 (b) 13.25 Fig. 25 (a) 16.74 Fig. 25 (c) 28.20 Fig. 25 (a) 19.87 Fig. 25 (c) 43.30 Fig. 25 (a) 31.02

Fig. 26 (a) 8.19

Fig. 26 (e) 20.90 Fig. 26 (a)

Fig. 26 (b) 18.99 Fig. 26 (e) 31.86 Fig. 26 (d) Fig. 26 (c) 29.91 Fig. 26 (e) 49.78 Fig. 26 (e) Fig. 26 (c) 31.28 Fig. 26 (e) 52.02 Fig. 26 (e) Fig. 26 (c) 31.96 Fig. 26 (e) 53.13 Fig. 26 (e) Fig. 26 (c) 32.63 Fig. 26 (e) 54.25 Fig. 26 (e) Fig. 26 (c) 33.99 Fig. 26 (e) 56.47 Fig. 26 (e) Fig. 26 (c) 35.35 Fig. 26 (e) 58.70 Fig. 26 (e) Fig. 26 (c) 36.73 Fig. 26 (e) 60.96 Fig. 26 (e)

Fig. 25 (a) 29.27 Fig. 25 (c) 45.19 Fig. 25 (a) 32.33 Fig. 25 (a) 29.89 Fig. 25 (c) 46.13 Fig. 25 (a) 32.98 Fig. 25 (a) 30.52 Fig. 25 (c) 47.07 Fig. 25 (a) 33.62 Fig. 25 (a) 31.78 Fig. 25 (c) 48.95 Fig. 25 (a) 34.92 Fig. 25 (a) 33.04 Fig. 25 (c) 50.83 Fig. 25 (a) 36.22

2.45

Fig. 25 (a) 34.3

Fig. 25 (c) 52.73 Fig. 25 (a) 37.55

Note: Stress values given in [2] are acting in Fig [3] and so on [4] [5]; [6] [7];

Figure 25 & 26 shows where the layers start to fail in the curved region of zones 2,3 and 4respectively. The results are tabulated in Table 2 for all 9 classes of displacement values given in Table 1 for quick reference. Figure 27 shows the comparison of strain values of the standalone L angle with that of the wing box, which is within 8-10% of the difference range. Figure 28 shows the Failure Index values extracted for each sub-laminate that consists of different layers and orientations as per the modelling approach adopted in this study. Figure 29 shows the variation of the maximum resultant shear stress, normal stress, and axial stress values computed by applying various displacement values on a standalone L angle. Figure 29 shows two constant horizontal lines at different levels, indicating the line of allowable shear strength value of 23 MPa, and allowable transverse tensile strength value of 53 MPa. These allowable strength values remain the same for all cases considered. It is inferred from Figure 29 that the L angle fails in shear between layers (1-2) when the applied internal fuel pressure increases from 12.5 PSI to 26.78 PSI (12.50 PSI x1.50 mm / 0.70 mm linearly interpolated between enforced displacements and applied internal fuel pressure), and the same L angle fails in tension across the composite layers (1-2) as it reaches its allowable transverse tensile strength value at 42 PSI arrived at by the similar linear interpolation at a displacement of 2.4 mm; however, the first ultimate failure load of the L angle is estimated to have occurred at an internal fuel pressure of 26.78 PSI which is more than