Journal articles: 'Stiffened composite panel'

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Shao, Qing, Yu Ting He, Teng Zhang, Hai Wei Zhang, and Qing Shan Kang. "Simulation of Compress Buckling Performance of Composite Stiffened Panel." Applied Mechanics and Materials 184-185 (June 2012): 1189–93. http://dx.doi.org/10.4028/www.scientific.net/amm.184-185.1189.

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Finite element method is applied to analyze the buckling performance of composite stiffened panel. Compress buckling critical loads of six types panels with T or Z-section stiffeners are calculated by FEM. The emulational calculation results show that with same cross section area, critical buckling load of panel with T-section stiffeners increases with the reduction of stiffener pitch and the increase of stiffener numbers, while the buckling load of panel with Z-section stiffeners increases to a certain level and then keep almost changeless. To T-section stiffener panels, the relation between thickness of skin and buckling load is approximately quadratic trinominal. Conclusions obtained can offer a referenced measure for the optimization design and engineering application of the structure.

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Moorhouse, Anna, Micah Shepherd, and Benjamin S. Beck. "Incorporation of acoustic black hole stiffeners into composite airframes for reduction of noise radiation." Journal of the Acoustical Society of America 151, no. 4 (April 2022): A129—A130. http://dx.doi.org/10.1121/10.0010875.

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Stiffened composite panels are commonly used in aerospace structures, because they are lightweight, while maintaining a high load-bearing ability. However, their high stiffness-to-mass ratio makes them efficient noise radiators. In rotorcraft cabins made with composite panels, for example, the internal noise levels can be quite high such that pilot and passenger communication and comfort are disrupted. This has led to a need for innovative noise reduction strategies for composite rotorcraft panels. A specialized stiffener, which incorporates the acoustic black hole (ABH) effect into the cross section, is proposed to improve the damping of stiffened composite panels. By incorporating the damping concept into the stiffeners, the panel’s radiated noise can be reduced while maintaining the weight advantages and panel strength. To determine the advantages and trade-offs of this concept, numerical models have been developed and incorporated into an optimization scheme. Computational studies reveal promising results from the optimized ABH stiffeners as compared to a baseline panel with traditional stiffeners.

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Zhao, Chuang, Zhi Dong Guan, and Xia Guo. "Compression Performance of Repaired Composite Stiffened Panels: Bolted Repair and Scarfed Bonded Repair." Materials Science Forum 813 (March 2015): 152–60. http://dx.doi.org/10.4028/www.scientific.net/msf.813.152.

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In this paper, compression performance of composite stiffened panels after bolted repair and scarfed bonded repair is presented. Several tests have been performed on the four kinds of stiffened panels. All the experiments show the influences of different damage actions, different repair methods for the mechanical properties of stiffened panels are different. Furthermore, numerical finite element models have been developed and solved using the software ABAQUS. In addition, stiffener-skin debonding was considered in the finite element models in order to ensure a better consistence with the fact. The comparison of simulation results and experiment shows that the finite element modeling method is effective. The result suggests that the failure load and the stiffness of the repaired stiffened panel are recovered close to the virgin ones’.

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Liu, Jun, Bao Zong Huang, and Ge Lan Li. "Secondary Buckling and Progressive Failure Analysis of Tailored Composite Panels in Thin-Walled Box-Shape Composite Beams under Bending and Torsion." Advanced Materials Research 41-42 (April 2008): 445–48. http://dx.doi.org/10.4028/www.scientific.net/amr.41-42.445.

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The postbuckling behavior and progressive failure of thin-walled box-shape composite beams (BSCBs) have been studied using a simplified composite panel model. It is shown that the carrying capacity of BSCBs under bending and torsion mainly depends on the postbuckling and progressive failure of panels. It is often necessary to identify and follow secondary buckling of tailored panels for correct estimation of carrying capacity. The failure process of stiffened composite panels is simulated in following path. The criterion of local failure in flange of stiffeners is applied to predict the ultimate failure of stiffened composite panels and thin-walled BSCBs. The comparison of the modeling and test shows that the prediction of BSCBs failure is satisfactory.

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Cherniaev, Aleksandr. "A new laminates encoding scheme for the genetic algorithm-based optimization of stiffened composite panels." Engineering Computations 31, no. 1 (February 25, 2014): 33–47. http://dx.doi.org/10.1108/ec-08-2011-0089.

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Purpose – The genetic algorithm (GA) technique is widely used for the optimization of stiffened composite panels. It is based on sequential execution of a number of specific operators, including the encoding of particular design variables. For instance, in the case of a stiffened composite panel, the design variables that need to be encoded are: the number of plies and their stacking sequences in the panel skin and stiffeners. This paper aims to present a novel, implicit, heuristic approach for encoding composite laminates and, through its use, demonstrates an improvement in the optimization process. Design/methodology/approach – The stiffened panel optimization has been formulated as a constrained discrete minimum-weight design problem. GAs, which use both new encoding schemes and those previously described in the literature, have been used to find near-optimal solutions to the formulated problem. The influence of the new encoding scheme on the searching capabilities of the GA has been investigated using comparative analysis of the optimization results. Findings – The new encoding scheme allows the definition of stacking sequences in composites using shorter symbolic representations as compared with standard encoding operators and, as a result of this, a reduction in the problem design space. According to numerical experiments performed in this work, this feature enables GA to obtain near-optimal designs using smaller population sizes than those required if standard encoding schemes are used. Originality/value – The approach to encoding laminates presented in this paper is based on the original heuristics. In the context of GA-based optimization of stiffened composite panels, the use of the new approach rather than the standard encoding technique can lead to a significant reduction in computational time employed.

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Sanchez-Carmona, Alejandro, and Cristina Cuerno-Rejado. "Composite stiffened panel sizing for conceptual tail design." Aircraft Engineering and Aerospace Technology 90, no. 8 (November 5, 2018): 1272–81. http://dx.doi.org/10.1108/aeat-05-2017-0129.

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Purpose A conceptual design method for composite material stiffened panels used in aircraft tail structures and unmanned aircraft has been developed to bear compression and shear loads. Design/methodology/approach The method is based on classical laminated theory to fulfil the requirement of building a fast design tool, necessary for this preliminary stage. The design criterion is local and global buckling happen at the same time. In addition, it is considered that the panel does not fail due to crippling, stiffeners column buckling or other manufacturing restrictions. The final geometry is determined by minimising the area and, consequently, the weight of the panel. Findings The results obtained are compared with a classical method for sizing stiffened panels in aluminium. The weight prediction is validated by weight reductions in aircraft structures when comparing composite and aluminium alloys. Research limitations/implications The work is framed in conceptual design field, so hypotheses like material or stiffeners geometry shall be taken a priori. These hypotheses can be modified if it is necessary, but even so, the methodology continues being applicable. Practical implications The procedure presented in this paper allows designers to know composite structure weight of aircraft tails in commercial aviation or any lifting surface in unmanned aircraft field, even for unconventional configurations, in early stages of the design, which is an aid for them. Originality/value The contribution of this paper is the development of a new rapid methodology for conceptual design of composite panels and the feasible application to aircraft tails and also to unmanned aircraft.

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Sahadevan, Vijay, Yoann Bonnefon, and Tim Edwards. "A Meta-Heuristic Based Weight Optimisation for Composite Wing Structural Analysis." Applied Mechanics and Materials 5-6 (October 2006): 305–14. http://dx.doi.org/10.4028/www.scientific.net/amm.5-6.305.

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This paper presents a two-stage meta-heuristic approach to producing weight-optimised solutions needed prior to the detailed finite element analysis of composite wing. Composite wing covers are assumed to take the form of a group of stiffened sub-panels with varying skin and stiffener geometries according to the wing layout and loads. A population of limited solutions satisfying various design constraints was created using layout (skin and stiffener geometry), selected lay-ups, rule based stacking sequence and various assumed loads. The closed form analytical solutions of flat stiffened orthotropic plates are used for calculating buckling reserve factors and strength margins. For each sub-panel, a meta-heuristic rule was imposed to search for a suitable combination of skin and stiffener geometry. The criterion used was minimum weight satisfying laminate continuity accounting for manufacturability. Later, the optimised solutions for each sub-panel are converted into a format supported by the conventional finite element tool (NASTRAN). The use of meta-heuristic approach and their automation in Visual Basic for Applications resulted in fast convergence and potential time-saving compared to genetic algorithms.

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Li, Yeou-Fong, Habib Meda, and Walter Chen. "The Design and Analysis of Internally Stiffened GFRP Tubular Decks—A Sustainable Solution." Sustainability 10, no. 12 (December 1, 2018): 4538. http://dx.doi.org/10.3390/su10124538.

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The aim of this paper was to find an optimal stiffener configuration of thin-wall tubular panels made by glass fiber reinforced polymer (GFRP) composite material, which is a low carbon emission, low life cycle cost, and sustainable material. Finite-element analysis (FEA) was used to investigate the flexural and torsional stiffness of various internally stiffened sections of thin-wall GFRP decks. These decks consist of internally stiffened tubular profiles laid side by side and bonded together with epoxy to ensure the panel acts as an assembly. Three-dimensional models of the seven proposed decks were assembled with tubular profiles of different stiffener patterns. First, the non-stiffened tube profile was tested experimentally to validate the parameters used in the subsequent numerical analysis. Then, the finite element software, ANSYS, was used to simulate the flexural and torsional behavior of the decks with different stiffener patterns under bending and torsional loads. The decks with stiffener patterns such as “O” type, “V” type, and “D” type were found to be the most effective in bending. For torsion, there was a distinct tendency for deck panels with closed shaped stiffener patterns to perform better than their counterparts. Overall, the “O” type deck panel was an optimal stiffener configuration.

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Guofan, Zhang, Wan Chunhua, and Nie Xiaohua. "Study on compressive bearing capacity and efficiency of composite stiffened panels with different cross-sections." Journal of Physics: Conference Series 2336, no. 1 (August 1, 2022): 012027. http://dx.doi.org/10.1088/1742-6596/2336/1/012027.

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Abstract The post-buckling behavior and load bearing efficiency of composite panels with three kinds of stiffeners subject to the axial compression load were investigated. The progressive damage analysis models considering the failure of fiber/matrix, debonding of adhesive interface and stiffness degradation were established. The buckling load, failure load and failure process for three types of stiffened panels were derived. The numerical results coincide with tests, and the relative errors are all less than 6%, which indicates the FEM model given in this paper is reliable and practical. Further, the load bearing efficiency of these three kinds of stiffened panels with the same weight was discussed. Results reveal that the bearing capacity and efficiency of I-shape stiffened panel is the highest, and its bearing efficiency reaches 154.5N/g. The results of the paper provide a reference for engineering application of composite stiffened structure design.

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KALNINS, KASPARS, ROLANDS RIKARDS, JANIS AUZINS, CHIARA BISAGNI, HAIM ABRAMOVICH, and RICHARD DEGENHARDT. "METAMODELING METHODOLOGY FOR POSTBUCKLING SIMULATION OF DAMAGED COMPOSITE STIFFENED STRUCTURES WITH PHYSICAL VALIDATION." International Journal of Structural Stability and Dynamics 10, no. 04 (October 2010): 705–16. http://dx.doi.org/10.1142/s0219455410003695.

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A metamodeling methodology has been proposed for postbuckling simulation of stiffened composite structures with integrated degradation scenarios. The presence of artificial damage between the outer skin and stiffeners has been simulated as softening of the material properties in predetermined regions of the structure. The proposed methodology for the fast design procedure of axially or torsionally loaded stiffened composite structures is based on response surface methodology (RSM) and design and analysis of computer experiments (DACE). Numerical analyses have been parametrically sampled by means of the ANSYS/LS-DYNA probabilistic design toolbox extracting the load-shortening response curves in the preselected domain of interest. These response curves have been simplified using piecewise linear approximation identifying the buckling and postbuckling stiffness ratios along with the values of the skin and the stiffener buckling loads. Three stiffened panel designs and a closed box structure with preselected damage scenarios have been elaborated and validated with the tests performed within the COCOMAT project. The resulting design procedure provides a time-effective design tool for preliminary study and for elaboration of the optimum design guidelines for composite stiffened structures with material degradation restraints.

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Liang, Li, Pu Rong Jia, and Gui Qiong Jiao. "Progressive Failure Study of Discrete-Source Damage in Stiffened Composite Panels." Advanced Materials Research 314-316 (August 2011): 963–67. http://dx.doi.org/10.4028/www.scientific.net/amr.314-316.963.

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In this paper experimental and FEA methods were used to study the damage propagation and failure properties of stiffened composite panels with discrete-source damage. The influence of discrete-source damage on residual strength of the composite panel was also investigated. The research results indicated that notched composite panel was suitable to simulate the damage property of stiffened composite panel with discrete-source damage. It has showed that there is high strain concentration at the notch end. And break-through of the stringer can make the load path changed. Based on different failure criteria, FEA method with progressive failure procedure was used to simulate the damage progression and failure procedure of the notched stiffened composite panel effectively. The analytical result was in good correspondence with the experimental data.

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Ryabov, Alexander, Evgeny Maslov, Dmitry Strelets, and Vladimir Slobodchikov. "Computational Analysis of Compressed Stiffened Composite Panels with Impact Damage." Aerospace 6, no. 3 (February 27, 2019): 25. http://dx.doi.org/10.3390/aerospace6030025.

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A complex modeling technique is presented in this paper for a numerical analysis of compressed stiffened composite panels with impact damage. The numerical technique is based on the LS-Dyna code application, which simulates both the dynamic deformation of the panel subjected to a local impact and the quasi-static uniform compression of the panel within the local damage zone. The technique has been validated by both impact and compression experimental tests of the stiffened composite panel. The obtained numerical results show that impact damage to the composite panel can reduce the carrying capacity in more than 50% of damaged panels compared to undamaged panels.

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Bai, Yujiao, Zhonghai Xu, Jieren Song, Linlin Miao, Chaocan Cai, Fan Yang, Rongguo Wang, Xiaodong He, Yi Hong, and Xulun Dong. "Experimental and numerical analyses of stiffened composite panels with delamination under a compressive load." Journal of Composite Materials 54, no. 9 (September 23, 2019): 1197–216. http://dx.doi.org/10.1177/0021998319875209.

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L-shaped stiffened composite panels provide an efficient structure for engineering applications. However, they often produce delamination in the preparation and service process due to a series of factors. To study the effect of different types of delamination on the compressive strength of stiffened composite panels, ABAQUS finite element software was used in combine with the progressive damage subroutine user-defined field variable (USDFLD), and the finite element model was established based on cohesive theory to realize the prediction of the progressive failure process and strength of the stiffened composite panels. The results showed that the delamination of a stringer had a greater impact on the strength of the stiffened composite panels than did the debonding between the skin panel and a stringer and the delamination of the skin panel. The debonding delamination and delamination of a stringer exhibited delamination growth near the damage position during static compression, but delamination of the skin panel exhibited no delamination growth. The experimental results were in good agreement with the finite element simulation results, which verified the validity of the finite element model.

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Ye, Qiang, San San Xiao, and Pu Hui Chen. "Experimental and Numerical Study on Six- Point Bending Test of a Composite Stiffened Panel." Advanced Materials Research 328-330 (September 2011): 1309–12. http://dx.doi.org/10.4028/www.scientific.net/amr.328-330.1309.

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A six-point bending test was presented to simulate skin/stiffener debonding under anti-symmetrical loading conditions. A novel rig was design via which the anti-symmetrical bending deformations can be forced on to the specimens. Experimental study on six-point bending test of composite stiffened panels of T700/QY8911 was done by using this rig. The tests are numerically analyzed using the finite element code ABAQUS, modeling the entire stiffened panel by shell elements, and investigating the progressive delamination by means of the cohesive zone model. The results of numerical analyses are compared to the experimental ones in terms of load-displacement curves and debonding positions between skin and stringer. The experimental and numerical resulits show that the anti-symmetrical bending deformation is the main factor which results in the asymmetrical propagation of the debonding between the skin and the stiffener. The failure mechanisms of the test are similar to the ones which induces skin/stiffener debonding during post-buckling in the anti-symmetrical buckling mode.

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Bisagni, C., and R. Vescovini. "A fast procedure for the design of composite stiffened panels." Aeronautical Journal 119, no. 1212 (February 2015): 185–201. http://dx.doi.org/10.1017/s0001924000010332.

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AbstractThis paper describes the analysis and the minimum weight optimisation of a fuselage composite stiffened panel made from carbon/epoxy material and stiffened by five omega stringers. The panel investigated inside the European project MAAXIMUS is studied using a fast tool, which relies on a semi-analytical procedure for the analysis and on genetic algorithms for the optimisation. The semi-analytical approach is used to compute the buckling load and to study the post-buckling response. Different design variables are considered during the optimisation, such as the stacking sequences of the skin and the stiffener, the geometry and the cross-section of the stiffener. The comparison between finite element and fast tool results reveals the ability of the formulation to predict the buckling load and the post-buckling response of the panel. The reduced CPU time necessary for the analysis and the optimisation makes the procedure an attractive strategy to improve the effectiveness of the preliminary design phases.

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Meng, Shi Yang, Zahra Sharif Khodaei, and M. H. Aliabadi. "Localization of Barely Visible Impact Damage (BVID) in Composite Plates." Key Engineering Materials 627 (September 2014): 217–20. http://dx.doi.org/10.4028/www.scientific.net/kem.627.217.

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This paper exploits the implementation of a delay-and-sum imaging method using Lamb wave signals to localise barely visible impact damage (BVID) in quasi-isotropic composite panels. The structural discontinuities, such as opening and stiffener, has also been tested to reflect the common structural features of an aircraft and to examine the feasibility of the proposed detection technique. The prediction results are compared with ultrasonic C-scan images, which indicate location error for three different panels –flat panel, flat panel with an opening and stiffened panel. The accuracy is believed to be improved by increasing the number of transducers. Overall the proposed damage detection technique, with the use of only four transducers, demonstrated sufficient accuracy and efficiency in impact damage detection and can be applied alongside the traditional NDT inspections for providing a priori information of the impact damage location.

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ABRAMOVICH, HAIM, and TANCHUM WELLER. "REPEATED BUCKLING AND POSTBUCKLING BEHAVIOR OF LAMINATED STRINGER-STIFFENED COMPOSITE PANELS WITH AND WITHOUT DAMAGE." International Journal of Structural Stability and Dynamics 10, no. 04 (October 2010): 807–25. http://dx.doi.org/10.1142/s0219455410003750.

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Eight curved blade stringer-stiffened composite panels were tested under axial compression to obtain the "first" buckling and postbuckling behavior till collapse. Except for one panel, used as a reference panel, all of the panels had stringers without dropoff layers. Four panels contained either artificial damage or both artificial and impact-induced damage. Cyclic/repeated buckling was applied well in a relatively "deep" postbuckling region. It was demonstrated that neither repeated buckling, within the number of cycles applied in the present program, nor artificial damage and impact-induced damage, which were introduced into the panels, resulted in stiffness degradation of the panels. No premature failure of any of the tested panels was observed within their expected life cycle, i.e. exposure to a few hundred cycles deep in the postbuckling region, even in the presence of either type or a combination of the damage. All of the tested panels sustained repeated postbuckling loading till they were subjected to static loading aimed at determining their collapse loads. In spite of the present design, i.e. stiffeners with no dropoff plies aimed amongst others at providing a mechanism for initiating stiffener debonding, no skin–stringer separation was encountered till collapse of the panels. It was found that composite stringer-stiffened panels can be safely and repeatedly loaded in their deep postbuckling range with no degradation in their stiffness. Damage, due to either manufacturing or impact, which usually will result in rejection of a structural element, affected neither the load-carrying capacity nor the capability to withstand repeated loading in the relatively very deep postbuckling range within the present designed life cycle of the element. It was realized that manufacturing complexities and consequently costs can be reduced by employing a simplified design configuration where the use of a dropoff ply of the stringer base has been eliminated.

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Kumar, N. Jeevan, and P. Ramesh Babu. "Impact analysis of embedded delamination growth in stiffener of hybrid laminated composite stiffened panel." International Journal of Computational Materials Science and Engineering 06, no. 01 (March 2017): 1750001. http://dx.doi.org/10.1142/s2047684117500014.

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Modern aerostructures are predominantly of curved construction characterized by a skin and stiffeners. The latest generation of large passenger aircraft also mostly uses composite materials in their primary structure and there is trend toward the utilization of bonding of subcomponents. A stiffener is a simple beam attached to the skin to support the applied load. The stiffeners are provided to improve the bending stiffness of the panel which then improves the structural efficiency. Metals, unlike composites, offer plasticity effects to evade high stress concentrations during postbuckling. Under compressive load, composite structures show a wide range of damage mechanisms. A pre-damaged configuration is loaded to study the effect of delamination location and mode for delamination initiation and propagation. A parametric study is conducted to investigate the effect of the location of the delamination propagation when delamination is embedded inbetween plies of the stiffener, with the cases (i) delamination embedded at front and inbetween plies of the stiffener, (ii) delamination embedded at middle and inbetween plies of the stiffener, (iii) delamination embedded at the end and inbetween plies of the stiffener. Further, the influence of the location of the delamination in between plies of stiffener on load carrying capacity of the panel is investigated. The effect of delamination location on crack growth and collapse behavior is analyzed using analysis tool. The effect of delamination location on crack growth and collapse behavior is analyzed with numerical method of Virtual Crack Closure Technique.

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Kumar, Shashi, Rajesh Kumar, Sasankasekhar Mandal, and Atul K. Rahul. "The Prediction of Buckling Load of Laminated Composite Hat-Stiffened Panels Under Compressive Loading by Using of Neural Networks." Open Civil Engineering Journal 12, no. 1 (December 31, 2018): 468–80. http://dx.doi.org/10.2174/1874149501812010468.

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Background:Stiffened panels are being used as a lightweight structure in aerospace, marine engineering and retrofitting of building and bridge structure. In this paper, two efficient analytical computational tools, namely, Finite Element Analysis (FEA) and Artificial Neural Network (ANN) are used to analyze and compare the results of the laminated composite 750-hat-stiffened panels.Objective:Finite Element (FE) is an efficient and versatile method for the analysis of a complex problem. FE models have been used to generate data set of four different parameters. The four parameters are extensional stiffness ratio of skin in the longitudinal direction to the transverse direction, orthotropy ratio of the panel, the ratio of twisting stiffness to transverse flexural stiffness and smeared extensional stiffness ratio of stiffeners to that of the plate.Results and Conclusion:For training of ANN, multilayer feedforward back-propagation has been used as a network function with two-hidden layers in the neural network. The good network architecture is achieved after several iterations to predict the buckling load of the stiffened panel. ANN prediction for unknown new data set is in good agreement with FEA results of different cases, which show that ANN tool can be used for the design of complex structural problems in civil engineering and optimization of the laminated composite stiffened panel.

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Chang, Fei, Shu Lin Li, Xiao Peng Shi, Jun Jie Yin, and Dong Liang Bian. "Study of Hygrothermal Environment Effect on Compressive Strength of Stiffened Composite Panel." Applied Mechanics and Materials 575 (June 2014): 382–88. http://dx.doi.org/10.4028/www.scientific.net/amm.575.382.

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Focus on the hygrothermal environment problems of CF3031/CCF300 stiffened composite panel, the stiffened composite panel absorption were studied on in the 85%RH, 70°C hygrothermal environment, the moisture equilibrium content of the composites were calculated. The compressive test were carried on after hygrothermal aging, observed the change of the strain, and draw the conclusion of the flexural load and the breaking load, contrasted with the panel of before hygrothemal aging. The results show that the rate of remained flexural load was about 85% and the rate of remained breaking load was about 75% after hygrothemal aging.

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Yeh, Hsien-Yang, and Victor L. Chen. "Experimental Study and Simple Failure Analysis of Stitched J-Stiffened Composite Shear Panels." Journal of Reinforced Plastics and Composites 15, no. 11 (November 1996): 1070–87. http://dx.doi.org/10.1177/073168449601501101.

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Two fuselage type, stitched composite shear panels with J-shape stiffeners manufactured through the process of Resin Transfer Molding technique were tested. Through the diagonal tension tests, it was found that failure of each panel occurred at approximately 3.5 times its initial buckling load, indicating significant diagonal tension strength. Neither of the panels failed by stiffener “pop-off,” showing that the stitching helps to inhibit this failure mode. No other damage or delamination of the panel was visible as well. The newly developed generalized Yeh-Stratton (Y-S) failure criterion was used to evaluate the failure of these composite panels. Compared with several other failure criteria, the calculated failure loads are quite close to experimental results and all are conservative. The Y-S criterion is the most conservative for this case.

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Jin, Da Feng, Zhe Liu, and Zhi Rui Fan. "Optimization Methodology for Composite Stiffened Panel Based on Genetic Algorithm." Advanced Materials Research 952 (May 2014): 34–37. http://dx.doi.org/10.4028/www.scientific.net/amr.952.34.

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A novel optimization methodology for stiffened panel is proposed in this paper. The purpose of the optimization methodology is to improve the first buckling load of the panel which is obtained by finite element method. The stacking sequence of the stiffeners is taken as design variables. In order to ensure the manufacturability of design, the design guidelines of stacking sequence are taken into account. A DOE based on Halton Sequence makes the initial points of genetic algorithm spread more evenly in the design space of laminate parameters and consequently accelerates the search to convergence. The numerical example verifies the efficiency of this method.

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Riccio, Aniello, A. Raimondo, and F. Scaramuzzino. "Skin Stringer Debonding Evolution in Stiffened Composite Panels under Compressive Load: A Novel Numerical Approach." Key Engineering Materials 577-578 (September 2013): 605–8. http://dx.doi.org/10.4028/www.scientific.net/kem.577-578.605.

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In this paper, a numerical study, on the compressive behaviour of stiffened composite panels with skin-stringer debonding has been carried out. The analysis has been performed by adopting a novel robust (mesh and time step independent) finite elements based numerical model on a single stiffener panel with an artificial debonding. In order to prove the effectiveness of the proposed numerical tool, the results in terms of debonded area growth and compressive load versus applied displacement, have been compared with experimental data available in literature.

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Arranz, Sofía, Abdolrasoul Sohouli, and Afzal Suleman. "Buckling Optimization of Variable Stiffness Composite Panels for Curvilinear Fibers and Grid Stiffeners." Journal of Composites Science 5, no. 12 (December 15, 2021): 324. http://dx.doi.org/10.3390/jcs5120324.

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Automated Fiber Placement (AFP) machines can manufacture composite panels with curvilinear fibers. In this article, the critical buckling load of grid-stiffened curvilinear fiber composite panels is maximized using a genetic algorithm. The skin is composed of layers in which the fiber orientation varies along one spatial direction. The design variables are the fiber orientation of the panel for each layer and the stiffener layout. Manufacturing constraints in terms of maximum curvature allowable by the AFP machine are imposed for both skin and stiffener fibers. The effect of manufacturing-induced gaps in the laminates is also incorporated. The finite element method is used to perform the buckling analyses. The panels are subjected to in-plane compressive and shear loads under several boundary conditions. Optimization results show that the percentage difference in the buckling load between curvilinear and straight fiber panels depends on the load case and boundary conditions.

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Liu, Bin, Xiaoduan Zhang, and Yordan Garbatov. "Multi-Scale Analysis for Assessing the Impact of Material Composition and Weave on the Ultimate Strength of GFRP Stiffened Panels." Journal of Marine Science and Engineering 11, no. 1 (January 5, 2023): 108. http://dx.doi.org/10.3390/jmse11010108.

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A micro-meso-macro analysis framework based on the multi-scale method was employed to analyse the mechanical behaviour of marine GFRP stiffened panels. The study aims to establish a procedure for assessing the impact of material composition and weave on the ultimate strength of GFRP stiffened panels. The ultimate strength assessment was an essential step in the design process, and the investigation of construction materials has a great benefit to the lightweight design of marine composite structures. The micro- and meso-scale RVE models of components used in GFRP materials are established, and their failure criteria and stiffness degradation models are created using the user-defined material subroutine VUMAT in ABAQUS. The equivalent material properties at the micro-scale (meso-scale) obtained by a homogenisation method are used to define the meso-scale (macro-scale) mechanical properties in the finite element analyses. The multi-scale method assesses the macro-mechanics of composites, and it is shown that the ultimate strength of GFRP stiffened panels is mainly determined by the failure of CSM fibre bundles and WR yarns. Parametric study of the meso-mechanics of composite materials can provide an analysis tool to obtain the optimal macro ultimate strength of the composite stiffened panel.

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Ouadia, Mouhat, Khamlichi Abdellatif, Hasnae Boubel, Oumnia Elmrabet, Mohamed Rougui, El Mehdi Echebba, and Ahmed El Bouhmidi. "Dynamic buckling of laminated composite stringer stiffened CFRP panels under axial compression." MATEC Web of Conferences 149 (2018): 01044. http://dx.doi.org/10.1051/matecconf/201814901044.

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In this work, the dynamic buckling of stiffened panels is evolved numerically through a nonlinear incremental expression through making use of a specific time integration procedure via the finite element software program. the buckling and post-buckling behaviours of hat-stringer-stiffened composite curved panel under axial compression load .Dynamic buckling is extracted from the curve abandoning the very last shortening as a characteristic of time while the shape is subjected with the aid of a square compression pulse movement carried out inside the axial direction. The duration of the heart beat and the amplitude of curvature of decreasing of the cloth inside the band tormented by the warmth, the dynamic buckling motion, are constant. The method approach was proposed to predict the dynamic buckling load of curved panel. Finite element analysis was used to investigate these tests and the FE models were performed by ABAQUS.Approach to determine the reliability of the stiffened panel in dynamic buckling state.

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Chetwynd, D., F. Mustapha, K. Worden, J. A. Rongong, S. G. Pierce, and J. M. Dulieu-Barton. "Damage Localisation in a Stiffened Composite Panel." Strain 44, no. 4 (August 2008): 298–307. http://dx.doi.org/10.1111/j.1475-1305.2007.00371.x.

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Minnetyan, Levon, James M. Rivers, Christos C. Chamis, and Pappu L. N. Murthy. "Discontinuously Stiffened Composite Panel under Compressive Loading." Journal of Reinforced Plastics and Composites 14, no. 1 (January 1995): 85–98. http://dx.doi.org/10.1177/073168449501400106.

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29

Wang, Hong Qing, Qun Yan, Zong Hong Xie, and Xiao Dong Sui. "Modeling and Experimental Validation of Composite Stiffened Panels under Uniaxial Tension." Advanced Materials Research 791-793 (September 2013): 415–18. http://dx.doi.org/10.4028/www.scientific.net/amr.791-793.415.

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A finite element model implemented with a progressive damage propagation mechanism was generated to study the mechanical behavior of stiffened composite panels under uniaxial tension. Typical damage modes including fiber breakage, matrix crushing and delamination were considered in the model. Failure criteria with corresponding stiffness degradation technologies was used to predict the initiation and evolution of intra-laminar damage modes by a user-defined subroutine. Cohesive elements with thickness of 0.01mm were defined along the interface areas between the filler and the adjacent laminate layers for predicting the initiation and propagation of delamination. Corresponding tests on composite stiffened panel with a web cut-out were conducted. A good correlation between the numerical results and test data was obtained, which validated the finite element models. Both the numerical and experimental results conclude that the delamination in the flange around the cut-out region is the most critical failure mode for the composite stiffened panel under the uniaxial tensile load.

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Zhu, He. "Research of the Stability of Stiffened Composite Panel." Advanced Materials Research 834-836 (October 2013): 191–94. http://dx.doi.org/10.4028/www.scientific.net/amr.834-836.191.

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Composite stiffened plate structure is widely used in aviation, aerospace and marine areas , in order to ensure the structural reliability ,stability analysis is absolutely necessary. The finite element models are established. In this paper, subspace method and arc-length method are applied to analyze the stability of stiffened panel structure, both two methods are effective. Different stringer parameters will have some impact on the stability, including the number and height. Through analyzing the results, some conclusions are drew for increasing the stability.They have a certain value for engineering application.

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Wang, Zhe, Xiangming Chen, Xinxiang Li, Peng Zou, Junchao Yang, and Xue Bi. "An improved engineering method for bearing capacity calculation of stiffened curved composite panels." Journal of Physics: Conference Series 2338, no. 1 (September 1, 2022): 012005. http://dx.doi.org/10.1088/1742-6596/2338/1/012005.

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Abstract The application of composite materials in the primary structures of large aircraft fuselage is a development trend in recent years. As an indispensable structure in design and structure selection, accurate and efficient evaluation of its performance is of great significance to reduce the research cost. In this paper, the bearing capacity of stiffened curved composite panel is calculated through an improved engineering calculation method, and the effectiveness of the method is verified by test, which provides technical support for the rapidly design and selection of stiffened curved composite panels.

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Sharif-Khodaei, Z., Ramon Rojas-Diaz, and M. H. Aliabadi. "Lamb-Wave Based Technique for Impact Damage Detection in Composite Stiffened Panels." Key Engineering Materials 488-489 (September 2011): 5–8. http://dx.doi.org/10.4028/www.scientific.net/kem.488-489.5.

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The propagation characteristic of Lamb waves activated by Piezoelectric actuators and collected by sensors in a stiffened panel has been investigated. A network of actuators is used to scan the structure before and after the presence of damage. A diagnostic imaging algorithm has been developed based on the probability of damage at each point of the structure measured by the signal reading of sensors in the benchmark and damaged structure. A damage localization image is then reconstructed by superimposing the image obtained from each sensor-actuator path. Three-dimensional finite element model with a transducer network is modeled. Damage is introduced as a small softening area in the stiffened panel. Applying the imaging algorithm, the damage location was predicted with good accuracy. This method proves to be suitable for stiffened panels, where the complicated geometry and boundary reflections make the signal processing more complicated.

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Wang, Chen, Erming He, Zhibin Zhao, Zhiqi Liu, Xiaofeng Xue, and Yunwen Feng. "Method for Determining Repair Tolerance of Civil Aircraft Composite Structure." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 38, no. 4 (August 2020): 695–704. http://dx.doi.org/10.1051/jnwpu/20203840695.

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In order to solve the problem of repair tolerance in the actual repair process of civil aircraft composite structure, by which it is judged whether to repair or not. Based on the existing results and shortcomings, this paper proposes two complete solutions and procedures for the repair tolerance of composite structure, focusing on how to determine the upper limit of repair under the consideration of both economy and safety. Then taking the composite stiffened wall panel under compression as an example, considering the damage of perforation, combined with finite element analysis, these two methods are used to try to calculate the repair tolerance of the stiffened panel respectively. The results show that the repair tolerance results of the same composite stiffened wall panel obtained by the two schemes are relatively close and reasonable, which proves the feasibility of these two methods, and provides a reference for the engineering damage assessment of civil aircraft composite structure.

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Ružek, Roman, Konstantinos Tserpes, and Evaggelos Karachalios. "Compression after impact and fatigue behavior of CFRP stiffened panels." International Journal of Structural Integrity 6, no. 2 (April 13, 2015): 176–93. http://dx.doi.org/10.1108/ijsi-06-2014-0030.

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Purpose – Impact and fatigue are critical loading conditions for composite aerostructures. Compression behavior after impact and fatigue is a weak point for composite fuselage panels. The purpose of this paper is to characterize experimentally the compression behavior of carbon fiber reinforced plastic (CFRP) stiffened fuselage panels after impact and fatigue. Design/methodology/approach – In total, three panels were manufactured and tested. The first panel was tested quasi-statically to measure the reference compression behavior. The second panel was subjected to impact so as to create barely visible impact damage (BVID) at different locations, then to fatigue and finally to quasi-static compression. Finally, the third panel was subjected to impact so as to create visible impact damage (VID) at different locations and then to quasi-static compression. The panels were tested using ultrasound inspection just after manufacturing to check material quality and between different tests to detect impact and fatigue damage accumulation. During tests the mechanical behavior of the panel was monitored using an optical displacement measurement system. Findings – Experimental results show that the presence of impact damage significantly degrades the compression behavior of the panels. Moreover, the combined effect of BVID and fatigue was proven more severe than VID. Originality/value – The paper gives information about the compression after impact and fatigue behavior of CFRP fuselage stiffened panels, which represent the most realistic loading scenario of composite aerostructures, and describes an integrated experimental procedure for obtaining such information.

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Li, Si Qi, Zhi Dong Guan, Heng Chang Nie, Zeng Shan Li, and Xia Guo. "Compression Performance of Composite Stiffened Panels with Scarfed Holes and Scarfed Bonded Repairs." Advanced Materials Research 1095 (March 2015): 561–68. http://dx.doi.org/10.4028/www.scientific.net/amr.1095.561.

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This paper presents the experiment results conducted on three kinds of composite stiffened panels under compressive static load. The panels with scarfed holes and scarfed bonded repairs are analyzed with the virgin panels. The analysis shows that the scarfed bonded repair method has recovered the failure load and the stiffness above those of the virgin panel while the panel with a scarfed hole has remained 70% strength of the virgin panel. The scarfed hole has an effect on the buckling load as well as the stiffness and the failure load.

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Bhowmik, Krishnendu, Shamim Akhtar, Raj Kumar Kalshyan, Niloy Khutia, and Amit Roy Choudhury. "CNT Reinforced Laminated Composite under In-Plane Tensile Loading: A Finite Element Study." Materials Science Forum 978 (February 2020): 323–29. http://dx.doi.org/10.4028/www.scientific.net/msf.978.323.

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The present study is mainly aimed at investigating the distribution of in-plane stresses of a rectangular plate under localized uniform in-plane tensile loading through finite element analysis. The configuration used in the analysis is analogous to the case of premature failure of stiffened panel due to the termination of a stiffener in aircraft wing structure. In this current work, three different types of materials namely, isotropic, plain woven and transversely isotropic materials are being considered. Aluminium is taken as isotropic; high strength carbon/epoxy is being assigned as plain woven composite and carbon nanotube based hybrid composite is used as transversely isotropic material, due to their wide range of applications in aircraft structures. The effect of different materials on overall axial, transverse and shear stress distributions at different layers of the stiffened composite panels are demonstrated using finite element analyses. Further, the variations of these stresses along axial and transverse directions are also compared for different materials. It can be concluded from the present study that the peak stress developed near the load application zone should be incorporated in the design criteria of such plates to avoid failure.

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37

Liu, Tie Jun, Yong Zhang, Gang Li, and Feng Hui Wang. "Dynamic Response Analysis for the Solar-Powered Aircraft Composite Wing Panel with Viscoelastic Damping Layer." Applied Mechanics and Materials 105-107 (September 2011): 491–94. http://dx.doi.org/10.4028/www.scientific.net/amm.105-107.491.

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In design of solar powered aircraft wing panel, vibration properties of wing panel should be considered, especially for the peak value of dynamic response. In this research, a viscoelastic damping layer is built for vibration isolation, wing panel finite element models of stiffened and no-stiffened structures base on fiber-reinforced laminates with damping layer in the middle are built. Natural frequency and displacement response are analyzed with different thickness of damping layer and structures. Result shows natural frequencies decrease as thickness increased, and that of laminates are lower than stiffened structure. The maximum displacement response value decreased when thickness increased and that of laminates is higher than structured with stiffer. The presented work is helpful for type selection and designing of solar powered aircraft wing panel.

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38

Feddal, Ikram, Abdellatif Khamlichi, and Koutaiba Ameziane. "Effects of plies orientations and initial geometric imperfections on buckling strength of a composite stiffened panel." MATEC Web of Conferences 191 (2018): 00008. http://dx.doi.org/10.1051/matecconf/201819100008.

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The use of composite stiffened panels is common in several activities such as aerospace, marine and civil engineering. The biggest advantage of the composite materials is their high specific strength and stiffness ratios, coupled with weight reduction compared to conventional materials. However, any structural system may reach its limit and buckle under extreme circumstances by a progressive local failure of components. Moreover, stiffened panels are usually assembled from elementary parts. This affects the geometric as well as the material properties resulting in a considerable sensitivity to buckling phenomenon. In this work, the buckling behavior of a composite stiffened panel made from carbon Epoxy Prepregs is studied by using the finite element analysis under Abaqus software package. Different plies orientations sets were considered. The initial distributed geometric imperfections were modeled by means of the first Euler buckling mode. The nonlinear Riks method of analysis provided by Abaqus was applied. This method enables to predict more consistently unstable geometrically nonlinear induced collapse of a structure by detecting potential limit points during the loading history. It was found that plies orientations of the composite and the presence of geometric imperfections have huge influence on the strength resistance.

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D. Anitha et al.,, D. Anitha et al ,. "Analysis of Composite Orthogonal Grid Stiffened Flat Panel." International Journal of Mechanical and Production Engineering Research and Development 8, no. 2 (2018): 65–74. http://dx.doi.org/10.24247/ijmperdapr20187.

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40

Arakaki, Francisco K., and Alfredo R. Faria. "Composite-stiffened panel design under shear postbuckling behavior." Journal of Composite Materials 50, no. 26 (July 28, 2016): 3643–62. http://dx.doi.org/10.1177/0021998315623625.

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41

Akula, Venkata M. K. "Multiscale reliability analysis of a composite stiffened panel." Composite Structures 116 (September 2014): 432–40. http://dx.doi.org/10.1016/j.compstruct.2014.06.001.

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42

An, Haichao, Shenyan Chen, and Hai Huang. "Concurrent optimization of stacking sequence and stiffener layout of a composite stiffened panel." Engineering Optimization 51, no. 4 (August 8, 2018): 608–26. http://dx.doi.org/10.1080/0305215x.2018.1492570.

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43

Teyi, Nabam, and Santosh Kumar Tamang. "Stress Analysis of a Flange Stiffened FRP Composite Panel with Varying Stacking Sequence." Key Engineering Materials 847 (June 2020): 9–14. http://dx.doi.org/10.4028/www.scientific.net/kem.847.9.

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Large panel structures made of composites are common building units in aerospace industries. In order to increase the stiffness of such structures, the panel or skin is adhered to a flange and supported by a web. Such a stiffened panel is modeled as an arrangement of web, flange and panel with an interface between the flange and the panel. In this paper, three dimensional stress analysis of one such stiffened panel has been carried out using the finite element analysis. The geometric non-linearity has been assumed in the analysis. The effect of material anisotropy and the laminate stacking sequence on the stress components has been studied. Material Graphite Fiber Reinforced Polymeric (GFRP) composite has been used and, two different configurations were considered while the unidirectional prepregs were laid up in quasi-isotropic [0/0/0/0]2 and cross-ply [0/90/90/0]2. Subsequently, the coupled stress failure criterion has been used to predict the critical location of damage onset. When component damage indicator attained the value of 1.0, the component was considered to lose stiffness and structural integrity.

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44

Ghajari, M., Z. Sharif-Khodaei, and M. H. Aliabadi. "Impact Identification in Composite Stiffened Panels." Key Engineering Materials 525-526 (November 2012): 565–68. http://dx.doi.org/10.4028/www.scientific.net/kem.525-526.565.

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A number of small mass and large mass impacts on a sensorised aircraft stiffened panel were numerically simulated. Sensor signals and the contact force history were recorded during each impact. A significant difference was noticed between the small mass and large mass impacts with respect to the contact force. To distinguish between these two types of impacts, the Fast Fourier Transform was performed on the sensor signals and a categorisation criterion was defined. Finally, two separate Artificial Neural Networks were trained to approximate the peak contact force for each type of impact. It was found that the performance of these ANNs were better than a single ANN trained for both small and large mass impacts.

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Tuan, Trinh Anh, Tran Huu Quoc, and Tran Minh Tu. "FREE VIBRATION ANALYSIS OF LAMINATED STIFFENED CYLINDRICAL PANELS USING FINITE ELEMENT METHOD." Vietnam Journal of Science and Technology 54, no. 6 (December 7, 2016): 771. http://dx.doi.org/10.15625/0866-708x/54/6/8214.

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A study on the free vibration analysis of stiffened laminated composite cylindrical shell is described in this paper. The eight-noded isoparametric degenerated shell element is developed to model both shell panel and stiffeners by using the degenerated solid concept based on Reissner-Mindlin assumptions which taking to account the shear deformation and rotatory effect. Numerical results are presented and comparison is made with the published results from the literature and the good agreement is found. Parametric studies considering different geometrical variables of shell and stiffeners have also been carried out.

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Yue, Nan, Zahra Sharif Khodaei, and M. H. Ferri Aliabadi. "Passive Sensing of Sensorized Composite Panels: Support Vector Machine." Key Engineering Materials 713 (September 2016): 199–202. http://dx.doi.org/10.4028/www.scientific.net/kem.713.199.

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Strain readings recorded by surface mounted piezoelectric sensors due to impact events on composite panel are used to detect and characterize the impact. Sensor signals on a composite stiffened panels have been simulated by a valid numerical model. Applicability of least square support vector machines (LSSVM) on creating a meta-model to detect and characterize impact event has been investigated. In particular, the main advantage of LSSVM over other meta-modeling technique was found to be the smaller number of training data that is required. Experimental results on a composite panel has been used to validate the findings.

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Poulin, Kevin C., and Rimas Vaicaitis. "Vibrations of Stiffened Composite Panels With Smart Materials." Journal of Vibration and Acoustics 126, no. 3 (July 1, 2004): 370–79. http://dx.doi.org/10.1115/1.1760566.

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A numerical study of the use of electrorheological (ER) fluids and piezoelectric (PZT) actuators to control random vibrations of stiffened composite panels is presented. Active control of stiffness and damping is provided by the ER fluids and direct feedback control is provided by the PZT’s. New forms of transfer matrices are developed to include the effects of these smart materials. The modal equations of an equivalent uniform panel are converted into state-space form and digital stochastic feedback control is implemented. PZT direct feedback control is compared with digital stochastic feedback control. Parametric studies quantify the effect of actuator size and number, and ER fluid action.

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Sheinman, Izhak, and Yeoshua Frostig. "Post-Buckling Analysis of Stiffened Laminated Panel." Journal of Applied Mechanics 55, no. 3 (September 1, 1988): 635–40. http://dx.doi.org/10.1115/1.3125841.

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An analytical-numerical procedure is applied to investigate the post-buckling behavior of a composite laminated stiffened panel. The panel is modeled by plate elements for which the nonlinear equations are derived (via a variational principle) in terms of the lateral displacement and Airy stress function, and treated by resolving the variables into eigenfunctions in conjunction with a finite-difference scheme.

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Chen, Nian-Zhong, and C. Guedes Soares. "Ultimate Longitudinal Strength of Ship Hulls of Composite Materials." Journal of Ship Research 52, no. 03 (September 1, 2008): 184–93. http://dx.doi.org/10.5957/jsr.2008.52.3.184.

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A progressive collapse analysis method is proposed to predict the ultimate longitudinal strength of ship hulls of composite materials. The load-average strain curve derived from a progressive failure nonlinear finite element analysis is adopted for representing the behavior of each stiffened composite panel forming a hull cross section. The bending moment of the ship hull under a prescribed curvature is achieved by integrating the reaction force of each stiffened panel over a hull cross section based on the load-average strain curves. The ultimate longitudinal strength of a ship hull is obtained from the moment-curvature relationship of the ship hull, which is established by imposing progressively increasing curvatures of a hull cross section. An all-composite ship is analyzed as an application.

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KATO, Yoko, Ning HU, Masaki KAMEYAMA, and Hisao FUKUNAGA. "Optimum Design of Composite Wing Considering Stiffened Panel Buckling." Proceedings of Conference of Tohoku Branch 2002.37 (2002): 208–9. http://dx.doi.org/10.1299/jsmeth.2002.37.208.

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Journal articles on the topic 'Stiffened composite panel'

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