Journal articles: 'Stiffened Composite Wing'

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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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Butler, R. "Optimum design of composite stiffened wing panels — a parametric study." Aeronautical Journal 99, no. 985 (May 1995): 169–77. http://dx.doi.org/10.1017/s0001924000028335.

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AbstractThe program VICONOPT is used to find the optimum (least mass) dimensions of a range of stiffened wing panels which are subject to buckling and material strength constraints and are loaded in axial compression with a sinusoidal manufacturing imperfection. Design plots are presented to show the effects that various rib spacings and stiffener types have on optimum design mass. A simplified model of a complete wing box is used to illustrate the design of a full wing panel and plots of optimum values of design variables at various stations along the wing have been obtained. The results were chosen to illustrate the practicality of optimisation with reference to manufacture of a full wing panel and to show the effect of changing the sophistication of modelling and theory used for the range of panels considered. The important aspects of the choice of design variables and design concepts are highlighted and percentage savings in mass, compared with an optimum metal panel design, are given for the various (global) optima found along with some examples of (rejected) local optima.

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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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Hwu, Chyanbin, and Z. S. Tsai. "Aeroelastic Divergence of Stiffened Composite Multicell Wing Structures." Journal of Aircraft 39, no. 2 (March 2002): 242–51. http://dx.doi.org/10.2514/2.2945.

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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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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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KATO, Yoko, Masaki KAMEYAMA, Ning HU, and Hisao FUKUNAGA. "Optimum Design of Composite Wing Considering Stiffened Panel Buckling." Transactions of the Japan Society of Mechanical Engineers Series A 70, no. 691 (2004): 479–86. http://dx.doi.org/10.1299/kikaia.70.479.

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Yang, Xue-Yong, and Jun Xiao. "Research Progress on Analytical and Numerical Prediction of Curing Deformation in Thermoset for Large Composite Parts." Science of Advanced Materials 14, no. 4 (April 1, 2022): 669–81. http://dx.doi.org/10.1166/sam.2022.4247.

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Solidification deformation will produce certain drawbacks, so that a composite material part may not meet the requirements of a stress-free assembly for a modern aircraft. This issue holds particularly in the composite material part of large aircrafts. To predict and control this deformation, a novel method is applied for shifting the relaxation times of the composite based on its temperature and degree of cure. The choice of a suitable material model to simulate induced distortions is important to achieve the right-first-time approach. This work investigates the ability of the multi-physics model within a linear viscoelastic material model to predict induced distortions into an aerospace composite wing. It is shown that a L-shaped stiffened wall was less dominated by all deformations, but two stiffened wall panels were more dominated. Yet, wing box panels with four stiffened wall panels reduced the contribution to deformation. Their effects were included in the theory reported for the curing, and found to account for approximately 6.25% of the part deformation. The deformation effect could be analyzed by the proposed analytical solution, which was coupled with a cure kinetics model and a chemical shrinkage model to capture the multi-physics that take place during the curing.

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Romano, Fulvio, Monica Ciminello, Assunta Sorrentino, and Umberto Mercurio. "Application of structural health monitoring techniques to composite wing panels." Journal of Composite Materials 53, no. 25 (April 10, 2019): 3515–33. http://dx.doi.org/10.1177/0021998319843333.

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This detailed study proposes a structural health monitoring system which enables the identification, localisation, and correct measurement analysis, in relation to the damage and debonding induced by low energy impacts within aircraft composite wing panels. The said system has been envisaged as an offline system which aims to be considered as a valid alternative method in relation to the current first two maintenance approach levels: visual inspection, which is to be followed if necessary by ultrasonic scanning techniques. The architecture includes two different technologies which act at different frequency ranges: high-frequency sensors/actuators piezoceramics and low-frequency distributed fiber optic sensors. Experimental and numerical results on small stiffened panels are illustrated in this study, where technological verification and validation have been assessed within a laboratory-controlled environment. In addition, the potential benefit by utilising such techniques within the design of the aircraft composite structures has also been illustrated; in comparison with the current aircraft composite structures, a higher weight saving and better performing structures is foreseen.

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De Angelis, Giovanni, Michele Meo, D. P. Almond, S. G. Pickering, and U. Polimeno. "Impact Damage Detection in a Stiffened Composite Wing Panel Using Digital Shearography and Thermosonics." Key Engineering Materials 471-472 (February 2011): 904–9. http://dx.doi.org/10.4028/www.scientific.net/kem.471-472.904.

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There has been a growing interest in the use of composites especially in structural application ranging from aerospace to automotive and marine sectors. However, their performances under impact loading represent one of the major concerns as impacts may occur during manufacture, normal operations and maintenance. This paper presents two novel NDT techniques, thermosonics and digital shearography (DISH) to detect and assess barely visible impact damage (BVID) produced on a stiffened composite wing panel by unknown low energy impacts. Thermosonics is based on synchronized infrared imaging and ultrasonic excitation. Despite the apparent simplicity of the experimental setup, thermosonics involves a number of factors, e.g. acoustic horn location, horn crack proximity, horn-sample coupling etc., that significantly tend to influence both the degree and the period of the excitation. Then, a numerical-experimental procedure for the assessment of the size and depth of delamination by digital shearography (DISH) is proposed. The flaw detection capabilities of DISH have been evaluated by measuring the dynamic response of the delaminated area to applied stresses. The shearographic methodology is based on the recognition of the (0 1) resonance mode per defect. A simplified model of thin circular plate, idealized above each impacted area, is used to calculate the natural frequency of vibrating delamination. The numerical difference between experimental resonance frequencies and those computationally obtained is minimized using an unconstrained optimization algorithm in order to calculate the delamination depth. The results showed that thermosonics is a quick and effective method to detect and localize BVID damage while the combined shearography and optimization methodology was able to size and localize delamination due to low velocity impacts.

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Kumar, K. C. Nithin, Gopal Gupta, Saurabh Lakhera, and Amir Shaikh. "Structural Optimization of Composite Stiffened Panel of a Transport Aircraft Wing using CAE Tools." Materials Today: Proceedings 2, no. 4-5 (2015): 2588–94. http://dx.doi.org/10.1016/j.matpr.2015.07.213.

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Wiggenraad, J. F. M., X. Zhang, and G. A. O. Davies. "Impact damage prediction and failure analysis of heavily loaded, blade-stiffened composite wing panels." Composite Structures 45, no. 2 (June 1999): 81–103. http://dx.doi.org/10.1016/s0263-8223(98)00132-9.

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K.C., Nithin Kumar, Subhash Chavadaki, Amir Shaikh, Durgeshwar Pratap Singh, and Shwetank Avikal. "Weight optimization of a hat stiffened panel of a typical transport aircraft composite wing." Materials Today: Proceedings 26 (2020): 471–74. http://dx.doi.org/10.1016/j.matpr.2019.12.087.

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Skrna-Jakl, I. C., M. A. Stiftinger, and F. G. Rammerstorfer. "Numerical investigations of an imperfect stringer-stiffened composite wing torsion box—an analysis concept." Composites Part B: Engineering 27, no. 1 (January 1996): 59–69. http://dx.doi.org/10.1016/1359-8368(95)00007-0.

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dos Santos, Rogério Rodrigues, Tulio Gomes de Paula Machado, and Saullo Giovani Pereira Castro. "Support Vector Machine Applied to the Optimal Design of Composite Wing Panels." Aerospace 8, no. 11 (November 2, 2021): 328. http://dx.doi.org/10.3390/aerospace8110328.

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One of the core technologies in lightweight structures is the optimal design of laminated composite stiffened panels. The increasing tailoring potential of new materials added to the simultaneous optimization of various design regions, leading to design spaces that are vast and non-convex. In order to find an optimal design using limited information, this paper proposes a workflow consisting of design of experiments, metamodeling and optimization phases. A machine learning strategy based on support vector machine (SVM) is used for data classification and interpolation. The combination of mass minimization and buckling evaluation under combined load is handled by a multi-objective formulation. The choice of a deterministic algorithm for the optimization cycle accelerates the convergence towards an optimal design. The analysis of the Pareto frontier illustrates the compromise between conflicting objectives. As a result, a balance is found between the exploration of new design regions and the optimal design refinement. Numerical experiments evaluating the design of a representative upper skin wing panel are used to show the viability of the proposed methodology.

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Pigazzini, Marco S., Yuri Bazilevs, Andrew Ellison, and Hyonny Kim. "Isogeometric analysis for simulation of progressive damage in composite laminates." Journal of Composite Materials 52, no. 25 (April 22, 2018): 3471–89. http://dx.doi.org/10.1177/0021998318770723.

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The increasing popularity of composite materials in aerospace applications is creating the need for a new class of predictive methods and tools for the simulation of progressive damage in laminated fiber-reinforced composite structures. The unique challenges associated with modeling damage in these structures may be addressed by means of thin-shell formulations which are naturally developed in the context of Isogeometric Analysis. In this paper, we further validate our recently developed Isogeometric Analysis-based multi-layer shell model for progressive damage simulations using experimental data for low-velocity impact on a 24-ply flat panel. The validation includes a careful comparison of delamination and matrix damage patterns predicted by the Isogeometric Analysis-based simulation and those obtained from post-impact non-destructive evaluation of the damaged coupon. The Isogeometric Analysis-based formulation is then deployed on two additional examples: a stiffened panel and a full-scale UAV wing, to demonstrate its suitability for, and ease of application to, typical aerospace composite structures.

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Romano, Fulvio, Marco Barile, Gianpaolo Cacciapuoti, Jean-Luc Godard, Paolo Vollaro, and Philippe Barabinot. "Advanced OoA and Automated Technologies for the Manufacturing of a Composite Outer Wing Box." MATEC Web of Conferences 233 (2018): 00005. http://dx.doi.org/10.1051/matecconf/201823300005.

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This work resumes the results achieved until today in the European project AirGreen 2 of Clean Sky 2 programme, deriving from the application of two different dry preforming processes for the manufacturing of a composite outer wing box of the next generation turboprop aircrafts. Liquid Resin Infusion and Out of Autoclave techniques, by Hand-Layup and Automated Fiber Placement, are considered. The optimisation and validation of the manufacturing processes have been done according to key performance indexes: weight and cost reduction, lower energy consumption, high productivity and minimal reworking time, less intensive labour, minimal scrap and less waste of materials. The work has been performed through manufacturing tests and optimisation of the process parameters, implementation of several bagging techniques, numerical simulations of the infusion process and material characterization tests in operative conditions, from coupons level up to details and elements level (flat stiffened panels). Pro and cons, suggestions and technical considerations useful for the next step of the project (final manufacturing of large parts and components) are assessed.

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Jayasree, Nithin, Sadik Omairey, and Mihalis Kazilas. "Novel multi-zone self-heated composites tool for out-of-autoclave aerospace components manufacturing." Science and Engineering of Composite Materials 27, no. 1 (October 7, 2020): 325–34. http://dx.doi.org/10.1515/secm-2020-0033.

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AbstractIn this paper, a multi-zone self-heating composite tool is developed to manufacture out-of-autoclave complex and high-quality business jet lower wing stiffened composite panel. Autoclave manufacturing is regarded as a benchmark for manufacturing aerospace-grade composite parts. However, high accruing operational costs limit production rates thereby not being practical for smaller-scale companies. Therefore, significant work towards developing out-of-autoclave manufacturing is underway. In this study, a production line tool is developed with embedded heating fabric that controls temperature at the desired zones, replacing the need for autoclave cure. It investigates and identifies the optimal design parameters of the self-heating setup namely the placement of the heating fabric, zones, thermal management system, temperature distribution, heating rate and thermal performance using a thermal FEA model. The associated thermal characterisation of the tooling material and the part are measured for accurate simulation results. The design developed in this study will be used as production guideline for the actual tool.

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Riccio, Aniello, Andrea Sellitto, Salvatore Saputo, Giovanni Conte, and Mauro Zarrelli. "Thermo-Mechanical Behaviour of a Composite Stiffened Panel Undergoing the Tail-Pipe Fire Event." Key Engineering Materials 774 (August 2018): 101–6. http://dx.doi.org/10.4028/www.scientific.net/kem.774.101.

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In this paper, a numerical/experimental study is introduced on a carbon fibre reinforced polymer composite panel experiencing the tail-pipe fire phenomenon. This phenomenon, being of main concern for an Airplane during the engine-starting phase, can be qualitatively described as the ignition of a flame outside the nozzle impacting the adjacent aircraft structural components, such as the wing or the tail. The tail-pipe fire phenomenon is characterized by a variable duration and may cause the overheating or even the damage of the aircraft components. A numerical model, able to simulate the thermomechanical behaviour of composite structures under fire, is proposed. The presented approach, considering a strong coupling between the thermal and the structural fields and taking into account thermal and mechanical properties degradation, has been implemented in the commercial FEM software ABAQUS and applied to a stiffened composite panel. The numerical model has been validated by comparing the ABAQUS numerical results to the experimental data obtained by an ad-hoc campaign including mechanical tests under variable thermal conditions. The comparisons, performed in terms of Temperature-Time and Load-Strain evolution histories, showed a good agreement between numerical and experimental data, confirming the robustness of the proposed numerical tool and its effectiveness in predicting damage onset and propagation due to the presence of high thermal gradients.

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Esposito, Marco, Rinto Roy, Cecilia Surace, and Marco Gherlone. "Hybrid Shell-Beam Inverse Finite Element Method for the Shape Sensing of Stiffened Thin-Walled Structures: Formulation and Experimental Validation on a Composite Wing-Shaped Panel." Sensors 23, no. 13 (June 27, 2023): 5962. http://dx.doi.org/10.3390/s23135962.

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This work presents a novel methodology for the accurate and efficient elastic deformation reconstruction of thin-walled and stiffened structures from discrete strains. It builds on the inverse finite element method (iFEM), a variationally-based shape-sensing approach that reconstructs structural displacements by matching a set of analytical and experimental strains in a least-squares sense. As iFEM employs the finite element framework to discretize the structural domain and as the displacements and strains are approximated using element shape functions, the kind of element used influences the accuracy and efficiency of the iFEM analysis. This problem is addressed in the present work through a novel discretization scheme that combines beam and shell inverse elements to develop an iFEM model of the structure. Such a hybrid discretization paradigm paves the way for more accurate shape-sensing of geometrically complex structures using fewer sensor measurements and lower computational effort than traditional approaches. The hybrid iFEM is experimentally demonstrated in this work for the shape sensing of bending and torsional deformations of a composite stiffened wing panel instrumented with strain rosettes and fiber-optic sensors. The experimental results are accurate, robust, and computationally efficient, demonstrating the potential of this hybrid scheme for developing an efficient digital twin for online structural monitoring and control.

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Daraji, Ali H., Jack M. Hale, and Jianqiao Ye. "Optimisation of energy harvesting for stiffened composite shells with application to the aircraft wing at structural flight frequency." Thin-Walled Structures 161 (April 2021): 107392. http://dx.doi.org/10.1016/j.tws.2020.107392.

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Hiremath, Pavan, Sathyamangalam Ramanarayanan Viswamurthy, Manjunath Shettar, Nithesh Naik, and Suhas Kowshik. "Damage Tolerance of a Stiffened Composite Panel with an Access Cutout under Fatigue Loading and Validation Using FEM Analysis and Digital Image Correlation." Fibers 10, no. 12 (December 8, 2022): 105. http://dx.doi.org/10.3390/fib10120105.

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Aircraft structures must be capable of performing their function throughout their design life while meeting safety objectives. Such structures may contain defects and/or damages that can occur for several reasons. Therefore, aircraft structures are inspected regularly and repaired if necessary. The concept of combining an inspection plan with knowledge of damage threats, damage growth rates, and residual strength is referred to as “damage-tolerant design” in the field of aircraft design. In the present study, we fabricated a composite panel with a cutout (which is generally found in the bottom skin of the wing) using a resin infusion process and studied the damage tolerance of a co-cured skin-stringer composite panel. The composite panel was subjected to low-velocity impact damage, and the extent of damage was studied based on non-destructive inspection techniques such as ultrasonic inspection. Fixtures were designed and fabricated to load the composite panel under static and fatigue loads. Finally, the panel was tested under tensile and fatigue loads (mini TWIST). Deformations and strains obtained from FE simulations were compared and verified against test data. Results show that the impact damages considered in this study did not alter the load path in the composite panel. Damage did not occur under the application of one block (10% life) of spectrum fatigue loads. The damage tolerance of the stiffened skin composite panel was demonstrated through test and analysis.

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Cestino, Enrico, Giacomo Frulla, Paolo Piana, and Renzo Duella. "Numerical/Experimental Validation of Thin-Walled Composite Box Beam Optimal Design." Aerospace 7, no. 8 (July 31, 2020): 111. http://dx.doi.org/10.3390/aerospace7080111.

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Thin-walled composite box beam structural configuration is representative of a specific high aspect ratio wing structure. The optimal design procedure and lay-up definition including appropriate coupling necessary for aerospace applications has been identified by means of “ad hoc” analytical formulation and by application of commercial code. The overall equivalent bending, torsional and coupled stiffness are derived and the accuracy of the simplified beam model is demonstrated by the application of Altair Optistruct. A simple case of a coupled cantilevered beam with load at one end is introduced to demonstrate that stiffness and torsion angle distribution does not always correspond to the trends that one would intuitively expect. The maximum of torsional stiffness is not obtained with fibers arranged at 45° and, at the maximum torsional stiffness, there is no minimum rotation angle. This observation becomes essential in any design process of composite structures where the constraints impose structural couplings. Furthermore, the presented theory is also extended to cases in which it is necessary to include composite/stiffened hybrid configurations. Good agreement has been found between the theoretical simplified beam model and numerical analysis. Finally, the selected composite configuration was compared to an experimental test case. The numerical and experimental validation is presented and discussed. A good correlation was found confirming the validity of the overall optimization for the optimal lay-up selection and structural configuration.

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Wang, Zhe, Xiangming Chen, Peng Zou, Xinxiang Li, Qingxianglong Liang, and Junchao Yang. "Parametric Analysis on Three-Points Bending Test of Typical Skin-Stringer Structure." Journal of Physics: Conference Series 2085, no. 1 (November 1, 2021): 012043. http://dx.doi.org/10.1088/1742-6596/2085/1/012043.

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Abstract Stringer-skin debonding was one of the most important failure models in stiffened composite panels. In this paper, three-points bending tests were performed on representative stringer-skin structure of composite wing to simulate the flange-skin interface behavior and to obtain the failure mode and failure load. A 3D finite element model was built by using ABAQUS software to simulate interface failure with cohesive zone model. The numerical results agree well with test data, which validate the rationality of the finite element model. Hence the influence of factors during manufacture, installation and test in three-points bending tests, such as off-axis displacement, inclination loading and span, is studied. Results show that the initial debonding load and failure load of specimen decrease as the displacement from loading axis to central axis increases. The load of specimen decreases as the span increases. The influence of inclination loading is insignificant when the inclination angle is less than 6 degree. However, the initial debonding load and failure load of specimen decreases in varying degrees as the inclination loads increases. Furthermore, the initial debonding load decreases rapidly.

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Fenner, Patrick, Andrew Watson, and Carol Featherston. "Modelling Infinite Length Panels Using the Finite Element Method." International Journal of Structural Stability and Dynamics 16, no. 07 (August 3, 2016): 1750038. http://dx.doi.org/10.1142/s0219455417500389.

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This paper compares three finite element models for determining the buckling and post-buckling performance of infinite length thin walled composite and metal stiffened panels — such as for modeling theoretical aircraft upper wing skin panels — namely single bay, double half-bay and quad half-bay models. The quad half-bay model is shown to be the ideal model as all wavelengths of buckling are permitted. This model gives an accurate estimate of postbuckling behavior that can include advanced behavior such as mode jumping or collapse while the single bay and double half-bay models are more restrictive and do not allow for accurate mode jumping to take place. Sample panels are analyzed for buckling performance using the computer program VICONOPT, which assumes an infinite length structure based on exact strip theory. This analysis is then compared to results from the quad half-bay FEM model, using the Abaqus solver, where the two models are in good agreement for the initial buckling performance for both the metal and composite panels. Buckling prediction for the quad half-bay model is within [Formula: see text] for the critical buckling mode, and within [Formula: see text] of all compared modes; and postbuckling performance compares well with the results of previous investigation of the same sample panel geometry.

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Capriotti, Margherita, and Francesco Lanza di Scalea. "Robust non-destructive inspection of composite aerospace structures by extraction of ultrasonic guided-wave transfer function in single-input dual-output scanning systems." Journal of Intelligent Material Systems and Structures 31, no. 5 (January 9, 2020): 651–64. http://dx.doi.org/10.1177/1045389x19898266.

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Ultrasonic guided-wave testing is widely utilized for damage detection in plate-like structural components, including composite aircraft fuselage and wing panels. Many guided-wave tests involve transducer scanning to cover finite areas, a task that is most effectively performed by non-contact wave transduction means. The most common guided-wave test implementation consists of a “single-input single-output” scheme. The single-input single-output scheme leads to a transfer function that is convolved with the particular frequency response of the transducers and that of the transducer-to-structure paths (in both excitation and detection). These responses can be unknown or generally variable, especially in non-contact scanning systems and impact excitations. This article proposes a “single-input dual-output” scheme for ultrasonic guided-wave testing in scanning systems that is based on a deconvolution operation. The single-input dual-output scheme better isolates the structural transfer function that is the only property affected by the presence of possible damage. The single-input dual-output scheme was applied to two guided-wave scanning systems under development to detect impact-type damage in stiffened skin-to-stringer panels representative of modern composite aircraft construction. The results demonstrate the dual-output technique and also shed some light on the role of the different frequency bands for the detection of damage at different locations of the skin-to-stringer assembly.

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Yu, Fei, Fei Yuan, Zhe Wang, and Xiangming Chen. "Experimental and Numerical Investigation on the Failure Behaviour of Multi-Stiffener Composite Panel Under Compression." Journal of Physics: Conference Series 2557, no. 1 (July 1, 2023): 012098. http://dx.doi.org/10.1088/1742-6596/2557/1/012098.

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Abstract Multi-stiffener composite panels are the primary components widely used in aircraft structures such as fuselage, wing box, and rear bulkhead, but they are typically sensitive to compressive load-induced buckling as the thin-walled nature of these components. This paper presents a numerical and experimental study on the compressive failure behaviour of multi-stiffener composite panels. A progressive damage model based on Hashin’s failure criteria was employed to model the failure behaviour of the laminates, while a cohesive zone model (CZM) was used to replicate the delamination between the stiffeners and the flat skin. The predicted initial buckling threshold and ultimate failure load agree well with the experimental data, which validated the simulation tool developed in this study. The failure mode of the MSCP was then dominated by the breakage of the stiffeners at their middle cross-section, simultaneously accompanied by delamination at the skin/stiffener interfaces. Results indicated that the initial out-of-plane buckling can cause a high level of peeling and shearing effects at the skin/stiffener interfaces, leading to skin/stiffener debonding as the compression increase. This subsequently introduced an additional bending effect to the stiffener, which eventually caused breakage in the middle stiffeners.

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LINDE, PETER. "VIRTUAL TESTING OF STIFFENED COMPOSITE PANELS AT AIRBUS." International Journal of Structural Stability and Dynamics 10, no. 04 (October 2010): 589–600. http://dx.doi.org/10.1142/s0219455410003634.

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The steady increase of composite parts in civil aircraft over the last three decades has recently been followed by a radical increase of weight percentage composites in the structure. In the most recent long range aircraft under development by both Boeing and Airbus, most major structure components, not only of the wings but also of the fuselage, now consist of composites. This necessitates an increased use of efficient structural simulation capabilities. One important aspect of this is the virtual testing of shear compression panels at Airbus, which will be presented here. Having served well for Glare during the A380 development, it is currently undergoing considerable development to extend its capacity to composites. Summarized under the designation "Simulation of Panels in AirCraft", (SIMULPAC), this platform is here introduced and described in terms of functioning and methodology. It has played a major role in the initial A350 developments: in the first designs, during virtual testing of the configuration of new components and for predictions of the first real shear compression panels.

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Papadopoulos, F., D. Aiyappa, R. Shapriya, E. Sotirchos, H. Ghasemnejad, and R. Benhadj-Djilali. "Advanced Natural Stitched Composite Materials in Skin-Stiffener of Wind Turbine Blade Structures." Key Engineering Materials 525-526 (November 2012): 45–48. http://dx.doi.org/10.4028/www.scientific.net/kem.525-526.45.

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In this paper the failure behaviour of natural stitched composite materials in the skin-stiffener of wind turbine blade structures are investigated. For this study, the laminated composite beams were stitched using Flax yarns before curing process. Two stiffener structures of T-beam and Box-beam are studied in this paper. These specimens were tested under quasi-static loading condition to compare the failure resistance of adhesive and stitched bonding methods. Furthermore, the cohesive zone modelling (CZM) which is known as a variation in the cohesive stresses with the interfacial opening displacement along the localised fracture process zone is used to predict bonding failure in the skin-stiffener of wind turbine blade structures.

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Doggui, Mohamed Montassar, Wael Touihri, Mondher Yahiaoui, and Moez Chafra. "Numerical investigation on aircraft wing stiffener composite material integration." Aerospace Systems 2, no. 2 (June 4, 2019): 137–45. http://dx.doi.org/10.1007/s42401-019-00027-9.

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Plenča, Stipe, and Albert Zamarin. "Structural Design of a Composite Trimaran." Journal of Maritime & Transportation Science 2, Special edition 2 (April 2018): 71–88. http://dx.doi.org/10.18048/2016-00.63.

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This paper presents a project of a composite trimaran structure, designed and built for competing at the Hydro Contest 2016 competition at Geneva Lake. Concept of the contest is to raise the awareness of tomorrow’s engineers, industrialists, opinion leaders and the public of what is at stake with regard to energy efficiency in the sea transportation of goods and passengers. In addition, to be the laboratory of tomorrow’s boats, particularly enabling the most innovative ideas to be developed in collaboration with the industrial partners. Designed boats must have technological innovations enabling them to achieve the most efficient use of energy. Therefore, the goal was to design, construct lightweight structure, within simple closed rules, with a satisfactory stiffens, and strength as well as to strive for more efficient transport, which means higher speed with minimal energy consumption. An analysis of project variants was made with regard to the hull shape, material, and technology of the fabrication and for the adopted variant, a computer structure model was developed, and the FEA was carried out. The structure is divided into three main sections analysed individually: hulls, front wing and rear wing along with rudder. Calculation was made for the worst load case, i.e. mass transfer, while wings were analysed at the highest advancing speed. The boat has structurally met all requirements since there were no structural problems in testing and competing.

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Butler, R., M. Lillico, J. R. Banerjee, M. H. Patel, and G. T. S. Done. "Sequential use of conceptual MDO and panel sizing methods for aircraft wing design." Aeronautical Journal 103, no. 1026 (August 1999): 389–97. http://dx.doi.org/10.1017/s0001924000064617.

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Abstract The optimisation results for composite and metallic versions of a regional aircraft wing are compared using the multidisciplinary optimisation (MDO) program CALFUNOPT. The program has been developed for the conceptual design stage and models the wing using just 11 beam elements. The wing has been optimised for three combinations of the following constraint cases: static strength; aeroelastic roll efficiency (represented by limiting the twist of the wing for an aileron loading) and aeroelastic divergence. As expected, comparison shows that the composite wing designs are significantly lighter than the metallic ones, due to the well-known tailoring of the composite material. However, the simple model reveals some insight that may be useful to the designer, and which could be lost within a more detailed finite element approach. The upper-skin compression panels produced by the conceptual MDO program, for both versions of the wing, have then been optimised using the more detailed and accurate panel sizing tool VICONOPT, which takes buckling into account. Such optimisation increases the panel mass by 5-10% and also provides a suitable ratio of stiffener to skin area for use in the conceptual MDO model.

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Dervilis, Nikolaos, R. Barthorpe, Wieslaw Jerzy Staszewski, and Keith Worden. "Structural Health Monitoring of Composite Material Typical of Wind Turbine Blades by Novelty Detection on Vibration Response." Key Engineering Materials 518 (July 2012): 319–27. http://dx.doi.org/10.4028/www.scientific.net/kem.518.319.

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New generations of offshore wind turbines are playing a leading role in the energy arena. One of the target challenges is to achieve reliable Structural Health Monitoring (SHM) of the blades. Fault detection at the early stage is a vital issue for the structural and economical success of the large wind turbines. In this study, experimental measurements of Frequency Response Functions (FRFs) are used and identification of mode shapes and natural frequencies is accomplished via an LMS system. Novelty detection is introduced as a robust statistical method for low-level damage detection which has not yet been widely used in SHM of composite blades. Fault diagnosis of wind turbine blades is a challenge due to their composite material, dimensions, aerodynamic nature and environmental conditions. The novelty approach combined with vibration measurements introduces an online condition monitoring method. This paper presents the outcomes of a scheme for damage detection of carbon fibre material in which novelty detection approaches are applied to FRF measurements. The approach is demonstrated for a stiffened composite plate subject to incremental levels of impact damage.

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Kennedy, D., M. Fischer, and C. A. Featherston. "Recent developments in exact strip analysis and optimum design of aerospace structures." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 221, no. 4 (April 1, 2007): 399–413. http://dx.doi.org/10.1243/0954406jmes432.

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The current paper outlines recent developments to algorithms and software for critical buckling and natural vibration analysis and optimum design of prismatic plate assemblies, based on the exact strip approach and the Wittrick—Williams algorithm. The current paper acts as a single source document discussing recent progress and planned future explorations in: initial local postbuckling of stiffened panels; discrete optimization of composite structures to satisfy manufacturing requirements; discontinuous cost functions; constraints on fundamental natural frequencies and frequency-free bands; a feasibility study of response surface optimization; and multi-level optimization of composite aircraft wings. The numerous references provide fuller technical details and illustrative examples.

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Shahriari, Behrooz, Ali Nazari, and Mostafa Sahraei. "Buckling Analysis of A Composite Stiffend Panel Structure In The Aircraft’s Wing." Mechanic of Advanced and Smart Materials 2, no. 3 (November 22, 2022): 328–46. http://dx.doi.org/10.52547/masm.2.3.328.

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Navadeh, Navid, Ivan Goroshko, Yaroslav Zhuk, Farnoosh Etminan Moghadam, and Arash Soleiman Fallah. "Finite Element Analysis of Wind Turbine Blade Vibrations." Vibration 4, no. 2 (April 5, 2021): 310–22. http://dx.doi.org/10.3390/vibration4020020.

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The article is devoted to the practical problem of computer simulation of the dynamic behaviour of horizontal axis wind turbine composite rotor blades. This type of wind turbine is the dominant design in modern wind farms, and as such its dynamics and strength characteristics should be carefully studied. For this purpose, in this paper the mechanical model of a rotor blade with a composite skin possessing a stiffener was developed and implemented as a finite element model in ABAQUS. On the basis of this computer model, modal analysis of turbine blade vibrations was performed and benchmark cases for the dynamic response were investigated. The response of the system subjected to a uniform underneath pressure was studied, and the root reaction force and blade tip displacement time histories were obtained from the numerical calculations conducted.

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Tripathy, A. K., H. J. Patel, and S. S. Pang. "Bending Analysis of Laminated Composite Box Beams." Journal of Engineering Materials and Technology 116, no. 1 (January 1, 1994): 121–29. http://dx.doi.org/10.1115/1.2904247.

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Box beams are widely used in weight reduction structures such as aircraft wings. The use of composite box beams further reduces the weight factor for such structures with the same deflection and stress as that of isotropic box beams. The difference in the behavior of composite box beams with different fiber orientation, number of plies, and number of stringers also provides a wide range of designing parameters to achieve the required performance for a given problem. A bending analysis has been carried out for the study of deflections and stresses for box beams of different material (isotropic and laminated composites), size, and number of stringers subjected to different kinds of loading conditions. A finite element model has been developed based on the strain energy principle, and the results are compared with an available commercial code “COSMOS/M.” Experiments using aluminum and scotchply composite laminates were conducted to verify the results. An optimal design for size and number of stiffeners for a given loading condition has been achieved. Investigations have also been carried out to find the effect of transverse shear on the span-wise normal stress.

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Zucco, G., V. Oliveri, M. Rouhi, R. Telford, G. Clancy, C. McHale, R. O’Higgins, T. M. Young, P. M. Weaver, and D. Peeters. "Static test of a variable stiffness thermoplastic composite wingbox under shear, bending and torsion." Aeronautical Journal 124, no. 1275 (January 22, 2020): 635–66. http://dx.doi.org/10.1017/aer.2019.161.

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AbstractAutomated manufacturing of thermoplastic composites has found increased interest in aerospace applications over the past three decades because of its great potential in low-cost, high rate, repeatable production of high performance composite structures. Experimental validation is a key element in the development of structures made using this emerging technology. In this work, a $750\times640\times240$ mm variable-stiffness unitised integrated-stiffener out-of-autoclave thermoplastic composite wingbox is tested for a combined shear-bending-torsion induced buckling load. The wingbox is manufactured by in-situ consolidation using a laser-assisted automated tape placement technique. It is made and tested as a demonstrator section located at 85% of the wing semi-span of a B-737/A320 sized aircraft. A bespoke in-house test rig and two aluminium dummy wingboxes are also designed and manufactured for testing the wingbox assembly which spans more than 3m. Prior to testing, the wingbox assembly and the test rig were analysed using a high fidelity finite element method to minimise the failure risk due to the applied load case. The experimental test results of the wingbox are also compared with the predictions made by a numerical study performed by nonlinear finite element analysis showing less than 5% difference in load-displacement behaviour and buckling load and full agreement in predicting the buckling mode shape.

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Son, C. Y., H. I. Byun, K. H. Kim, J. K. Choi, and J. Y. Shin. "An Analysis and Experimental Study of the Rotor Blade with Composite Material Fiber Reinforced Plastics." Key Engineering Materials 306-308 (March 2006): 851–56. http://dx.doi.org/10.4028/www.scientific.net/kem.306-308.851.

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In these days, large-scale wind turbines are being made of the Glass Fiber Reinforced Plastic (hereinafter F.R.P). Some reinforcement stiffeners such as carbon fiber and polyamide (Kevlar) are not economical for the wind turbine. In addition, the steel or aluminum alloy, featuring heavy weight and metallic fatigue load, is not suitable for global use, except very small-scale wind turbines. In this study, we manufactured a 10kW-grade small Rotor Blade with the F. R. P featuring high stiffness and good dynamic behavior characteristic, and carried out experiments for understanding the bending behavior characteristic of the fatigue load and bending load. And, we examined the experiment results through the Finite Element Method. We compared the experiment results and FEM analysis outputs using the commercial ANSYS FEM program.

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Qiu, Zixue, Lei Qiu, Jiang Yuan, and Guan Lu. "On research of a phase synthesis time reversal focusing method for damage imaging of complex composite structures." Journal of Intelligent Material Systems and Structures 24, no. 2 (September 27, 2012): 209–25. http://dx.doi.org/10.1177/1045389x12461078.

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Time reversal focusing method has been proved to be an effective method for active Lamb wave–based structural health monitoring. In this article, aiming at developing a practical method for online localization of damage on aircraft composite structures that can take advantage of time reversal focusing and do not rely on the transfer function, a phase synthesis–based time reversal focusing method was proposed. In this method, damage images are given out directly through time reversal focusing, and the other imaging processes such as the delay-and-sum imaging method adopted in many researches of time reversal focusing are not needed. Based on the damage imaging method, a structural health monitoring demonstration system was built on a composite panel of an aircraft wing box with many bolt holes and stiffeners. The demonstrated results show that this method can estimate the positions of damages efficiently with a low sensitivity of group velocity errors and a high antijamming capability.

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41

Lee, Sang-Lae. "Active vibration suppression of wind turbine blades integrated with piezoelectric sensors." Science and Engineering of Composite Materials 28, no. 1 (January 1, 2021): 402–14. http://dx.doi.org/10.1515/secm-2021-0039.

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Abstract As the wind turbine size gets larger, the optimal design of blades, which is a major source of energy for the wind turbines and also the cause of loads, is becoming more important than anything else. Therefore, reducing the load on the blade should be the top priority in designing a blade. In this article, we studied the vibration control of the stiffened wind blades subjected to a wind load with piezoelectric sensors and actuators to mitigate fluctuations in loading and adding damping to the blade. The model is a laminated composite blade with a shear web and the PZT piezomaterial layers embedded on the top and bottom surfaces act as a sensor and actuator, respectively. A uniformly distributed external wind load is assumed over the entire plate surface for simplicity. The first-order shear deformation (FSDT) theory is adopted, and Hamilton’s principle is used to derive the finite element equation of motion. The modal superposition technique and the Newmark- β \beta method are used in numerical analysis to calculate the dynamic response. Using the constant gain negative velocity feedback control algorithm, vibration characteristics and transient responses are compared. Furthermore, vibration control at various locations of the shear webs subjected to an external load is discussed in detail. Through various calculation results performed in this study, this article proposes a method of designing a blade that can reduce the load by actively responding to the external load acting on the wind turbine blade.

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Park, Sunghyun, In-Gul Kim, Seokje Lee, and Oo-Chul Jun. "Optimal Design of the Composite Hat-shaped Stiffeners for Simplified Wing Box with Embedded Array Antenna." Journal of The Korean Society for Composite Materials 25, no. 6 (December 31, 2012): 224–29. http://dx.doi.org/10.7234/kscm.2012.25.6.224.

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43

Qiu, Lei, Shen Fang Yuan, and Tian Xiang Huang. "A Time Reversal Imaging Method without Relying on Transfer Function for Impact and Damage Monitoring of Composite Structures." Applied Mechanics and Materials 330 (June 2013): 542–48. http://dx.doi.org/10.4028/www.scientific.net/amm.330.542.

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Composite structures adopted in aerospace structures have attracted much interest to structural health monitoring (SHM) for localization of impact and damage positions due to their poor impact resistance properties. Propagation mechanism and frequency dispersion characteristics of Lamb wave signals on composite structures are more complicated than that on simple aluminum plates. Recently, much attention has been paid to the research of time reversal focusing method because this method shows a promising advantage to give a focusing image of the structural damage, improve the signal-to-noise ratio and compensate the dispersion of Lamb wave signals. In this paper, aiming at developing a practical method for on-line localization of impact and damage on aircraft composite structures which can take advantage of time reversal focusing and does not rely on the transfer function, a new phase synthesis based time reversal focusing method is proposed. Impact and damage images are given out directly through time reversal focusing based on phase synthesis process of the signals. A SHM demonstration system is built on a composite panel of an aircraft wing box with many bolt holes and stiffeners using the phase synthesis based time reversal focusing method. The demonstration results show that this method can estimate the positions of impact and damage efficiently with a low sensitivity of velocity errors.

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Mandell, John F., Douglas S. Cairns, Daniel D. Samborsky, Robert B. Morehead, and Darrin J. Haugen. "Prediction of Delamination in Wind Turbine Blade Structural Details." Journal of Solar Energy Engineering 125, no. 4 (November 1, 2003): 522–30. http://dx.doi.org/10.1115/1.1624613.

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Delamination between plies is the root cause of many failures of composite material structures such as wind turbine blades. Design methodologies to prevent such failures have not been widely available for the materials and processes used in blades. This paper presents simplified methodologies for the prediction of delamination in typical structural details in blades under both static and fatigue loading. The methodologies are based on fracture mechanics. The critical strain-energy release rate, GIC and GIIC, are determined for opening mode (I) and shearing mode (II) delamination cracks; fatigue crack growth in each mode is also characterized. These data can be used directly for matrix selection and as properties for the prediction of delamination in structural details. The strain-energy release rates are then determined for an assumed interlaminar flaw in a structural detail. The flaw is positioned based on finite-element analysis (FEA), and the strain-energy release rates are calculated using the virtual crack closure feature available in codes like ANSYS®. The methodologies have been validated for a skin-stiffener intersection. Two prediction methods differing in complexity and data requirements have been explored. Results for both methods show good agreement between predicted and experimental delamination loads under both static and fatigue loading.

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Liu, Dianzi, Chuanwei Zhang, Z. Wan, and Z. Du. "Topology optimization of a novel fuselage structure in the conceptual design phase." Aircraft Engineering and Aerospace Technology 90, no. 9 (November 14, 2018): 1385–93. http://dx.doi.org/10.1108/aeat-04-2017-0100.

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Purpose In recent years, innovative aircraft designs have been investigated by researchers to address the environmental and economic issues for the purpose of green aviation. To keep air transport competitive and safe, it is necessary to maximize design efficiencies of the aircrafts in terms of weight and cost. The purpose of this paper is to focus on the research which has led to the development of a novel lattice fuselage design of a forward-swept wing aircraft in the conceptual phase by topology optimization technique. Design/methodology/approach In this paper, the fuselage structure is modelled with two different types of elements – 1D beam and 2D shell – for the validation purpose. Then, the finite element analysis coupled with topology optimization is performed to determine the structural layouts indicating the efficient distributed reinforcements. Following that, the optimal fuselage designs are obtained by comparison of the results of 1D and 2D models. Findings The topological results reveal the need for horizontal stiffeners to be concentrated near the upper and lower extremities of the fuselage cross section and a lattice pattern of criss-cross stiffeners should be well-placed along the sides of the fuselage and near the regions of window locations. The slight influence of windows on the optimal reinforcement layout is observed. To form clear criss-cross stiffeners, modelling the fuselage with 1D beam elements is suggested, whereas the less computational time is required for the optimization of the fuselage modelled using 2D shell elements. Originality/value The authors propose a novel lattice fuselage design in use of topology optimization technique as a powerful design tool. Two types of structural elements are examined to obtain the clear reinforcement detailing, which is also in agreement with the design of the DLR (German Aerospace Center) demonstrator. The optimal lattice layout of the stiffeners is distinctive to the conventional semi-monocoque fuselage design and this definitely provides valuable insights into the more efficient utilization of composite materials for novel aircraft designs.

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Stanford, Bret K., and Christine V. Jutte. "Comparison of curvilinear stiffeners and tow steered composites for aeroelastic tailoring of aircraft wings." Computers & Structures 183 (April 2017): 48–60. http://dx.doi.org/10.1016/j.compstruc.2017.01.010.

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47

Barkanov, E., E. Eglītis, F. Almeida, M. C. Bowering, and G. Watson. "Optimal design of composite upper covers of lateral wings with the effect of rib attachment to stiffener webs." Mechanics of Composite Materials 49, no. 3 (July 2013): 285–96. http://dx.doi.org/10.1007/s11029-013-9345-3.

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Silva, Fernando José, Bernardo Félix Santiago Lana, Francisco Carlos Rodrigues, and Luís Eustáquio Moreira. "Buckling of Bamboo Masts with Interposed Spacers." Key Engineering Materials 634 (December 2014): 389–99. http://dx.doi.org/10.4028/www.scientific.net/kem.634.389.

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A hollow bamboo in its raw state, from the geometrical point of view and language engineering can be defined as an element of tubular bar, not prismatic, approximately circular cross sections, stiff by intermittent internal disks positioned along the bar. The decrease in the diameter and wall thickness usually happens from the bottom up, the basal part may contain some exceptions, with sections of the base with diameters smaller than the second, then to diminish steadily to the top. This architecture has a genetic component that resulted from constant interactions of bamboo with the actions of wind, which stimulated increased local resistance of the most requested points mechanically, not only by concentrated lignification in cellulosic tissues, such as the geometric localized variations. From the viewpoint of composite materials science, bamboo can be defined as a composition of two different materials, a first fiber and vessels oriented and aligned along the internodal stem sections, which connect to another material with fibers and vessels tangled - anastonose - intermittent stiffeners composing the above cited nodes. In both materials vessels and fibers are surrounded by a parenchymathous matrix of hollow cells that store sugars. The mechanical point of view, bamboo, due to the configuration, the rigidity and strength components and a tubular geometry is a flexible structural element with high mechanical resistance. This flexibility makes the long elements have low load capacity in flexion compression, if the goal is the application of bamboo in construction structures. This issue can be circumvented by systemic compositions, called masts, [1,2] which can be applied in a single long bamboo element resistant to relatively high compressive loads. In the present investigation, four parallel bamboos 5.5 meters long, the speciesPhyllostachys pubescensare discontinuously connected by bamboo segments interposed fixed by steel pins, achieving this composition with a load limit of 48 kN in controlled experiments. These experiments were also used and motivated to find the numerical modeling by the MEF, whose results were widely discussed.

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Petrolo, Marco, and Erasmo Carrera. "High-Fidelity and Computationally Efficient Component-Wise Structural Models: An Overview of Applications and Perspectives." Applied Mechanics and Materials 828 (March 2016): 175–96. http://dx.doi.org/10.4028/www.scientific.net/amm.828.175.

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The Component-Wise approach (CW) is a novel structural modeling strategy that stemmed from the Carrera Unified Formulation (CUF). This work presents an overview of the enhanced capabilities of the CW for the static and dynamic analysis of structures, such as aircraft wings, civil buildings, and composite plates. The CW makes use of the advanced 1D CUF models. Such models exploit Lagrange polynomial expansions (LE) to model the displacement field above the cross-section of the structure. The use of LE allows the improvement of the 1D model capabilities. LE models provide 3D-like accuracies with far fewer computational costs. The use of LE leads to the CW. Although LE are 1D elements, every component of an engineering structure can be modeled via LE elements independently of their geometry, e.g. 2D transverse stiffeners and panels, and of their scale, e.g. fiber/matrix cells. The use of the same type of finite elements facilitates the finite element modeling to a great extent. For instance, no interface techniques are necessary. Moreover, in a CW model, the displacement unknowns are placed along the physical surfaces of the structure with no need for artificial lines and surfaces. Such a feature is promising in a CAD/FEM coupling scenario. The CW approach can be considered as an accurate and computationally cheap analysis tool for many structural problems. Such as progressive failure analyses, multiscale, impact problems and health-monitoring.

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Qian, Xiuyang, Yushan Zhou, Menghui Wang, Liya Cai, and Feng Pei. "Structural design of composite stiffened panel for a flat wing micro-aircraft." SN Applied Sciences 2, no. 4 (March 20, 2020). http://dx.doi.org/10.1007/s42452-020-2559-9.

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Journal articles on the topic 'Stiffened Composite Wing'

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