A Numerical and Experimental Study of Compression-Loaded Composite Panels with Cutouts (original) (raw)
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Buckling Analysis of Laminated Composite Panel with Elliptical Cutout Subject to Axial Compression
Modelling and Simulation in Engineering, 2012
A buckling analysis has been carried out to investigate the response of laminated composite cylindrical panel with an elliptical cutout subject to axial loading. The numerical analysis was performed using the Abaqus finite-element software. The effect of the location and size of the cutout and also the composite ply angle on the buckling load of laminated composite cylindrical panel is investigated. Finally, simple equations, in the form of a buckling load reduction factor, were presented by using the least square regression method. The results give useful information into designing a laminated composite cylindrical panel, which can be used to improve the load capacity of cylindrical panels.
Journal of Mechanics of Materials and Structures, 2009
It is well known [1] that non-closely stiffened panels can have considerable postbuckling reserve strength, enabling them to carry loads significantly in excess of their initial buckling load. If appropriately designed, their load carrying capacity will even appreciably exceed that corresponding to an equivalent weight unstiffened shell (i.e. a shell of identical radius and thicker skin and which is also more sensitive to geometrical imperfections). In these shells, initial buckling of their panels takes place in a local mode, i.e. skin buckling between stiffeners, and not in an overall mode, i.e., an Euler or wide column mode. The design of aerospace structures places great emphasis on exploiting the behavior and on mass minimization of such panels to reduce lifecycle costs. An optimum (minimum mass) design approach based on initial buckling, stress or strain, and stiffness constraints, typically yields an idealized structural configuration characterized by almost equal critical loads for local and overall buckling. This, of course, results in little postbuckling strength capacity and susceptibility to premature failure. However, an alternative optimum design approach can be imposed to achieve lower mass designs for a given loading by requiring the initial local buckling to occur considerably below the design load and allowing for the response characteristics known to exist in postbuckled panels [2] ,i.e. capability to carry loads higher than their initial buckling load. To meet the requirements of low structurally weight, advanced lightweight laminated composite elements are increasingly being introduced into new designs of modern aerospace structures for enhancing both their structural efficiency and performance. In recognition of the numerous advantages that composites offer, there is a steady growth in replacement of metallic components by composite ones in marine structures, ground transportation, robotics, sports and other fields of engineering. Many theoretical and experimental studies have been performed on buckling and postbuckling behavior of flat stiffened composite panels (see for example Refs.3-8). Recently, a wide body of description and detailed data on buckling and postbuckling tests has been compiled [9] (see chaps.12-14). However, studies on cylindrical composite shells and curved stiffened composite panels are still quite scarce (see for example Refs.10-15). Most of them have been discussed in detail in Ref. 9 (see chap. 14). In light of the above discussion, it has been suggested that permitting postbuckling to take place under ultimate load of fuselage structures, i.e. alleviation of design constraints, may provide a means for meeting the objectives for the design of next generation aircraft, where the demand is reduction of weight without prejudice to cost and structural life (see paper Vision 2020 of the European Community). This approach has been undertaken in an experimental study (Improved POstbuckling SImulation for Design of Fibre COmposite Stiffened Fuselage Structures -POSICOSS project) as a part of an ongoing effort on design of low cost low weight airborne structures initiated by the 5 th European Initiative Program. It was aimed at supporting the development of improved, fast and reliable procedures for analysis and simulation of postbuckling behavior of fiber composite stiffened panel of future generation fuselage structures and their design. Within the POSICOSS project, the Technion performed a long test series, on curved laminated composite stringer stiffened panels under axial compression, shear load introduced by torsion and combined axial compression and shear. The buckling and postbuckling behavior of these panels was recorded till their final collapse. The first part of this test series, dealing with panels PSC1-PSC9 was summarized in Ref. . The results of the tests with panels BOX1-BOX4, which deal with two identical panels, combined together by two flat non-stiffened aluminum panels, to form a torsion box,
Design of a laminated composite variable curvature panel under uniaxial compression
Engineering Computations, 2012
Purpose-The purpose of this paper is to present the optimal design of a simply supported variable curvature laminated angle-ply composite panel under uniaxial compression. The objective is to maximize the failure load which is defined as the minimum of the buckling load and the first-ply failure load. Design/methodology/approach-The numerical results presented are obtained using a shear deformable degenerated shell finite element, a brief formulation of which is given. Some verification problems are solved and a convergence study is conducted in order to assess the accuracy of the element. The design procedure is presented and optimization results are given for a simply supported symmetric eight layer angle-ply panel composed of a flat and two cylindrical sections. Findings-The influences of the stacking sequence and panel thickness on optimization are investigated and the effects of various problem parameters on the optimization procedure are discussed. Originality/value-The paper shows that the load carrying capacity of thicker panels is considerably reduced when the first-ply failure constraint is taken into account.
Progressive failure analyses of compression-loaded composite curved panels with and without cutouts
Composite Structures, 2004
Progressive failure analyses results are presented for composite curved panels with and without a circular cutout and subjected to axial compression loading well into their postbuckling regime. Ply damage modes such as matrix cracking, fiber-matrix shear, and fiber failure are modeled by degrading the material properties. Results from finite element analyses are compared with experimental data. Good agreement between experimental data and numerical results are observed for most part of the loading range for the structural configurations considered. Modeling of initial geometric imperfections may be required to obtain accurate analysis results depending on the ratio of the cutout width to panel width.
Dynamic Stability of Laminated Composite Curved Panels with Cutouts
Journal of Engineering Mechanics, 2003
The present investigation deals with the dynamic stability behavior of laminated composite curved panels with cutouts subjected to in-plane static and periodic compressive loads, analyzed using the finite element method. A generalized shear deformable Sanders' theory with tracers is used in this study. Numerical results obtained for vibration and buckling of composite panels with cutouts compare well with literature. The principal dynamic instability region of composite perforated panels is obtained using Bolotin's approach. The study reveals that curved panels with cutouts depict higher stiffness with the addition of curvatures. The laminated hyperbolic paraboloid panel shows the highest stiffness with the onset of instability at higher excitation frequencies. The effect of curvature in laminated composite curved panels is reduced with an increase in size of the cutout. The principal instability regions are influenced by the lamination parameters. Thus, the laminate construction, coupled with cutout geometry, can be used to the advantages of tailoring during design of composite structures for practical applications.
Buckling analysis of composite panels
Composite Structures, 2006
The aim of this paper is to present the buckling analysis of a laboratory tested composite panel under axial compression by means of a simple shell finite element that is developed and presented herein. The tests were performed in the Aircraft Structure Laboratory of the Faculty of Aerospace Engineering at the Technion. Buckling is achieved via incremental geometrically nonlinear analysis and monitoring of the tangent stiffness matrix at each increment. The performance of the finite element is further validated by solving a complex multisnap example from the literature.
Effect of rotational restraints on the stability of curved composite panels under shear loading
Acta Mechanica, 2020
The present work focuses on the effect of rotational restraints on the shear buckling of symmetrically laminated curved composite panels. The Sanders-Koiter shell theory and a first-order shear deformation scheme were used for the mathematical representation of the deformation kinematics of cylindrical shells. The eigenvalue buckling equations were obtained using the principle of minimum total potential energy and by employing the Ritz method. The solution for the deformed shape was approximated as a series of trigonometric functions compatible with the essential boundary conditions of the problem. The effect of the rotational and torsional springs was incorporated by adding their corresponding potential energy to the total potential energy of the panel-loading system. Using the developed formulation, the effect of the influential parameters such as aspect ratio, panel curvature and restraint stiffness on the buckling strength of a specific class of laminates was studied extensively. To present the results in a more insightful manner, non-dimensional parameters were used in parametric studies. To normalize the effect of torsional elements, a new non-dimensional parameter was introduced.
Buckling of laminated composite cylindrical skew panels
Experimental studies were made on isotropic cylindrical skew panels made of Aluminum 7075-T6 and laminated composite cylindrical skew panels under uniaxial compression. The experimental values of the critical buckling load (P cr) were determined using five different methods. The values of P cr were also determined using MSC/Nastran and CQUAD8 finite element. The experimental values of the P cr obtained by different methods were compared with the finite element solution. The effects of the skew angle and aspect ratio on the critical buckling load of isotropic cylindrical skew panels made of Aluminum 7075-T6 were studied. The effects of the skew angle, aspect ratio, and the laminate stacking sequence on the critical buckling load of laminated composite cylindrical skew panels were also studied. It is found that the method IV (based on a plot of applied load (P) vs. average axial strain) yields the highest value for P cr and method III (based on a plot of P vs. square of out-of-plane-deflection) the lowest value for P cr. The experimental values given by method IV are seen to be closest to the finite element solution, the discrepancy being in the range of 5–23% for laminated composite cylindrical skew panels. For isotropic panels, it is found that the value P cr initially increases with an increase in the skew angle and later decreases as the skew angle increases beyond 15. For laminated composite panels, the P cr value decreases as the aspect ratio increases for all laminate stacking sequences.