Design of a laminated composite variable curvature panel under uniaxial compression (original) (raw)
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Failure analysis of curved composite panels based on first-ply and buckling failures
Procedia Engineering, 2011
Curved panels are used extensively in several branches of engineering and in particular in marine and aerospace engineering working mostly under compressive loads. Failure of these components by buckling or excessive stress is an important design consideration. In the present study the effect of fiber orientation is studied on the failure load of a laminated curved panel subject to uniaxial compression. The failure modes are specified as first-ply failure and buckling with the failure load defined as the minimum of these two loads. The panel is taken as a symmetrically laminated angle-ply plate and the failure load is determined for different aspect ratios, panel thicknesses and boundary conditions (simply supported and clamped panels). The failure load is maximized for a set of selected stacking sequences by determining the best ply angle for each stacking sequence giving the highest failure load.
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.
Thin-Walled Structures, 1996
Optimal buckling designs of symmetrically laminated rectangular plates under in-plane uniaxial loads" which have a nonuniform distribution along the edges are presented. In particular, point loads, partial uniJorm loads and nonuniform loads" are considered in addition to uniform O' distributed inplane loads" which provide the benchmark solutions. Poisson's effect is" taken into account when in-plane restraints are present along the unloaded edges. Restraints give rise to in-plane loads" at unloaded edges which lead to biaxial loading, and may cause premature instability. The laminate behaviour with respect to fiber orientation changes significantly in the presence of Poisson's eJfi, ct as compared to that o/'a laminate where this" ~Jfect is neglected. This change in behaviour has significant implications Jor design optimisation as the optimal values of design variables with or without restraints differ substantially. In the present study, the design objective is" the maximisation of the uniaxial buckling load by optimally determining the fiber orientations. The )qnite element method, coupled with an optimisation routine, is employed in analysing and optimising the laminates. Numerical results are given for a number of boundary conditions and fi)r uniJormly and non-uniformly distributed buckling loads. Copyright :~'~ 1996 Elsevier Science Ltd.
Buckling optimization of laminated cylindrical panels subjected to axial compressive load
Composite Structures, 2007
The buckling resistance of fiber-reinforced laminated cylindrical panels with a given material system and subjected to uniaxial compressive force is maximized with respect to fiber orientations by using a sequential linear programming method together with a simple move-limit strategy. The significant influences of panel thicknesses, curvatures, aspect ratios, cutouts and end conditions on the optimal fiber orientations and the associated optimal buckling loads of laminated cylindrical panels have been shown through this investigation.
A Numerical and Experimental Study of Compression-Loaded Composite Panels with Cutouts
47th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference<BR> 14th AIAA/ASME/AHS Adaptive Structures Conference<BR> 7th, 2006
Results from a numerical and experimental study on the effects of laminate orthotropy and circular cutout size on the response of compression-loaded composite curved panels are presented. Several 60-in-radius composite panels with four different laminate configurations were tested with cutout diameters that range from 10% to 60% of the panel width. Finite-element analyses were performed for each panel in order to identify the effects boundary conditions, measured initial geometric imperfections and thickness variations had on the nonlinear and buckling behavior of the panels. The compression-loaded panels considered herein exhibited two separate types of behavior depending on the laminate stacking sequence and cutout size. More specifically, some of the panels exhibited the classical "snap-through" type buckling response; however, some of the panels exhibited a monotonically increasing stable response and achieved compressive loads in excess of twice the predicted linear bifurcation buckling load. In general, the finite-element analyses were able to predict accurately the nonlinear response and buckling loads of the panels and the prebuckling and postbuckling out-of-plane deformations and strains.
BUCKLING ANALYSIS OF STIFFENED COMPOSITE PANELS FOR DIFFERENT PLY ORIENTATIONS
Buckling and post-buckling analysis was performed on composite stiffened panel using Abaqus/CAE to obtain the critical load and modes of failures, with different parameters like ply orientation, different composite materials, stiffeners & by changing the number of stiffeners were derived. To investigate the buckling behavior of composite curved stiffened panels, the nonlinear FE tools Abaqus/Explicit are employed. Studies were conducted using the analytical tool in order to understand the structural behavior in the post buckling range and to determine the critical parameters.
2018
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.
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,
1994
Optimal buckling designs of symmetrically laminated rectangular plates under in-plane uniaxial loads" which have a nonuniform distribution along the edges are presented. In particular, point loads, partial uniJorm loads and nonuniform loads" are considered in addition to uniform O' distributed inplane loads" which provide the benchmark solutions. Poisson's effect is" taken into account when in-plane restraints are present along the unloaded edges. Restraints give rise to in-plane loads" at unloaded edges which lead to biaxial loading, and may cause premature instability. The laminate behaviour with respect to fiber orientation changes significantly in the presence of Poisson's eJfi, ct as compared to that o/'a laminate where this" ~Jfect is neglected. This change in behaviour has significant implications Jor design optimisation as the optimal values of design variables with or without restraints differ substantially. In the present study, the design objective is" the maximisation of the uniaxial buckling load by optimally determining the fiber orientations. The)qnite element method, coupled with an optimisation routine, is employed in analysing and optimising the laminates. Numerical results are given for a number of boundary conditions and fi)r uniJormly and non-uniformly distributed buckling loads.
IRJET, 2020
Composite materials are widely used in aerospace industry due to their high performance and characteristics like strength to weight ratio, stiffness to weight ratio etc. In this paper, finite element analysis for flat rectangular panel is done with different stacking sequences (symmetric) and different ply orientations, the ply orientations considered for the analysis are 0 0 , ±45 0 and 90 0 with a total number of 28 plies. Roark's formula is considered for the calculation of critical buckling stress when the panel is under equal uniform compression on two opposite short edges and is simply supported on all edges. After the analysis, buckling factor and buckling mode shapes are observed for the different stacking sequences and ply orientations. ABD matrix as well as elastic modulus in longitudinal direction E11, modulus in transvers direction E22, Poisson's ratio Nu11 and shear modulus G12 are tabulated for each stacking sequence. From the analysis best stacking sequence for the applied loading condition is summarized.