The Experts below are selected from a list of 309 Experts worldwide ranked by ideXlab platform
Ronald C Averill - One of the best experts on this subject based on the ideXlab platform.
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Structural Optimization of Composite Aircraft
12th AIAA Aviation Technology Integration and Operations (ATIO) Conference and 14th AIAA ISSMO Multidisciplinary Analysis and Optimization Conference, 2012Co-Authors: Nathan J. Chase, Ronald C Averill, Ranny SidhuAbstract:The aim of this study was to minimize the mass of the structural components in a composite aircraft in order to maximize its payload capacity and flight range. In particular, the detailed lamination schemes in various regions of the fuselage, tail, and wing structural components were optimized using a multi-disciplinary optimization strategy. At every point in the structure, the lamination scheme was defined as an assembly of Sublaminates, each of which spanned different regions of the vehicle. The lamination schemes were automatically generated in terms of the Sublaminate definitions. Thus, the design variables included the number of plies in each Sublaminate and the orientation and material (unidirectional or fabric) in each ply. With hundreds of design variables and a large number of load cases, it was not possible to design an efficient composite aircraft structure using a manual design process and intuition alone. However, by taking advantage of the interactions among the many components, the multi-disciplinary optimization strategy employed here was able to achieve a significant overall mass reduction by making some regions heavier and others lighter.
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An improved thermal lamination model for analysis of heat transfer in composite structures
Journal of Composite Materials, 2002Co-Authors: Antonio Pantano, Ronald C AverillAbstract:A new thermal lamination model and its associated finite element model are presented for analysis of heat transfer in laminated composite structures. The form of the present model closely resembles that of recent zig-zag Sublaminate structural laminate theories. The through-thickness distribution of temperature is assumed to vary linearly within each ply, and continuity of transverse flux at ply interfaces is enforced analytically. Thus, the number of computational degrees-offreedom (DOFs) is made independent of the number of plies in the Sublaminate. In its present form, the model contains only two computational DOFs in each Sublaminate–the temperature at the top and bottom surface of the Sublaminate. This model lends itself well to development of convenient and efficient finite element models, as demonstrated herein. Both linear and nonlinear numerical results are presented to demonstrate the effectiveness of the present approach.
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first order zig zag Sublaminate plate theory and finite element model for laminated composite and sandwich panels
Composite Structures, 2000Co-Authors: Ronald C AverillAbstract:A refined laminated plate theory and three-dimensional finite element based on first-order zig-zag Sublaminate approximations has been developed. The in-plane displacement fields in each Sublaminate are assumed to be piecewise linear functions and vary in a zig-zag fashion through-the-thickness of the Sublaminate. The zig-zag functions are evaluated by enforcing the continuity of transverse shear stresses at layer interfaces. This in-plane displacement field assumption accounts for discrete layer effects without increasing the number of degrees of freedom as the number of layers is increased. The transverse displacement field is assumed to vary linearly through-the-thickness. The transverse normal strain predictions are improved by assuming a constant variation of transverse normal stress in each Sublaminate. In the computational model, each finite element represents one Sublaminate. The finite element is developed with the topology of an eight-noded brick, allowing the thickness of the plate to be discretized into several elements, or Sublaminates, where each Sublaminate can contain more than one physical layer. Each node has five engineering degrees of freedom, three translations and two rotations. Thus, this element can be conveniently implemented into general purpose finite element codes. The element stiffness coefficients are integrated exactly, yet the element exhibits no shear locking due to the use of an interdependent interpolation scheme and consistent shear strain fields. Numerical performance of the current element is investigated for a composite armored vehicle panel and a sandwich panel. These tests demonstrate that the element is very accurate and robust.
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a 3d zig zag Sublaminate model for analysis of thermal stresses in laminated composite and sandwich plates
Journal of Sandwich Structures and Materials, 2000Co-Authors: Antonio Pantano, Ronald C AverillAbstract:A laminated plate theory and 3D finite element model based on first-order zig-zag Sublaminate approximations are presented for thermal stress analysis of composite laminates and sandwich plates. The finite element is developed with the topology of an eight-noded brick, allowing the thickness of the plate to be discretized into several elements, or Sublaminates, where each Sublaminate can contain more than one physical layer. The temperature field is first computed by a thermal model, where the through-thickness distribution of temperature is assumed to vary linearly within each ply, and continuity of transverse flux at ply interfaces is enforced analytically. Similarly, the in-plane displacement fields in each Sublaminate are assumed to be piecewise linear functions and vary in a zig-zag fashion through the thickness of the Sublaminate. The zig-zag functions are evaluated by enforcing the continuity of transverse shear stresses at layer interfaces. The formulation also enforces continuity of the transvers...
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zigzag Sublaminate model for nonlinear analysis of laminated panels
Journal of Aerospace Engineering, 2000Co-Authors: Ronald C AverillAbstract:A geometrically nonlinear first-order zigzag Sublaminate theory and finite-element model are presented that account for moderately large displacements and moderate rotations using a total Lagrangian formulation. The model contains special laminated plate bending kinematics but is cast in the form of a 3D eight-noded brick finite-element topology with five engineering degrees of freedom per node—three translations and two rotations. This permits discretization through the thickness of a laminate to obtain higher accuracy of displacements and stresses when required. The accuracy of the present model is demonstrated by comparing its structural response predictions with results from previous experimental investigations and with numerical tests using a commercial finite-element code.
Biswajit Tripathy - One of the best experts on this subject based on the ideXlab platform.
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Stiffened composite axisymmetric shells-Optimum lay-up for buckling by ranking
Computers and Structures, 1993Co-Authors: Biswajit Tripathy, K.p. RaoAbstract:Linear bifurcation buckling of FRP axisymmetric shells with fully compatible FRP meridional and hoop stiffeners is studied using the finite element method. Eccentricity of the stiffeners is taken into account. The composite shell and the stiffener are assumed to be made of a repeated Sublaminate construction. This type of construction is used in industry to reduce manufacturing errors and to produce more damage-tolerant laminates. In this type of construction, the Sublaminate consists of smaller number of plies and the required thickness of the laminate is obtained by repeating the Sublaminate many times. This paper deals with the determination of the optimum lay-up scheme in the Sublaminate of a composite axisymmetric shell with composite stiffener elements so as to achieve maximum buckling load for a given geometry, loading and boundary conditions using the finite element method. A four-noded, 48-DOF doubly curved quadrilateral laminated anisotropic thin shell finite element with fully compatible two-noded, 16-DOF meridional stiffener elements (MSE) and parallel circle stiffener elements (PCSE) is used. The buckling loads computed for several cases of shells (solid/stiffened) of positive and negative Gaussian curvatures with different applied loads and boundary conditions compare well with existing results in the literature. Subsequently the computer program has been used to find the optimum lay-up scheme of the plies in the Sublaminate so as to achieve maximum buckling load for typical composite solid/stiffened shells. © 1993.
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Optimum design for buckling of plain and stiffened composite axisymmetric shell panels/shells
Composite Structures, 1993Co-Authors: Biswajit TripathyAbstract:The buckling of plain and discretely stiffened composite axisymmetric shell panels/shells made of repeated Sublaminate construction is studied using the finite element method. In repeated Sublaminate construction, a full laminate is obtained by repeating a basic Sublaminate, which has a smaller number of plies. The optimum design for buckling is obtained by determining the layup sequence of the plies in the Sublaminate by ranking, so as to achieve maximum buckling load for a specified thickness. For this purpose, a four-noded 48-dof quadrilateral composite thin shell element, together with fully compatible two-noded 16-dof composite meridional and parallel circle stiffener elements are used.
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Stiffened composite cylindrical panels—Optimum lay-up for buckling by ranking
Computers & Structures, 1992Co-Authors: Biswajit TripathyAbstract:Buckling of discretely stiffened composite cylindrical panels made of repeated Sublaminate construction is studied using a finite element method. In repeated Sublaminate construction, a full laminate is obtained by repeating a basic Sublaminate, which has a smaller number of plies. This paper deals with the determination of the optimum lay-up for buckling by ranking of such stiffened (longitudinal and hoop) composite cylindrical panels. For this purpose we use the particularized form of a four-noded, 48 degrees of freedom doubly curved quadrilateral thin shell finite element together with a fully compatible two-noded, 16 degrees of freedom composite stiffener element. The computer program developed has been used, after extensive checking for correctness, to obtain an optimum orientation scheme of the plies in the Sublaminate so as to achieve maximum buckling load for a specified thickness of typical stiffened composite cylindrical panels.
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Optimum Design for Buckling of Plain and Stiffened Composite Cylindrical Panels
Composite Structures, 1991Co-Authors: Biswajit TripathyAbstract:Buckling of plain and discretely stiffened composite cylindrical panels made of a repeated Sublaminate construction is studied using the finite element method. In repeated Sublaminate construction, a full laminate is obtained by repeating a basic Sublaminate, which has a smaller number of plies. The optimum design for buckling is obtained by determining the layup sequence of the plies in the Sublaminate by ranking, so as to achieve maximum buckling load for a specified thickness. For this purpose, a particularized form of a four- noded, 48-dof quadrilateral composite shell element together with a fully compatible two-noded, 16-dof composite stiffener element are used.
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Curved composite beams—optimum lay-up for buckling by ranking
Computers & Structures, 1991Co-Authors: Biswajit TripathyAbstract:Instability of laminated curved composite beams made of repeated Sublaminate construction is studied using finite element method. In repeated Sublaminate construction, a full laminate is obtained by repeating a basic Sublaminate which has a smaller number of plies. This paper deals with the determination of optimum lay-up for buckling by ranking of such composite curved beams (which may be solid or sandwich). For this purpose, use is made of a two-noded, 16 degress of freedom curved composite beam finite element. The displacements u, v, w of the element reference axis are expressed in terms of one-dimensional first-order Hermite interpolation polynomials, and line member assumptions are invoked in formulation of the elastic stiffness matrix and geometric stiffness matrix. The nonlinear expressions for the strains, occurring in beams subjected to axial, flexural and torsional loads, are incorporated in a general instability analysis. The computer program developed has been used, after extensive checking for correctness, to obtain optimum orientation scheme of the plies in the Sublaminate so as to achieve maximum buckling load for typical curved solid/sandwich composite beams.
Giles W Hunt - One of the best experts on this subject based on the ideXlab platform.
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Compressive strength of composite laminates with delamination-induced interaction of panel and Sublaminate buckling modes
Composite Structures, 2017Co-Authors: Andrew Rhead, Richard Butler, Giles W HuntAbstract:Compression After Impact (CAI) strength is critical to the safety and weight of carbon fibre aircraft. In this paper, the standard aerospace industry practice of using separate analyses and tests for panel buckling and CAI strength is challenged. Composite panels with a range of stacking sequences were artificially delaminated and subject to compression testing in a fixture that allowed local Sublaminate and global panel buckling modes to interact. Compared to panels without delamination, interaction of buckling modes reduced panel buckling strains by up to 29%. Similarly, compared to delaminated panels restrained against panel buckling, interaction reduced delamination propagation strains by up to 49%. These results are the first to indicate that restriction of interaction during CAI testing is unconservative and therefore potentially unsafe. A novel integration of an analytical Strip model, for Sublaminate buckling driven delamination propagation, and a Shanley model, for determining increased local strain due to Sublaminate-buckling-induced panel curvature, is used to calculate the reduction in strength due to buckling mode interaction. Assuming a typical Sublaminate post to pre-buckling stiffness ratio of 0.65, the difference in integrated model and experimental results is
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compressive strength of composite laminates with delamination induced interaction of panel and Sublaminate buckling modes
Composite Structures, 2017Co-Authors: Andrew Rhead, Richard Butler, Giles W HuntAbstract:Compression After Impact (CAI) strength is critical to the safety and weight of carbon fibre aircraft. In this paper, the standard aerospace industry practice of using separate analyses and tests for panel buckling and CAI strength is challenged. Composite panels with a range of stacking sequences were artificially delaminated and subject to compression testing in a fixture that allowed local Sublaminate and global panel buckling modes to interact. Compared to panels without delamination, interaction of buckling modes reduced panel buckling strains by up to 29%. Similarly, compared to delaminated panels restrained against panel buckling, interaction reduced delamination propagation strains by up to 49%. These results are the first to indicate that restriction of interaction during CAI testing is unconservative and therefore potentially unsafe. A novel integration of an analytical Strip model, for Sublaminate buckling driven delamination propagation, and a Shanley model, for determining increased local strain due to Sublaminate-buckling-induced panel curvature, is used to calculate the reduction in strength due to buckling mode interaction. Assuming a typical Sublaminate post to pre-buckling stiffness ratio of 0.65, the difference in integrated model and experimental results is <11% – a level of accuracy that will allow the integrated model to drive initial design studies.
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compressive strength following delamination induced interaction of panel and Sublaminate buckling
53rd AIAA ASME ASCE AHS ASC Structures Structural Dynamics and Materials Conference<BR>20th AIAA ASME AHS Adaptive Structures Conference<BR&g, 2012Co-Authors: Andrew Rhead, Richard Butler, Giles W HuntAbstract:A number of panels have been artificially delaminated to produce a range of thin anisotropic sub-laminates. Panels were subject to experimental compression testing in which Sublaminate and global panel buckling modes were allowed to interact. A comparison of experimental results with previous results where panel buckling was restrained, demonstrates that interaction of buckling modes causes significant reductions in both panel buckling and delamination propagation strains. This indicates that the current industry practice of using separate tests to define allowable strains for panel buckling and damage tolerance may be unconservative. A combination of a Shanley model with a computationally efficient analytical Strip model, conservatively predicts values of compressive strain below which propagation of delaminations will not occur.
Andrew Rhead - One of the best experts on this subject based on the ideXlab platform.
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A plate model for compressive strength prediction of delaminated composites
Composite Structures, 2019Co-Authors: R.s. Choudhry, Andrew Rhead, Mark Nielsen, Richard ButlerAbstract:Abstract Damage tolerance is of critical importance to laminated composite structures. In this paper, we present a new semi-analytical method for predicting the strain at which delamination propagation will initiate following Sublaminate buckling. The method uses a numerical strip model to determine the thin-film buckling strain of an anisotropic sub-laminate created by delamination, before evaluating the strain energy release rate for delamination propagation. The formulation assumes that all energy is available for propagation in a peeling mode (Mode I); avoiding an approximate mixed-mode criterion. Results are compared with twelve experimentally obtained propagations strains, covering a variety of laminates each containing a circular PTFE delamination. Comparison shows agreement to within 12% for balanced Sublaminate tests in which delamination propagation occurred before intra-ply cracking. The method can be used to significantly improve the damage tolerance of laminates, opening up new opportunities for structural efficiency using elastic tailoring, non-standard ply angles and material optimisation.
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Compressive strength of composite laminates with delamination-induced interaction of panel and Sublaminate buckling modes
Composite Structures, 2017Co-Authors: Andrew Rhead, Richard Butler, Giles W HuntAbstract:Compression After Impact (CAI) strength is critical to the safety and weight of carbon fibre aircraft. In this paper, the standard aerospace industry practice of using separate analyses and tests for panel buckling and CAI strength is challenged. Composite panels with a range of stacking sequences were artificially delaminated and subject to compression testing in a fixture that allowed local Sublaminate and global panel buckling modes to interact. Compared to panels without delamination, interaction of buckling modes reduced panel buckling strains by up to 29%. Similarly, compared to delaminated panels restrained against panel buckling, interaction reduced delamination propagation strains by up to 49%. These results are the first to indicate that restriction of interaction during CAI testing is unconservative and therefore potentially unsafe. A novel integration of an analytical Strip model, for Sublaminate buckling driven delamination propagation, and a Shanley model, for determining increased local strain due to Sublaminate-buckling-induced panel curvature, is used to calculate the reduction in strength due to buckling mode interaction. Assuming a typical Sublaminate post to pre-buckling stiffness ratio of 0.65, the difference in integrated model and experimental results is
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compressive strength of composite laminates with delamination induced interaction of panel and Sublaminate buckling modes
Composite Structures, 2017Co-Authors: Andrew Rhead, Richard Butler, Giles W HuntAbstract:Compression After Impact (CAI) strength is critical to the safety and weight of carbon fibre aircraft. In this paper, the standard aerospace industry practice of using separate analyses and tests for panel buckling and CAI strength is challenged. Composite panels with a range of stacking sequences were artificially delaminated and subject to compression testing in a fixture that allowed local Sublaminate and global panel buckling modes to interact. Compared to panels without delamination, interaction of buckling modes reduced panel buckling strains by up to 29%. Similarly, compared to delaminated panels restrained against panel buckling, interaction reduced delamination propagation strains by up to 49%. These results are the first to indicate that restriction of interaction during CAI testing is unconservative and therefore potentially unsafe. A novel integration of an analytical Strip model, for Sublaminate buckling driven delamination propagation, and a Shanley model, for determining increased local strain due to Sublaminate-buckling-induced panel curvature, is used to calculate the reduction in strength due to buckling mode interaction. Assuming a typical Sublaminate post to pre-buckling stiffness ratio of 0.65, the difference in integrated model and experimental results is <11% – a level of accuracy that will allow the integrated model to drive initial design studies.
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The influence of surface ply fibre angle on the compressive strength of composite laminates containing delamination
Aeronautical Journal, 2012Co-Authors: Andrew Rhead, Richard Butler, Neil BakerAbstract:A combination of uniaxial compression tests and Strip Model and Finite Element analyses of laminates artificially delaminated to create circular [±θ] Sublaminates is used to assess the influence of fibre angle on the compressive strength of composite laminates. Sublaminates with 0° < θ < 40° are found to fail by Sublaminate-buckling-driven delamination propagation and provide poor tolerance of delamination. This is a consequence of their relatively high axial stiffnesses, low Sublaminate buckling strains, Poisson’s ratio induced compressive transverse strains and extension-twist coupling which produces unexpected Sublaminate buckling mode shapes. Sublaminates with 40° < θ < 60° are most tolerant to delamination; axial and transverse stiffnesses are minimal, formation of Sublaminate buckles is resisted, high laminate buckling strains reduce interaction between laminate and Sublaminate buckling mode shapes and extension-twist coupling is minimal. Sublaminates with 60° < θ < 90° are shown to produce varied tolerance of delamination. Sublaminate buckling is generally prevented owing to transverse tensile strains induced by mismatches between laminate and Sublaminate Poisson’s ratios but may occur in laminates with low Poisson’s ratios.
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compressive strength following delamination induced interaction of panel and Sublaminate buckling
53rd AIAA ASME ASCE AHS ASC Structures Structural Dynamics and Materials Conference<BR>20th AIAA ASME AHS Adaptive Structures Conference<BR&g, 2012Co-Authors: Andrew Rhead, Richard Butler, Giles W HuntAbstract:A number of panels have been artificially delaminated to produce a range of thin anisotropic sub-laminates. Panels were subject to experimental compression testing in which Sublaminate and global panel buckling modes were allowed to interact. A comparison of experimental results with previous results where panel buckling was restrained, demonstrates that interaction of buckling modes causes significant reductions in both panel buckling and delamination propagation strains. This indicates that the current industry practice of using separate tests to define allowable strains for panel buckling and damage tolerance may be unconservative. A combination of a Shanley model with a computationally efficient analytical Strip model, conservatively predicts values of compressive strain below which propagation of delaminations will not occur.
K.p. Rao - One of the best experts on this subject based on the ideXlab platform.
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Stiffened composite axisymmetric shells-Optimum lay-up for buckling by ranking
Computers and Structures, 1993Co-Authors: Biswajit Tripathy, K.p. RaoAbstract:Linear bifurcation buckling of FRP axisymmetric shells with fully compatible FRP meridional and hoop stiffeners is studied using the finite element method. Eccentricity of the stiffeners is taken into account. The composite shell and the stiffener are assumed to be made of a repeated Sublaminate construction. This type of construction is used in industry to reduce manufacturing errors and to produce more damage-tolerant laminates. In this type of construction, the Sublaminate consists of smaller number of plies and the required thickness of the laminate is obtained by repeating the Sublaminate many times. This paper deals with the determination of the optimum lay-up scheme in the Sublaminate of a composite axisymmetric shell with composite stiffener elements so as to achieve maximum buckling load for a given geometry, loading and boundary conditions using the finite element method. A four-noded, 48-DOF doubly curved quadrilateral laminated anisotropic thin shell finite element with fully compatible two-noded, 16-DOF meridional stiffener elements (MSE) and parallel circle stiffener elements (PCSE) is used. The buckling loads computed for several cases of shells (solid/stiffened) of positive and negative Gaussian curvatures with different applied loads and boundary conditions compare well with existing results in the literature. Subsequently the computer program has been used to find the optimum lay-up scheme of the plies in the Sublaminate so as to achieve maximum buckling load for typical composite solid/stiffened shells. © 1993.
