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Published: 4 June 2021
Last updated: 7 February 2022

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1

Xu, Hailan. "Buckling, Postbuckling and Imperfection Sensitivity Analysis of Different Type of Cylindrical Shells by Hui's Postbuckling Method." ScholarWorks@UNO, 2013. http://scholarworks.uno.edu/td/1781.

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Hui and Chen (1986) were the first to show that the well-known Koiter’s General Theory of Elastic Stability of 1945 can be significantly improved by evaluating the postbuckling b coefficient at the actual applied load, rather than at the classical buckling load. Such improvement method was demonstrated to be (1) very simple to apply with no tedious algebra, (2) significant reduction in imperfection sensitivity and (3) although it is still asymptotically valid, there exists a significant extension of the range of validity involving larger imperfection amplitudes. Strictly speaking, Koiter’s theory of 1945 is valid only for vanishingly small imperfection amplitudes. Hence such improved method is termed Hui’s Postbuckling method. This study deals with the postbuckling and imperfection sensitivity of different kinds of cylinders, using the Hui’s postbuckling method. For unstiffened cylinder and laminate cylinder the results are compared with ABAQUS simulation results, and a parameter variation of stringer/ring stiffened cylinder is also evaluated. A significant positive shift of the postbuckling b coefficient is found which indicates that Koiter's general stability theory of 1945 has significantly overestimated the imperfection sensitivity of the structure. Also, compared with the Koiter's general stability theory, the valid region is significantly increased by using Hui's postbuckling method.
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2

Soncco, K., X. Jorge, and R. A. Arciniega. "Postbuckling Analysis of Functionally Graded Beams." Institute of Physics Publishing, 2019. http://hdl.handle.net/10757/625602.

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This paper studies the geometrically non-linear bending behavior of functionally graded beams subjected to buckling loads using the finite element method. The computational model is based on an improved first-order shear deformation theory for beams with five independent variables. The abstract finite element formulation is derived by means of the principle of virtual work. High-order nodal-spectral interpolation functions were utilized to approximate the field variables which minimizes the locking problem. The incremental/iterative solution technique of Newton's type is implemented to solve the nonlinear equations. The model is verified with benchmark problems available in the literature. The objective is to investigate the effect of volume fraction variation in the response of functionally graded beams made of ceramics and metals. As expected, the results show that transverse deflections vary significantly depending on the ceramic and metal combination.
Revisión por pares
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3

Faggiani, Andrea. "Optimisation of postbuckling stiffened composite structures." Thesis, Imperial College London, 2008. http://hdl.handle.net/10044/1/8001.

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The thesis starts off with an introductory chapter on composite materials. This includes a definition of composites, a brief history of composite materials, their use in aerostructures (primarily as stiffened structures), and also optimization of composite structures. A literature review is then presented on postbuckling stiffened structures. This includes both experimental investigations on stiffened composite panels and investigations into secondary instabilities and mode jumping as well as their numerical modelling. Next, the Finite Element (FE) modelling of posthuckling stiffened structures is discussed, relating how ABAQUS models are set up in order to trace stiffened composite panels' buckling and postbuckling responses. An experimental programme conducted on an I-stiffened panel is described, where the panel was tested in compression until collapse. The buckling and postbuckling characteristics of the panel are presented, and then an FE model is described together with its predicted numerical behaviour of the panel's buckling and postbuckling characteristics. Focus then shifts to the modelling of failure in composites, in particular delamination failure. A literature review is conducted, looking at the use of both the Virtual Crack Closure Technique (VCCT) and interface elements in delamination modelling. Two stiffener runout models, representing two specimens previously tested experimentally, are then developed to illustrate how interface elements may be used to model mixed mode delamination. The previously discussed panel is revisited, and a global-local modelling approach used to model the skin-stiffener interface. FE models of a stiffened cylindrical shell are also considered, and again the postbuckling characteristics of the shell are compared with experimental results. . The thesis then moves on to optimization of composite structures. This starts off with a literature review of existing optimization methodologies. A Genetic Algorithm (GA) is devised to increase the damage resistance of the I-stiffened panel. The global-local ABAQUS model discussed earlier is used in conjunction with the GA in order to find a revised stacking sequence of both the panel flanges and skin so as to minimize skin-stiffener debonding subject to a variety of design constraints. A second optimization is then presented, this time linked to the FE model of the stiffened cylindrical shell. The objective is to increase the collapse load of the shell, again subject to specific design constraints. The thesis concludes by summarising the importance of the work conducted. FE models were created and validated against experimental work in order to model a variety of composite stiffened structures in their buckling and postbuckling regimes. These models were able to capture the failure characteristics of these structures relating to delamination at the skin-stiffener interface, a phenomenon widely observed experimentally. Various optimizations, able to account for failure mechanisms which may occur prior to overall structural collapse, were then conducted on the analysed structures in order to obtain more damage resistant designs.
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4

Song, Yuzhan. "Thermo-elastoviscoplastic postbuckling behavior of shell-like structures." Diss., Georgia Institute of Technology, 1991. http://hdl.handle.net/1853/11703.

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5

Parsa, Kourosh. "Buckling/postbuckling of polymer composite continuum/skeletal structures." Thesis, University of Surrey, 1995. http://epubs.surrey.ac.uk/636/.

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6

Lee, Ho Hyung. "Postbuckling failure of composite plates with central holes." Diss., This resource online, 1991. http://scholar.lib.vt.edu/theses/available/etd-10022007-145159/.

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7

Seresta, Omprakash. "Buckling, Flutter, and Postbuckling Optimization of Composite Structures." Diss., Virginia Tech, 2007. http://hdl.handle.net/10919/26401.

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This research work deals with the design and optimization of a large composite structure. In design of large structural systems, it is customary to divide the problem into many smaller independent/semi-independent local design problems. For example, the wing structure design problem is decomposed into several local panel design problem. The use of composite necessitates the inclusion of ply angles as design variables. These design variables are discrete in nature because of manufacturing constraint. The multilevel approach results into a nonblended solution with no continuity of laminate layups across the panels. The nonblended solution is not desirable because of two reasons. First, the structural integrity of the whole system is questionable. Second, even if there is continuity to some extent, the manufacturing process ends up being costlier. In this work, we develop a global local design methodology to design blended composite laminates across the whole structural system. The blending constraint is imposed via a guide based approach within the genetic algorithm optimization scheme. Two different blending schemes are investigated, outer and inner blending. The global local approach is implemented for a complex composite wing structure design problem, which is known to have a strong global local coupling. To reduce the computational cost, the originally proposed local one dimensional search is replaced by an intuitive local improvement operator. The local panels design problem arises in global/local wing structure design has a straight edge boundary condition. A postbuckling analysis module is developed for such panels with applied edge displacements. A parametric study of the effects of flexural and inplane stiffnesses on the design of composite laminates for optimal postbuckling performance is done. The design optimization of composite laminates for postbuckling strength is properly formulated with stacking sequence as design variables. Next, we formulate the stacking sequence design (fiber orientation angle of the layers) of laminated composite flat panels for maximum supersonic flutter speed and maximum thermal buckling capacity. The design is constrained so that the behavior of the panel in the vicinity of the flutter boundary should be limited to stable limit cycle oscillation. A parametric study is carried out to investigate the tradeoff between designs for thermal buckling and flutter. In an effort to include the postbuckling constraint into the multilevel design optimization of large composite structure, an alternative cheap methodology for predicting load paths in postbuckled structure is presented. This approach being computationally less expensive compared to full scale nonlinear analysis can be used in conjunction with an optimizer for preliminary design of large composite structure with postbuckling constraint. This approach assumes that the postbuckled stiffness of the structure, though reduced considerably, remains linear. The analytical expressions for postbuckled stiffness are given in a form that can be used with any commercially available linear finite element solver. Using the developed approximate load path prediction scheme, a global local design approach is developed to design large composite structure with blending and local postbuckling constraints. The methodology is demonstrated via a composite wing box design with blended laminates.
Ph. D.
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8

Chung, Kwok Fai. "The elastic distortional and local plate buckling of slender web beam." Thesis, Imperial College London, 1988. http://hdl.handle.net/10044/1/7860.

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9

Jane, Kuo Chang. "Buckling, postbuckling deformation and vibration of a delaminated plate." Diss., Georgia Institute of Technology, 1989. http://hdl.handle.net/1853/19975.

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10

Cerini, Marco. "Investigation of secondary instabilities in postbuckling stiffened composite structures." Thesis, Imperial College London, 2006. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.429557.

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11

Lee, Jaehong. "Vibration, buckling and postbuckling of laminated composites with delaminations." Diss., This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-06062008-170322/.

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12

Meyers, Carol Ann. "Thermal buckling and postbuckling of symmetrically laminated composite plates." Thesis, Virginia Tech, 1990. http://hdl.handle.net/10919/42238.

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This paper discusses an investigation into thermal buckling and post-buckling of symmetrically laminated composite plates. In this study, thermal buckling is investigated for laminates under two different simple support conditions, fixed and sliding. These laminates are subjected to the conditions of a uniform temperature change and a linearly varying temperature change along the length of the plate. Postbuckling in the presence of a uniform temperature change and nonlinear response to imperfections in the form of a thermal gradient through the thickness of the plate and a lack of initial flatness are also studied. The buckling response is studied using variational methods, specifically the Trefftz criterion. Postbuckling and responses to imperfections are studied using nonlinear equilibrium conditions. A Rayleigh-Ritz formulation is used to obtain numerical results from the formulations for the prebuckling response, the buckling response, and the post-buckling and imperfection responses. The analyses are applied to graphite-reinforced materials with (± 45/0₂)s and (± 45/0/90)s lamination sequences. Numerical results are obtained for these laminates and also for the case of these laminates being rotated 30° inplane. For the first laminate, for example, such a rotation results in a (+75/ — 15/30₂)s. stacking sequence. Such skewing of the principal material directions may be encountered when using fiber-reinforced materials in a structurally tailored design. In addition, the influence on thermal buckling of a lack of ideal boundary conditions in the form of boundary compliance and thermal expansion, which would occur in any real set-up, are investigated.
Master of Science
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13

Qu, Shuang. "Multilevel optimisation of aerospace and lightweight structures incorporating postbuckling effects." Thesis, Cardiff University, 2011. http://orca.cf.ac.uk/55080/.

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The optimisation of aerospace structures is a very complex problem, due to the hundreds of design variables a multidisciplinary optimisation may contain, so that multilevel optimisation is required. This thesis presents the recent developments to the multilevel optimisation software VICONOPT MLO, which is a multilevel optimisation interface between the well established analysis and design software packages VICONOPT and MSC/NASTRAN. The software developed is called VICONOPT MLOP (Multilevel Optimisation with Postbuckling), and allows for postbuckling behaviour, using analysis based on the Wittrick-Williams algorithm. The objective of this research is to enable a more detailed insight into the multilevel optimisation and postbuckling behaviour of a complex structure. In VICONOPT MLOP optimisation problems, individual panels of the structural model are allowed to buckle before the design load is reached. These panels continue to carry load with differing levels of reduced stiffness. VICONOPT MLOP creates new MSC/NASTRAN data files based on this reduced stiffness data and iterates through analysis cycles to converge on an appropriate load re-distribution. Once load convergence has been obtained with an appropriate criterion, the converged load distribution is used as a starting point in the optimisation of the constituent panels, i.e. a new design cycle is started, in which the updated ply thicknesses for each panel are calculated by VICONOPT and returned to MSC/NASTRAN through VICONOPT MLOP. Further finite element analysis of the whole structure is then carried out to determine the new stress distributions in each panel. The whole process is repeated until a mass convergence criterion is met. A detailed overview of the functionality of VICONOPT MLOP is presented in the thesis. A case study is conducted into the multilevel optimisation of a composite aircraft wing, to demonstrate the capabilities of VICONOPT MLOP and identify areas for future studies. The results of the case study show substantial mass savings, proving the software's capabilities when dealing with such problems. The time taken for this multilevel optimisation also proves the efficiency of the software.
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14

Riddick, Jaret Cleveland. "Prebuckling, Buckling, and Postbuckling Response of Segmented Circular Composite Cylinders." Diss., Virginia Tech, 2001. http://hdl.handle.net/10919/29952.

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Discussed is a numerical and experimental characterization of the response of small-scale fiber-reinforced composite cylinders constructed to represent a fuselage design whereby the crown and keel consist of one laminate stacking sequence and the two sides consist of another laminate stacking sequence. This construction is referred to as a segmented cylinder. The response to uniform axial endshortening is discussed. Numerical solutions for the nonlinear prebuckling, buckling, and postbuckling responses are compared to experimental results. Focus is directed at the investigation of two specific cylinder configurations, referred to as axially-stiff and circumferentially-stiff cylinders. Small-scale cylinders, each having a nominal radius of 5 in., were fabricated on a mandrel by splicing adjacent segments together to form 0.5 in. overlaps. Finite-element models of both cylinder configurations, including the overlap regions, are developed using the STAGS finite-element code. Perfectly circular cylinder models are considered, as are models which include the measured geometry of the specimens as an imperfection. Prebuckling predictions show that the segmented cylinder response is characterized by the existence of circumferential displacement, and an axial boundary layer accompanied by circumferential gradients in radial displacement. Experimental measurements, taken with strain gages and displacement transducers, confirm these numerical findings. As the endshortening approaches the critical, or buckling, values, the response of the cylinders is characterized by wrinkling in the axial direction. In the axially-stiff cylinder, the crown and keel segments wrinkle, while in the circumferentially-stiff cylinder the side segments wrinkle. Experimental images taken from Moire interferometry show this response in the circumferentially-stiff cylinder. Four methods are used to predict the buckling values of endshortening and load for both cylinders, and the four values are in good agreement. The experimentally-measured buckling conditions, however, show that the models overpredict buckling values. For the axially-stiff cylinder, the difference could be due to the fact material failure not included in the model plays a role in the cylinder response. For the circumferentially-stiff cylinder, the difference is definitely due to material failure characteristics not included in the model. The predicted postbuckling response of the segmented cylinders is shown to be dominated by the existence of inward dimples in some or all of the segments. For the axially-stiff cylinder, the as-predicted dimpled crown and keel configuration is observed in the experiment but at a load 12 percent below predicted values. For the circumferentially-stiff cylinder material failure in the linear prebuckling range of response triggered buckling that resembled the predicted circumferential rings of dimples, but at a load 31 percent below predictions. Finally, it is shown that the effect of including the measured imperfections in the model has little observable effect on the circumferentially-stiff cylinder. For the axially-stiff cylinder the inclusion of the imperfections is found to effect the transition from buckling to postbuckling, but ultimately has little effect on postbuckling deformations.
Ph. D.
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15

Wolfe, David R. "Delamination buckling, postbuckling, and growth in axially loaded beam-plates." Thesis, Virginia Polytechnic Institute and State University, 1987. http://hdl.handle.net/10919/80114.

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The purpose of this study is to develop a simple one-dimensional model to analyze axially loaded beam-plates containing cracks which extend through the thickness of the beam-plates. Although the material analyzed is isotropic, these cracks will be referred to as delaminations. Buckling, postbuckling, and growth of delaminations in these beam-plates will be analyzed. A finite element method in which all of the terms of the stiffness matrices are obtained by exact integration is employed to determine the linear buckling load and postbuckling solution. The energy release rate is then determined using the postbuckling solution. Curves are provided to show the effect of delamination length and location on buckling loads, energy release rates, and strengths of the beam-plates. The problem of buckling and postbuckling of beams with multiple delaminations is also considered. A method of calculating the energy release rate for beams with multiple delaminations using numerical differentiation is introduced.
Master of Science
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16

Perry, Christine Ann. "Minimum-weight design of compressively loaded stiffened panels for postbuckling response." Thesis, This resource online, 1995. http://scholar.lib.vt.edu/theses/available/etd-02132009-172332/.

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17

DiNardo, Marc Thomas. "Buckling and postbuckling behavior of laminated composite plates with ply dropoffs." Thesis, Massachusetts Institute of Technology, 1986. http://hdl.handle.net/1721.1/14997.

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Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1986.
MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERO
Bibliography: leaves 320-324.
by Marc Thomas DiNardo.
M.S.
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18

Powell, Stephen. "Buckling and postbuckling of prismatic plate assemblies using exact eigenvalue theory." Thesis, Cardiff University, 1997. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.691256.

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19

Jensen, David Warren. "Buckling and postbuckling behavior of unbalanced and unsymmetric laminated graphite/epoxy plates." Thesis, Massachusetts Institute of Technology, 1986. http://hdl.handle.net/1721.1/15102.

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Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1986.
MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERO
Bibliography: v.1, leaves 287-291.
by David Warren Jensen.
Ph.D.
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20

Hause, Terry J. "Thermomechanical Postbuckling of Geometrically Imperfect Anisotropic Flat and Doubly Curved Sandwich Panels." Diss., Virginia Tech, 1998. http://hdl.handle.net/10919/30449.

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Sandwich structures constitute basic components of advanced supersonic/hypersonic flight and launch vehicles. These advanced flight vehicles operate in hostile environments consisting of high temperature, moisture, and pressure fields. As a result, these structures are exposed to large lateral pressures, large compressive edge loads, and high temperature gradients which can create large stresses and strains within the structure and can produce the instability of the structure. This creates the need for a better understanding of the behavior of these structures under these complex loading conditions. Moreover, a better understanding of the load carrying capacity of sandwich structures constitutes an essential step towards a more rational design and exploitation of these constructions. In order to address these issues, a comprehensive geometrically non-linear theory of doubly curved sandwich structures constructed of anisotropic laminated face sheets with an orthotropic core under various loadings for simply supported edge conditions is developed. The effects of the radii of curvature, initial geometric imperfections, pressure, uniaxial compressive edge loads, biaxial edge loading consisting of compressive/tensile edge loads, and thermal loads will be analyzed. The effect of the structural tailoring of the facesheets upon the load carrying capacity of the structure under these various loading conditions are analyzed. In addition, the movability/immovability of the unloaded edges and the end-shortening are examined. To pursue this study, two different formulations of the theory are developed. One of these formulations is referred to as the mixed formulation, While the second formulation is referred to as the displacement formulation. Several results are presented encompassing buckling, postbuckling, and stress/strain analysis in conjunction with the application of the structural tailoring technique. The great effects of this technique are explored. Moreover, comparisons with the available theoretical and experimental results are presented and good agreements are reported.
Ph. D.
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21

Namdar, Omer. "Buckling, Postbuckling And Progressive Failure Analyses Of Composite Laminated Plates Under Compressive Loading." Master's thesis, METU, 2012. http://etd.lib.metu.edu.tr/upload/12614924/index.pdf.

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The aim of this thesis is to investigate buckling, post-buckling behaviors and failure characteristics of composite laminated plates under compressive loading with the help of finite element method and experiments. In the finite element analyses, eigen value extraction method is used to determine the critical buckling loads and nonlinear Riks and Newton-Raphson methods are employed to obtain post-buckling behaviors and failure loads. The effects of geometric imperfection amplitude on buckling and post-buckling are discussed. Buckling load, post buckling loaddisplacement relations, out of plane displacements and end shortening of the plates are determined numerically. Furthermore, the numerical results are compared with experimental findings for two different laminates made of woven fabric and unidirectional tapes where buckling, post-buckling behavior and structural failure of laminated plates were determined. The comparisons show that there is a good agreement between numerical and experimental results obtained for buckling load and post-buckling range. However, 15 % - 22 % differences are predicted between the experimental and numerical results for failure of laminates made of woven fabric whereas the laminates with uni-directional tapes show good agreement.
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22

Ferrie, Catherine H. "Effect of transverse shear on the postbuckling and growth characteristics of delaminated composites." Diss., Georgia Institute of Technology, 1997. http://hdl.handle.net/1853/12355.

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23

Young, Richard Douglas. "Prebuckling and postbuckling behavior of stiffened composite panels with axial-shear stiffness coupling." Diss., This resource online, 1996. http://scholar.lib.vt.edu/theses/available/etd-06062008-144734/.

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24

Wieland, Todd M. "Scale effects in buckling, postbuckling and crippling of graphite-epoxy Z-section stiffeners." Diss., Virginia Tech, 1991. http://hdl.handle.net/10919/39973.

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25

Che, Bin. "Improved exact strip postbuckling analysis of anisotropic plate with combined load and edge cases." Thesis, Cardiff University, 2012. http://orca.cf.ac.uk/56759/.

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Minimisation of the mass of aerospace structures has been investigated by researchers and designers for many years. It is an efficient means to reduce the manufacturing costs, fuel consumption and environmental impact. To achieve this objective, high performance composite materials and optimised configurations are utilised in modern aircraft design. Additionally, use of the postbuckling reserve of strength has been considered during the preliminary design stage to obtain more efficient structures. The exact strip analysis and optimum design software VICONOPT has been developed and used in postbuckling analysis. VICONOPT is able to give a good initial evaluation of load versus end shortening when compared with experimental and finite element results. However it provides poor predictions of the stress and strain distributions in the postbuckling range. This is due to its assumptions concerning the longitudinal invariance of stress and the sinusoidal variation of buckling modes in the longitudinal direction. These assumptions are appropriate for initial buckling analysis but they limit the accuracy of subsequent postbuckling analysis. This thesis outlines some developments which improve the existing exact strip postbuckling analysis by improving the accuracy of mode shape prediction and stress and strain distributions. Based on previous research by Von Karman, improved governing equations are derived and solved for general anisotropic plates with different in- plane edge conditions. Implementation of the improved analysis in VICONOPT enhances the accuracy of mode shapes and stress and strain distributions in the postbuckling analysis.
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26

Lynch, Colum James. "A finite element study of the postbuckling behaviour of a typical aircraft fuselage panel." Thesis, Queen's University Belfast, 2000. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.343003.

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27

Kim, Yoon Duk. "Transverse Stiffener Requirements in Straight and Horizontally Curved Steel I-Girders." Thesis, Georgia Institute of Technology, 2004. http://hdl.handle.net/1853/4802.

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Recent research studies have confirmed that curved I-girders are capable of developing substantial shear postbuckling resistance due to tension field action and have demonstrated that the AASHTO LRFD equations for the tension field resistance in straight I-girders may be applied to curved I-girders within specific limits. However, the corresponding demands on intermediate transverse stiffeners in curved I-girders are still largely unknown. Furthermore, a number of prior research studies have demonstrated that transverse stiffeners in straight I-girders are loaded predominantly by bending induced by their restraint of web lateral deflections at the shear strength limit state, not by in-plane tension field forces. This is at odds with present Specification approaches for the design of transverse stiffeners, which are based on (1) providing sufficient stiffener bending rigidity only to develop the shear buckling strength of the web and (2) providing sufficient stiffener area to resist the in-plane tension field forces. In this research, the behavior of one- and two-sided intermediate transverse stiffeners in straight and horizontally curved steel I-girders is investigated by refined full nonlinear finite element analysis. Variations in stiffener rigidity, panel aspect ratio, panel slenderness, and stiffener type are considered. New recommendations for design of transverse stiffeners in straight and curved I-girder bridges are developed by combining the solutions from the above FEA studies with the results from prior research.
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28

Breivik, Nicole L. "Thermal and Mechanical Response of Curved Composite Panels." Diss., Virginia Tech, 1998. http://hdl.handle.net/10919/28015.

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Curved panels constructed of laminated graphite-epoxy composite material are of potential interest in airframe fuselage applications. An understanding of structural response at elevated temperatures is required for anticipated future high speed aircraft applications. This study concentrates on the response of unstiffened, curved composite panels subjected to combinations of thermal and mechanical loading conditions. Mechanical loading is due to compressive end-shortening and thermal loading is due to a uniform temperature increase. Thermal stresses, which are induced by mechanical restraints against thermal expansions or contractions, cause buckling and postbuckling panel responses. Panels with three different lamination sequences are considered, including a quasi-isotropic laminate, an axially soft laminate, and an axially stiff laminate. These panels were chosen because they exhibit a range of stiffnesses and a wide variation in laminate coefficients of thermal expansion. The panels have dimensions of 10 in. by 10 in. with a base radius of 60 in. The base boundary conditions are clamped along the curved ends, and simply supported along the straight edges. Three methods are employed to study the panel response, including a geometrically nonlinear Rayleigh-Ritz solution, a finite element solution using the commercially available code STAGS, and an experimental program. The effects of inplane boundary conditions and radius of curvature are studied analytically, along with consideration of order of application in combined loading. A substantial difference is noted in the nonlinear load vs. axial strain responses of panels loaded in end-shortening and panels loaded with uniform temperature change, depending on the specific lamination sequence, boundary conditions, and radius of curvature. Experiments are conducted and results are presented for both room temperature end-shortening tests and elevated temperature tests with accompanying end-shortening. The base finite element model is modified to include measured panel thicknesses, boundary conditions representative of the experimental apparatus, measured initial geometric imperfections, and measured temperature gradients. With these modifications, and including an inherent end displacement of the panel present during thermal loading, good correlation is obtained between the experimental and numerically predicted load vs. axial strain responses from initial loading through postbuckling.
Ph. D.
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29

Ragon, Scott Alan II. "Development of a Global/Local Approach and a Geometrically Non-linear Local Panel Analysis for Structural Design." Diss., Virginia Tech, 1998. http://hdl.handle.net/10919/30761.

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A computationally efficient analysis capability for the geometrically non-linear response of compressively loaded prismatic plate structures was developed. Both a "full" finite strip solution procedure and a "reduced" solution procedure were implemented in a FORTRAN 90 computer code, and comparisons were made with results available in the technical literature. Both the full and reduced solution procedures were demonstrated to provide accurate results for displacement and strain quantities through moderately large post-buckling loads. The full method is a non-linear finite strip analysis of the semi-analytical, multi-term type. Individual finite strips are modeled as balanced and symmetric laminated composite materials which are assumed to behave orthotropically in bending, and the structure is loaded in uniaxial or biaxial compression. The loaded ends of the structure are assumed to be simply supported, and geometric shape imperfections may be modeled. The reduced solution method makes use of a reduced basis technique in conjunction with the full finite strip analysis. Here, the potentially large set of non-linear algebraic equations produced by the finite strip method are replaced by a small set of system equations. In the present implementation, the basis vectors consist of successive derivatives of the non-linear solution vector with respect to a loading parameter. Depending on the nature of the problem, the reduced solution procedure is capable of computational savings of up to 60%+ compared to the full finite strip method. The reduced method is most effective in reducing the computational cost of the full method when the most significant portion of the cost of the full method is factorization of the assembled system matrices. The robustness and efficiency of the reduced solution procedure was found to be sensitive to the user specified error norm which is used during the reduced solution procedure to determine when to generate new sets of basis vectors. In parallel with this effort, a new method for performing global/local design optimization of large complex structures (such as aircraft wings or fuselages) was developed. A simple and flexible interface between the global and local design levels was constructed using response surface methodology. The interface is constructed so as to minimize the changes required in either the global design code or the local design codes(s). Proper coupling is maintained between the global and local design levels via a "weight constraint" and the transfer of global stiffness information to the local level. The method was verified using a simple isotropic global wing model and the local panel design code PASCO.
Ph. D.
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30

Shin, Dong Ku. "Minimum-weight design of symmetrically laminated composite plates for postbuckling performance under in-plane compression loads." Diss., This resource online, 1990. http://scholar.lib.vt.edu/theses/available/etd-07282008-135134/.

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31

Lin, Weiqing. "Buckling and postbuckling of flat and curved laminated composite panels under thermomechanical loadings incorporating non-classical effects." Diss., Virginia Tech, 1997. http://hdl.handle.net/10919/40240.

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Two structural models which can be used to predict the buckling, post buckling and vibration behavior of flat and curved composite panels under thermomechanical loadings are developed in this work. Both models are based on higher-order transverse shear deformation theories of shallow shells that include the effects of geometric nonlinearities and initial geometric imperfections. Within the first model (Model I), the kinematic continuity at the contact surfaces between the contiguous layers and the free shear traction condition on the outer bounding surfaces are satisfied, whereas in the second model (Model II), in addition to these conditions, the static interlaminae continuity requirement is also fulfilled. Based on the two models, results which cover a variety of problems concerning the postbuckling behaviors of flat and curved composite panels are obtained and displayed. These problems include: i) buckling and postbuckling behavior of flat and curved laminated structures subjected to mechanical and thermal loadings; ii)frequency-load/temperature interaction in laminated structures in both pre-buckling and post buckling range; iii) the influence of a linear/nonlinear elastic foundation on static and dynamic post buckling behavior of flat/curved laminated structures exposed to mechanical and temperature fields; iv) implication of edge constraints upon the temperature/load carrying capacity and frequencyload/ temperature interaction of flat/curved structures; v) elaboration of a number of methodologies enabling one to attenuate the intensity of the snap-through buckling and even to suppress it as well as of appropriate ways enabling one to enhance the load/temperature carrying capacity of structures.
Ph. D.
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32

Herrero, Javier. "Buckling, postbuckling and progressive failure analysis of hybrid composite shear webs using a continuum damage mechanics model." Diss., Wichita State University, 2007. http://hdl.handle.net/10057/1488.

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This dissertation presents an innovative analysis methodology to enhance the design of composite structures by extending their work range into the postbuckling regime. This objective is accomplished by using the numerical simulation capabilities of nonlinear finite element analysis combined with continuum damage mechanics models to simulate the onset of failure and the subsequent material properties degradation. A complete analysis methodology is presented with increasing levels of complexity. The methodology is validated by correlation of analytical results with experimental data from a set of hybrid carbon/epoxy glass/epoxy composite panels tested under shear loading using a picture frame fixture.
Thesis (Ph.D.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering
"December 2007."
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33

Herrero, Javier Locke James Yang Charles. "Buckling, postbuckling and progressive failure analysis of hybrid composite shear webs using a continuum damage mechanics model /." Diss., A link to full text of this thesis in SOAR, 2007. http://hdl.handle.net/10057/1488.

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34

Chai, Gin Boay. "The effect of geometry and prescribed delaminations on the postbuckling behaviour of laminated carbon-fibre reinforced plastic panels." Thesis, University of Strathclyde, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.303278.

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35

Lee, Merrill Cheng Wei Mechanical &amp Manufacturing Engineering Faculty of Engineering UNSW. "Stochastic analysis and robust design of stiffened composite structures." Awarded By:University of New South Wales. Mechanical & Manufacturing Engineering, 2009. http://handle.unsw.edu.au/1959.4/44217.

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The European Commission 6th Framework Project COCOMAT (Improved MATerial Exploitation at Safe Design of COmposite Airframe Structures by Accurate Simulation of COllapse) was a four and a half year project (2004 to mid-2008) aimed at exploiting the large reserve of strength in composite structures through more accurate prediction of collapse. In the experimental work packages, significant statistical variation in buckling behaviour and ultimate loading were encountered. The variations observed in the experimental results were not predicted in the finite element analyses that were done in the early stages of the project. The work undertaken in this thesis to support the COCOMAT project was initiated when it was recognised that there was a gap in knowledge about the effect of initial defects and variations in the input variables of both the experimental and simulated panels. The work involved the development of stochastic algorithms to relate variations in boundary conditions, material properties and geometries to the variation in buckling modes and loads up to first failure. It was proposed in this thesis that any future design had to focus on the dominant parameters affecting the statistical scatter in the results to achieve lower sensitivity to variation. A methodology was developed for designing stiffened composite panels with improved robustness. Several panels tested in the COCOMAT project were redesigned using this approach to demonstrate its applicability. The original contributions from this thesis are therefore the development of a stochastic methodology to identify the impact of variation in input parameters on the response of stiffened composite panels and the development of Robust Indices to support the design of new panels. The stochastic analysis included the generation of metamodels that allow quantification of the impact that the inputs have on the response using two first order variables, Influence and Sensitivity. These variables are then used to derive the Robust Indices. A significant outcome of this thesis was the recognition in the final report for COCOMAT that the development of a validated robust index should be a focus of any future design of postbuckling stiffened panels.
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36

Haynie, Waddy. "Torsion of Elliptical Composite Cylindrical Shells." Diss., Virginia Tech, 2007. http://hdl.handle.net/10919/28547.

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The response of elliptical composite cylindrical shells under torsion is studied. The torsional condition is developed by rotating one end of the cylinder relative to the other. Prebuckling, buckling, and postbuckling responses are examined, and material failure is considered. Four elliptical cross sections, defined by their aspect ratio, the ratio of minor to major radii, are considered: 1.00 (circular), 0.85, 0.70, and 0.55. Two overall cylinder sizes are studied; a small size with a radius and length for the circular cylinder of 4.28 in. and 12.85 in., respectively, and a large size with radii and lengths five times larger, and thicknesses two times larger than the small cylinders. The radii of the elliptical cylinders are determined so the circumference is the same for all cylinders of a given size. For each elliptical cylinder, two lengths are considered. One length is equal to the length of the circular cylinder, and the other length has a sensitivity of the buckling twist to changes in the length-to-radius ratio the same as the circular cylinder. A quasi-isotropic lamination sequence of a medium-modulus graphite-epoxy composite material is assumed. The STAGS finite element code is used to obtain numerical results. The geometrically-nonlinear static and transient, eigenvalue, and progressive failure analysis options in the code are employed. Generally, the buckling twist and resulting torque decrease with decreasing aspect ratio. Due to material anisotropy, the buckling values are generally smaller for a negative twist than a positive twist. Relative to the buckling torque, cylinders with aspect ratios of 1.00 and 0.85 show little or no increase in capacity in the postbuckling range, while cylinders with aspect ratios of 0.70 and 0.55 show an increase. Postbuckling shapes are characterized by wave-like deformations, with ridges and valleys forming a helical pattern due to the nature of loading. The amplitudes of the deformations are dependent on cross-sectional geometry. Some elliptical cylinders develop wave-like deformations prior to buckling. Instabilities in the postbuckling range result in shape changes and loss of torque capacity. Material failure occurs on ridges and in valleys. Cylinder size and cross-sectional geometry influence the initiation and progression of failure.
Ph. D.
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37

Lindström, Anders. "Strength of Sandwich Panels Loaded in In-plane Compression." Licentiate thesis, KTH, Aeronautical and Vehicle Engineering, 2007. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-4558.

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The use of composite materials in vehicle structures could reduce the weight and thereby the fuel consumption of vehicles.

As the road safety of the vehicles must be ensured, it is vital that the energy absorbing capability of the composite materials are similar to or better than the commonly used steel structures. The high specific bending stiffness of sandwich structures can with advantage be used in vehicles, provided that the structural behaviour during a crash situation is well understood and possible to predict. The purpose of this thesis is to identify and if possible to describe the failure initiation and progression in in-plane compression loaded sandwich panels.

An experimental study on in-plane compression loaded sandwich panels with two different material concepts was conducted. Digital speckle photography (DSP) was used to record the displacement field of one outer face-sheet surface during compression. The sandwich panels with glass fibre preimpregnated face-sheets and a polymer foam core failed due to disintegration of the face-sheets from the core, whereas the sandwich panels with sheet molding compound face-sheets and a balsa core failed in progressive end-crushing. A simple semi-empirical model was developed to describe the structural response before and after initial failure.

The postfailure behaviour of in-plane compression loaded sandwich panels was studied by considering the structural behaviour of sandwich panels with edge debonds. A parametrical finite element model was used to determine the influence of different material and geometrical properties on the buckling and postbuckling failure loads. The postbuckling failure modes studied were debond crack propagation and face-sheet failure. It could be concluded that the postbuckling failure modes were mainly determined by the ratio between the fracture toughness of the face-core interface and the bending stiffness of the face-sheets.

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38

Hart, Daniel Constantine. "Development of a Progressive Failure Finite Element Analysis For a Braided Composite Fuselage Frame." Thesis, Virginia Tech, 2002. http://hdl.handle.net/10919/34026.

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Short, J-section columns fabricated from a textile composite are tested in axial compression to study the modes of failure with and without local buckling occuring.The textile preform architecture is a 2x2, 2-D triaxial braid with a yarn layup of [0 deg 18k/+-64 deg 6k] 39.7% axial. The preform was resin transfer molded with 3M PR500 epoxy resin. Finite element analyses (FEA) of the test specimens are conducted to assess intra- and inter- laminar progressive failure models. These progressive failure models are then implemented in a FEA of a circular fuselage frame of the same cross section and material for which test data was available. This circular frame test article had a nominal radius of 120 inches, a forty-eight degree included angle, and was subjected to a quasi-static, radially inward load, which represented a crash type loading of the frame. The short column test specimens were cut from some of the fuselage frames. The branched shell finite element model of the frame included geometric nonlinearity and contact of the load platen of the testing machine with the frame. Intralaminar progressive failure is based on a maximum in-plane stress failure criterion followed by a moduli degradation scheme. Interlaminar progressive failure was implemented using an interface finite element to model delamination initiation and the progression of delamination cracks. Inclusion of both the intra- and inter- laminar progressive failure models in the FEA of the frame correlated reasonably well with the load-displacement response from the test through several major failure events.
Master of Science
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39

Wang, Yang. "Virtual testing of post-buckling behaviour of metallic stiffened panel." Thesis, Cranfield University, 2011. http://dspace.lib.cranfield.ac.uk/handle/1826/7291.

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The aim of the project presented in this thesis is to demonstrate a modelling method for predicting the variability in the ultimate load of stiffened panel under axial compression due to manufacturing variability. Bulking is sensitive to imperfections. In the case of a post-buckled panel, manu-facturing variability produces a scatter in the ultimate load. Thus, reasonable leeway for imperfections and inherent variability must be allowed in their design. Firstly, a finite element model of a particular stiffened panel was developed, and all nonlinearities within the material, boundary condition and geometry were considered. Verification and validation were performed to examine the accuracy of the buckling behaviour prediction, especially ultimate load. Experiments on 5 identical panels in design were performed to determine the level of panel-panel variation in geometry and collapse load. A data reduction programme based on the practical geometry scanning was developed, in addi-tion to which, the procedure of importing measured imperfection into Finite Ele-ment model was introduced. To identify and apply representative imperfections to the panel model, a double Fourier series representation of the random geometric distributions is attempt-ed, and was used thereby to derive a series of shapes representing random ge-ometry scatters. With these newly generated geometric imperfections, the variation in collapse load was determined, using the validated FE analysis. And also, the probability of these predicted loads was generalized.
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40

Byklum, Eirik. "Ultimate strength analysis of stiffened steel and aluminium panels using semi-analytical methods." Doctoral thesis, Norwegian University of Science and Technology, Department of Marine Technology, 2002. http://urn.kb.se/resolve?urn=urn:nbn:no:ntnu:diva-352.

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Buckling and postbuckling of plates and stiffened panels are considered. Computational models for direct calculation of the response are developed using large deflection plate theory and energy principles. Deflections are represented by trigonometric functions. All combinations of biaxial in-plane compression or tension, shear, and lateral pressure are included in the formulations. The procedure is semi-analytical in the sense that the incremental equilibrium equations are derived analytically, while a numerical method is used for solving the equation systems, and for incrementation of the solution.

Unstiffened plate models are developed both for the simply supported case and for the clamped case. For the simply supported case the material types considered are isotropic elastic, orthotropic elastic, and elastic-plastic. Two models are developed for analysis of local buckling of stiffened plates, one for open profiles and one for closed profiles. A global buckling model for stiffened panels is developed by considering the panel as a plate with general anisotropic stiffness. The stiffness coefficients are input from the local analysis. Two models are developed for combined local and global buckling, in order to account for interaction between local and global deflection. The first is for a single stiffened plate, and uses a column approach. The second is for a stiffened panel with several stiffeners.

Numerical results are calculated for a variety of plate and stiffener geometries for verification of the proposed model, and comparison is made with nonlinear finite element methods. Some examples are presented. For all models, the response in the elastic region is well predicted compared with the finite element method results. Also, the efficiency of the calculations is very high. Estimates of ultimate strength are found using first yield as a collapse criterion. In most cases, this leads to conservative results compared to predictions from finite element calculations.

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41

Obdržálek, Vít. "Boulení delaminovaných kompozitních desek." Doctoral thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2010. http://www.nusl.cz/ntk/nusl-233929.

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Chování laminátových desek namáhaných na tlak či na smyk může být výrazně ovlivněno přítomností delaminací, tedy oblastí, kde je porušena vazba mezi sousedními vrstvami. Cílem této práce je rozšířit znalosti o chování delaminovaných desek, a to především o chování desek s větším počtem delaminací a desek s delaminacemi libovolného tvaru, neboť taková podoba porušení laminátu více odpovídá poškození vznikajícího v důsledku nízkorychlostního dopadu cizího tělesa na laminátovou desku. Disertační práce se skládá ze tří hlavních částí. V první části jsou stručně nastíněny postupy využívané při analýze boulení delaminovaných desek a jsou diskutována omezení těchto analýz. Dále jsou v této části shrnuty hlavní poznatky o boulení delaminovaných desek. V druhé části práce je popsán výpočtový model použitý v rámci disertační práce pro analýzu boulení delaminovaných desek. Schopnost modelu předpovědět chování delaminovaných desek je pak dokumentována na několika ověřovacích úlohách. Třetí část disertační práce se skládá ze tří samostatných studií chování desek s několika delaminacemi eliptického či kruhového tvaru a jedné studie zabývající se možností náhrady obecného tvaru delaminace kruhem či elipsou. Je probírán vliv řady parametrů na chování delaminovaných desek, konkrétně vliv orientace vrstev laminátu a dále vliv počtu, tvaru, orientace a umístění delaminací. Na základě těchto studií jsou pak zformulována doporučení ohledně postupu při posuzování únosnosti delaminovaných konstrukcí.
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42

Elseifi, Mohamed A. "A new scheme for the optimum design of stiffened composite panels with geometric imperfections." Diss., Virginia Tech, 1998. http://hdl.handle.net/10919/29250.

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Thin walled stiffened composite panels, which are among the most utilized structural elements in engineering, possess the unfortunate property of being highly sensitive to geometrical imperfections. Existing analysis codes are able to predict the nonlinear postbuckling behavior of a structure with specified imperfections. However, it is impossible to determine the geometric imperfection profile of a nonexistent composite panel early in the design. This is due to the variety of uncertainties that are involved in the manufacturing of these panels. As a mater of fact, due to the very nature of the manufacturing processes, it is hard to imagine that a given manufacturing process could ever produce two identical panels. The objective of this study is to introduce a new design methodology in which a manufacturing model and a convex model for uncertainties are used in conjunction with a nonlinear design tool in order to obtain a more realistic, better performing final design. First a finite element code for the nonlinear postbuckling analysis of stiffened panels is introduced. Next, a manufacturing model for the simulation of the autoclave curing of epoxy matrix composites is presented. A convex model for the uncertainties in the imperfections is developed in order to predict the weakest panel profile among a family of panels. Finally, the previously developed tools are linked in a closed loop design scheme aimed at obtaining a final design that incorporates the manufacturing tolerances information through more realistic imperfections.
Ph. D.
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43

Orifici, Adrian Cirino, and adrian orifici@student rmit edu au. "Degradation Models for the Collapse Analysis of Composite Aerospace Structures." RMIT University. Aerospace, Mechanical and Manufacturing Engineering, 2007. http://adt.lib.rmit.edu.au/adt/public/adt-VIT20080619.090039.

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44

Ungwattanapanit, Tanut [Verfasser], Horst [Akademischer Betreuer] Baier, Horst [Gutachter] Baier, and Kai-Uwe [Gutachter] Bletzinger. "Optimization of Steered-Fibers Composite Stiffened Panels including Postbuckling Constraints handled via Equivalent Static Loads / Tanut Ungwattanapanit ; Gutachter: Horst Baier, Kai-Uwe Bletzinger ; Betreuer: Horst Baier." München : Universitätsbibliothek der TU München, 2017. http://d-nb.info/1152384082/34.

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45

Ferreira, Inês Oliveira de Vasconcelos. "Analysis of the structural behaviour of stiffened panels subjected to compressive loading conditions." Master's thesis, Universidade de Aveiro, 2015. http://hdl.handle.net/10773/16557.

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Mestrado em Engenharia Mecânica
Stiffened panels form the basic structural building blocks of airplanes, vessels and other structures with high requirements of strength-to-weight ratio. As a consequence it is crucial to understand the behaviour of these type of panels. Since buckling is the primary mode of failure of stiffened panels, it will be the focus in the present work. In the present work it was carried out several analysis, using the simulation software Abaqus, in order to study the buckling and postbuckling behaviour. Two different panels were tested in this thesis, the first one an aluminium stiffened panel, which its main goal was to understand the methodologies involved in the analysis of the buckling behaviour, and the second one a composite stiffened which its main goal was to find the proper tools to simulate its behaviour. Therefore, two different methods were used, the Riks method was used to analyse the aluminium panel and the Stabilize method to analyse the composite panel. The behaviour of stiffened panels are influenced by several parameters such as, the number and type of elements, the skin-stringer connection, the boundary conditions, the magnitude of imperfections, etc. So in the present work, those parameters were taken into account and its influence will be shown.
Os painéis reforçados formam as estruturas básicas de construção de aviões, navios e outras estruturas que exijam uma elevada relação entre resistência e peso. Deste modo, é crucial perceber o comportamento deste tipo de painéis. Tendo em conta que a encurvadura é o modo principal de falha deste tipo de painéis, será o foco de estudo desta dissertação. No trabalho presente, foram realizadas várias análises de forma a estudar o comportamento de encurvadura e pós-encurvadura de painéis reforçados, utilizando para isso o software de simulação Abaqus. Foram testados dois painéis diferentes, sendo que o primeiro foi um painél de alumínio, com o objectivo de perceber as metodologias envolvidas na simulação de placas reforçadas, e o segundo, um painel compósito, com o objecto de encontrar as ferramentas adequadas para simular o seu comportamento. Para isso, dois métodos distintos foram utilizados, sendo que foi utilizado o método de Riks para analisar a placa de alumínio e para analisar a placa compósita foi ultizado o método de estabilização. O comportamento dos painéis reforçados é influênciado por vários parâmetros tais como, modelo numérico, ligação entre placa e reforço, condições de fronteira, magnitude de imperfeições, etc. Assim, todos esses parâmetros foram tidos em conta e a sua influência irá ser mostrada no trabalho presente.
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46

Sahin, Mehmet. "Buckling and postbuckling behavior of cracked structures /." Diss., 2004. http://gateway.proquest.com/openurl?url_ver=Z39.88-2004&rft_val_fmt=info:ofi/fmt:kev:mtx:dissertation&res_dat=xri:pqdiss&rft_dat=xri:pqdiss:3147332.

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47

Lee, Da-Ching, and 李大慶. "Thermal Postbuckling of Doubly Curved Composite Shell." Thesis, 1996. http://ndltd.ncl.edu.tw/handle/05812895510900001633.

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48

Guo, Feng-Yu, and 郭鳳玉. "A parametric study on postbuckling of delaminated beams." Thesis, 2007. http://ndltd.ncl.edu.tw/handle/89198093849261074563.

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碩士
國立臺灣海洋大學
河海工程學系
95
Abstract Laminated beams possess great strength-to-weight ratio and good resistance to corrosion, and therefore have been widely used in a variety of engineering fields. The major problem encountered in the application of such composite structure is attributable to separation of adjoining plies commonly called delamination. This drawback of material largely arises from manufacturing imperfection or impact loading, and inevitable leads to reduction in the stiffness and resistance strength. Buckling and postbuckling of belaminated beams are investigated in this work to accurately predict the load capacity. On the basis of geometrically nonlinear finite element method and perturbation technique, postbuckling equilibrium paths are obtained as a number of asymptotic expansions of a gauge parameter. Several support conditions are considered for numerical studies.
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T''sai, Jhong-Jing, and 蔡仲景. "The Postbuckling Behavior of Beams on Elastic Foundation." Thesis, 1994. http://ndltd.ncl.edu.tw/handle/81418224662649639494.

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50

LI, ZHENG-LONG, and 李政隆. "Nonlinear solution methods and postbuckling behaviour of structures." Thesis, 1992. http://ndltd.ncl.edu.tw/handle/17828899485650256574.

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