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.

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.

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.

For the next generation of aircraft, the use of fibre-reinforced polymer composites and the design of

<p>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.</p><p>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.</p><p>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. </p>

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.

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.

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

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.

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.