Dissertations / Theses on the topic 'Stiffened composite panel'

Composite materials have been more and more used in aircraft primary structures such as wing and fuselage. The aim of this thesis is to identify an effective way to optimize composite wing structure, especially the stiffened skin panels for minimum weight subject to stiffness and strength constraints. Many design variables (geometrical dimensions, ply angle proportion and stacking sequence) are involved in the optimum design of a composite stiffened panel. Moreover, in order to meet practical design, manufacturability and maintainability requirements should be taken into account as well, which makes the optimum design problem more complicated. In this thesis, the research work consists of three steps: Firstly, attention is paid to metallic stiffened panels. Based on the study of Emero’s optimum design method and buckling analysis, a VB program IPO, which employs closed form equations to obtain buckling load, is developed to facilitate the optimization process. The IPO extends the application of Emero’s method to an optimum solution based on user defined panel dimensional range to satisfy practical design constraints. Secondly, the optimum design of a composite stiffened panel is studied. Based on the research of laminate layup effects on buckling load and case study of bucking analysis methods, a practical laminate database (PLDB) concept is presented, upon which the optimum design procedure is established. By employing the PLDB, laminate equivalent modulus and closed form equations, a VB program CPO is developed to achieve the optimum design of a composite stiffened panel. A multi-level and step-length-adjustable optimization strategy is applied in CPO, which makes the optimization process efficient and effective. Lastly, a composite outer wing box, which is related to the author’s GDP work, is optimized by CPO. Both theoretical and practical optimum solutions are obtained and the results are validated by FE analysis.

Large-aspect-ratio wings and composite structures both have been considered for the next-generation civil transport aircraft to achieve improved aerodynamic efficiency and to save aircraft structural weight. The use of the large-aspect-ratio and the light-weight composite wing can lead to an enhanced flexibility of the aircraft wing, which may cause many aeroelastic problems such as large deflections, increased drag, onset of flutter, loss of control authority, etc. Aeroelastic tailoring, internal structural layout design and aerodynamic wing shape morphing are all considered to address these aeroelastic problems through multidisciplinary design, analysis and optimization (MDAO) studies in this work. Performance Adaptive Aeroelastic Wing (PAAW) program was initiated by NASA to leverage the flexibility associated with the use of the large-aspect-ratio wings and light-weight composite structures in a beneficial way for civil transport aircraft wing design. The biologically inspired SpaRibs concept is used for aircraft wing box internal structural layout design to achieve the optimal stiffness distribution to improve the aircraft performance. Along with the use of the active aeroelastic wing concept through morphing wing shape including the wing jig-shape, the control surface rotations and the aeroelastic tailoring scheme using composite laminates with ply-drop for wing skin design, a MDAO framework, which has the capabilities in total structural weight minimization, total drag minimization during cruise, ground roll distance minimization in takeoff and load alleviation in various maneuver loads by morphing its shape, is developed for designing models used in the PAAW program. A bilevel programming (BLP) multidisciplinary design optimization (MDO) architecture is developed for the MDAO framework. The upper-level optimization problem entails minimization of weight, drag and ground roll distance, all subjected to both static constraints and the global dynamic requirements including flutter mode and free vibration modes due to the specified control law design for body freedom flutter suppression and static margin constraint. The lower-level optimization is conducted to minimize the total drag by morphing wing shape, to minimize wing root bending moment by scheduling flap rotations (a surrogate for weight reduction), and to minimize the takeoff ground roll distance. Particle swarm optimization and gradient-based optimization are used, respectively, in the upper-level and the lower-level optimization problems. Optimization results show that the wing box with SpaRibs can further improve the aircraft performances, especially in a large weight saving, as compared to the wing with traditional spars and ribs. Additionally, the nonuniform chord control surface associated with the wing with SpaRibs achieve further reductions in structural weight, total drag and takeoff ground roll distance for an improved aircraft performance. For a further improvement of the global wing skin panel design, an efficient finite element approach is developed in designing stiffened composite panels with arbitrarily shaped stiffeners for buckling and vibration analyses. The developed approach allows the finite element nodes for the stiffeners and panels not to coincide at the panel-stiffeners interfaces. The stiffness, mass and geometric stiffness matrices for the stiffeners can be transformed to those for the panel through the displacement compatibility at their interfaces. The method improves the feasible model used in shape optimizing by avoiding repeated meshing for stiffened plate. Also, it reduces the order of the finite element model, a fine mesh typically associated with the skin panel stiffened by many stiffeners, for an efficient structural analysis. Several benchmark cases have been studied to verify the accuracy of the developed approach for stiffened composite panel structural analyses. Several parametric studies are conducted to show the influence of stiffener shape/placement/depth-ratio on panel's buckling and vibration responses. The developed approach shows a potential benefit of using gradient-based optimization for stiffener shape design.
Ph. D.

The structural optimization of a cantilever aircraft wing with curvilinear spars and ribs and stiffeners is described. The design concept of reinforcing the wing structure using curvilinear stiffening members has been explored due to the development of novel manufacturing technologies like electron-beam-free-form-fabrication (EBF3). For the optimization of a complex wing, a common strategy is to divide the optimization procedure into two subsystems: the global wing optimization which optimizes the geometry of spars, ribs and wing skins; and the local panel optimization which optimizes the design variables of local panels bordered by spars and ribs. The stiffeners are placed on the local panels to increase the stiffness and buckling resistance. The panel thickness, size and shape of stiffeners are optimized to minimize the structural weight. The geometry of spars and ribs greatly influences the design of stiffened panels. During the local panel optimization, the stress information is taken from the global model as a displacement boundary condition on the panel edges using the so-called "Global-Local Approach". The aircraft design is characterized by multiple disciplines: structures, aeroelasticity and buckling. Particle swarm optimization is used in the integration of global/local optimization to optimize the SpaRibs. The interaction between the global wing optimization and the local panel optimization is usually computationally expensive. A parallel computing technology has been developed in Python programming to reduce the CPU time. The license cycle-check method and memory self-adjustment method are two approaches that have been applied in the parallel framework in order to optimize the use of the resources by reducing the license and memory limitations and making the code robust. The integrated global-local optimization approach has been applied to subsonic NASA common research model (CRM) wing, which proves the methodology's application scaling with medium fidelity FEM analysis. Both the global wing design variables and local panel design variables are optimized to minimize the wing weight at an acceptable computational cost. The structural weight of the wing has been, therefore, reduced by 40% and the parallel implementation allowed a reduction in the CPU time by 89%. The aforementioned Global-Local Approach is investigated and applied to a composite panel with crack at its center. Because of composite laminates' heterogeneity, an accurate analysis of these requires very high time and storage space. In the presence of structural discontinuities like cracks, delaminations, cutouts etc., the computational complexity increases significantly. A possible alternative to reduce the computational complexity is the global-local analysis which involves an approximate analysis of the whole structure followed by a detailed analysis of a significantly smaller region of interest. We investigate here the performance of the global-local scheme based on the finite element method by comparing it to the traditional finite element method. To do so, we conduct a 2D structural analysis of a composite square plate, with a thin rectangular notch at its center, subjected to a uniform transverse pressure, using the commercial software ABAQUS. We show that the presence of the thin notch affects only the local response of the structure and that the size of the affected area depends on the notch length. We investigate also the effect of the notch shape on the response of the structure. Stiffeners attached to composite panels may significantly increase the overall buckling load of the resultant stiffened structure. Buckling analysis of a composite panel with attached longitudinal stiffeners under compressive loads is performed using Ritz method with trigonometric functions. Results are then compared to those from ABAQUS FEA for different shell elements. The case of composite panel with one, two, and three stiffeners is investigated. The effect of the distance between the stiffeners on the buckling load is also studied. The variation of the buckling load and buckling modes with the stiffeners' height is investigated. It is shown that there is an optimum value of stiffeners' height beyond which the structural response of the stiffened panel is not improved and the buckling load does not increase. Furthermore, there exist different critical values of stiffener's height at which the buckling mode of the structure changes. Next, buckling analysis of a composite panel with two straight stiffeners and a crack at the center is performed. Finally, buckling analysis of a composite panel with curvilinear stiffeners and a crack at the center is also conducted. ABAQUS is used for these two examples and results show that panels with a larger crack have a reduced buckling load. It is shown also that the buckling load decreases slightly when using higher order 2D shell FEM elements. A damage tolerance framework, EBF3PanelOpt, has been developed to design and analyze curvilinearly stiffened panels. The framework is written with the scripting language PYTHON and it interacts with the commercial software MSC. Patran (for geometry and mesh creation), MSC. Nastran (for finite element analysis), and MSC. Marc (for damage tolerance analysis). The crack location is set to the location of the maximum value of the major principal stress while its orientation is set normal to the major principal axis direction. The effective stress intensity factor is calculated using the Virtual Crack Closure Technique and compared to the fracture toughness of the material in order to decide whether the crack will expand or not. The ratio of these two quantities is used as a constraint, along with the buckling factor, Kreisselmeier and Steinhauser criteria, and crippling factor. The EBF3PanelOpt framework is integrated within a two-step Particle Swarm Optimization in order to minimize the weight of the panel while satisfying the aforementioned constraints and using all the shape and thickness parameters as design variables. The result of the PSO is used then as an initial guess for the Gradient Based Optimization using only the thickness parameters as design variables. The GBO is applied using the commercial software VisualDOC.
Ph. D.

Thesis (M.S.)--Wichita State University, College of Engineering.
"May 2006." Title from PDF title page (viewed on October 29, 2006). Thesis adviser: Hamid M. Lankarani. Includes bibliographic references (leaves 57-59).

Rollovers tend to be very severe crashes because of the energy required to roll a vehicle over unlike front and side crashes. This study is an effort towards reducing the severity of a rollover crash by strengthening the roof of a passenger car. The main focus of this thesis is to study the effect of reinforcing the roof of a car, in the event of a rollover. An Eglass/polypropylene isogrid composite panel, which is known for high specific energy absorption under impact, is used in reinforcing the roof of a ford Taurus car and the force-displacement response of the roof structure is observed in contrast to the same while without the roof reinforcement. The non-linear Finite Element Analysis (FEA) simulation of this rollover event is performed in LS DYNA computer code. The simulation setup is done in accordance with Federal Motor Vehicle Standard (FMVSS) No. 216, which is a static study of roof strength in the case of rollover accidents. A study of roof strength characteristics under dynamic loading, involving rollover forces and velocities, is also carried out. The simulation in the latter case involves dropping the car, inverted at a certain pitch and roll offset against gravity from a certain predetermined height onto a concrete surface. The results of the above simulations show that Isogrid composite panels are excellent reinforcements for a car roof, stiffening the roof and increasing its resistance to crush in a rollover accident. This study will help in recommending the use of isogrid panels in the design of car roofs with better roof crush characteristics.
Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Mechanical Engineering
"May 2006."
Includes bibliographic references (leaves 57-59)

In this Thesis, the problem of the Free Vibrations Analysis of Doubly Stiffened Composite, Orthotropic and/or Isotropic, Base Plates or Panels (with Orthotropic Stiffening Plate Strips) is investigated. The composite plate or panel system is made of an Orthotropic and/or Isotropic Base Plate stiffened or reinforced by adhesively bonded Upper and Lower Orthotropic Stiffening Plate Strips. The plates are assumed to be the Mindlin Plates connected by relatively very thin adhesive layers. The general problem under study is considered in terms of three problems, namely Main PROBLEM I Main PROBLEM II and Main PROBLEM III. The theoretical formulation of the Main PROBLEMS is based on a First Order Shear Deformation Plate Theory (FSDPT) that is, in this case, the Mindlin Plate Theory. The entire composite system is assumed to have simple supports along the two opposite edges so that the Classical Levy'
s Solutions can be applied in that direction. Thus, the transverse shear deformations and the rotary moments of inertia of plates are included in the formulation. The very thin, yet elastic deformable adhesive layers are considered as continua with transverse normal and shear stresses. The damping effects in the plates and the adhesive layers are neglected. The sets of the systems of equations of the Mindlin Plate Theory are reduced to a set of the Governing System of First Order Ordinary Differential Equations in the state vector form. The sets of the Governing System for each Main PROBLEM constitute a Two-Point Boundary Value Problem in the y-direction which is taken along the length of the plates. Then, the system is solved by the Modified Transfer Matrix Method (with Interpolation Polynomials and/or Chebyshev Polynomials)which is a relatively semi-analytical and numerical technique. The numerical results and important parametric studies of the natural modes and the corresponding frequencies of the composite system are presented.

A method for computing the skin-stiffener interface stresses in stiffened composite panels is developed. Both geometrically linear and nonlinear analyses are considered. Particular attention is given to the flange termination region where stresses are expected to exhibit unbounded characteristics. The method is based on a finite-element analysis and an elasticity solution. The finite-element analysis is standard, while the elasticity solution is based on an eigenvalue expansion of the stress functions. The eigenvalue expansion is assumed to be valid in the local flange termination region and is coupled with the finite-element analysis using collocation of stresses on the local region boundaries. In the first part of the investigation the accuracy and convergence of the local elasticity solution are assessed using a geometrically linear analysis. It is found that the finite-element/local elasticity solution scheme produce a very accurate interface stress representation in the local flange termination region. The use of 10 to 15 eigenvalues, in the eigenvalue expansion series, and 100 collocation points results in a converged local elasticity solution. In the second part of the investigation, the local elasticity solution is extended to include geometric nonlinearities. Using this analysis procedure, the influence of geometric nonlinearities on skin-stiffener interface stresses is evaluated. It is found that in flexible stiffened skin structures, which exhibit out-of-plane deformation on the order of 2 to 4 times the skin thickness, inclusion of geometrically nonlinear effects in the calculation of interface stresses is very important. Thus, the use of a geometrically linear analysis, rather than a nonlinear analysis, can lead to considerable error in the computation of the interface stresses. Finally, using the analytical tool developed in this investigation, it is possible to study the influence of stiffener parameters on the state of interface stresses.
Ph. D.

A study aimed at determining the thermal deformation response and thermal buckling loads of rectangular grid-stiffened composite panels is presented. Two edge conditions are considered for the panel, one in which all panel edges are free to deform, and another when all the edges are restrained. In the first case panel deformations due to a uniformly distributed thermal load are analyzed. In the latter case, thermal loads causing buckling failure due to the suppressed in-plane deformations are determined. The panel is composed of a skin and a network of stiffeners, which are all made of the same graphite-epoxy composite material. Kirchhoff's Theory is used to determine the pre-buckling deformations and load distributions of the composite laminates for a panel with free to deform edges. To illustrate both the in-plane and out-of-plane deformations of plate structures under uniform thermal loads, two thermal coefficient vectors, thermal expansion and thermal bending coefficient vectors are introduced. Linear panel buckling analysis performed by assuming a linear undeformed prebuckling state. Rayleigh-Ritz Method, which utilizes minimization of the total energy of a structure to determine the buckling loads, is used to govern the buckling analysis of composite laminates forming the panel. Lagrange Multiplier Method is used along with the Rayleigh-Ritz Method to enforce the deformation continuity constraints at discrete locations along the skin and stiffener interface. As a result, graphical and numerical presentations of the effects of skin and stiffener laminate stacking sequences on the thermal deformations and on the thermal buckling load of the grid-stiffened panel are given.
Master of Science

Stiffeners attached to composite panels and shells may significantly increase the overall buckling load of the resultant stiffened structure. Initially, an extensive literature review was conducted over the past ten years of published work wherein research was conducted on grid stiffened composite structures and stiffened panels, due to their applications in weight sensitive structures. Failure modes identified in the literature had been addressed and divided into a few categories including: buckling of the skin between stiffeners, stiffener crippling and overall buckling. Different methods have been used to predict those failures. These different methods can be divided into two main categories, the smeared stiffener method and the discrete stiffener method. Both of these methods were used and compared in this thesis. First, a buckling analysis was conducted for the case of a grid stiffened composite pressure vessel. Second, a buckling analysis was conducted under the compressive load on the composite stiffened panels for the case of one, two and three longitudinal stiffeners and then, using different parameters, stiffened panels under combined compressive and shear load for the case of one longitudinal centric stiffener and one longitudinal eccentric stiffener, two stiffeners and three stiffeners.
Master of Science

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.

This investigation was aimed at understanding the global and local strain and deflection responses of stiffened skins. Global deformations of the stiffened skins, under load, produce high local stresses in the interface region between the stiffener and skin. Test panels were designed to study the stiffener and skin interactions using parameters typical of stiffened skins for aircraft fuselages. A total of six panels were tested. Two skin laminates, both 0.04 in. thick, and three stiffener configurations were studied. The panels, having clamped edge boundary conditions, were subjected to pressure loads of up to 14.5-14.8 psi. Out-of-plane deflections and longitudinal and transverse strains were measured in several locations. The deflection responses showed a strongly nonlinear behavior at pressure loads of less than 5 psi. In addition relatively severe gradients of both longitudinal and transverse strains were measured in the interface region of the stiffener and skin. Finite element models incorporating geometric nonlinearities were made of four of the panels. Results of these models substantiated the overall findings of the experimental measurements.
Master of Science
incomplete_metadata

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.

In this study, the &ldquo
Free Flexural (or Bending) Vibrations of Stiffened Plates or Panels&rdquo
are investigated in detail. Two different Groups of &ldquo
Stiffened Plates&rdquo
will be considered. In the first group, the &ldquo
Type 4&rdquo
and the &ldquo
Type 6&rdquo
of &ldquo
Group I&rdquo
of the &ldquo
Integrally-Stiffened and/or Stepped-Thickness Plate or Panel Systems&rdquo
are theoretically analyzed and numerically solved by making use of the &ldquo
Mindlin Plate Theory&rdquo
. Here, the natural frequencies and the corresponding mode shapes, up to the sixth mode, are obtained for each &ldquo
Dynamic System&rdquo
. Some important parametric studies are also presented for each case. In the second group, the &ldquo
Class 2&rdquo
and the &ldquo
Class 3&rdquo
of the &ldquo
Bonded and Stiffened Plate or Panel Systems&rdquo
are also analyzed and solved in terms of the natural frequencies with their corresponding mode shapes. In this case, the &ldquo
Plate Assembly&rdquo
is constructed by bonding &ldquo
Stiffening Plate Strips&rdquo
to a &ldquo
Base Plate or Panel&rdquo
by dissimilar relatively thin adhesive layers. This is done with the purpose of reinforcing the &ldquo
Base Plate or Panel&rdquo
by these &ldquo
Stiffening Strips&rdquo
in the appropriate locations, so that the &ldquo
Base Plate or Panel&rdquo
will exhibit satisfactory dynamic response. The forementioned &ldquo
Bonded and Stiffened Systems&rdquo
may also be used to repair a damaged (or rather cracked) &ldquo
Base Plate or Panel&rdquo
. Here in the analysis, the &ldquo
Base Plate or Panel&rdquo
, the &ldquo
Stiffening Plate Strips&rdquo
as well as the in- between &ldquo
adhesive layers&rdquo
are assumed to be linearly elastic continua. They are assumed to be dissimilar &ldquo
Orthotropic Mindlin Plates&rdquo
. Therefore, the effects of shear deformations and rotary moments of inertia are considered in the theoretical formulation. In each case of the &ldquo
Group I&rdquo
and &ldquo
Group II&rdquo
problems, the &ldquo
Governing System of Dynamic Equations&rdquo
for every problem is reduced to the &ldquo
First Order Ordinary Differential Equations&rdquo
. In other words the &ldquo
Free Vibrations Problem&rdquo
, in both cases, is an &ldquo
Initial and Boundary Value Problem&rdquo
is reduced to a &ldquo
Two- Point or Multi-Point Boundary Value Problem&rdquo
by using the present &ldquo
Solution Technique&rdquo
. For this purpose, these &ldquo
Governing Equations&rdquo
are expressed in &ldquo
compact forms&rdquo
or &ldquo
state vector&rdquo
forms. These equations are numerically integrated by the so-called &ldquo
Modified Transfer Matrix Method (MTMM) (with Interpolation Polynomials)&rdquo
. In the numerical results, the mode shapes together with their corresponding non-dimensional natural frequencies are presented up to the sixth mode and for various sets of &ldquo
Boundary Conditions&rdquo
for each structural &ldquo
System&rdquo
. The effects of several important parameters on the natural frequencies of the aforementioned &ldquo
Systems&rdquo
are also investigated and are graphically presented for each &ldquo
Stiffened and Stiffened and Bonded Plate or Panel System&rdquo
. Additionally, in the case of the &ldquo
Bonded and Stiffened System&rdquo
, the significant effects of the &ldquo
adhesive material properties&rdquo
(i.e. the &ldquo
Hard&rdquo
adhesive and the &ldquo
Soft&rdquo
adhesive cases) on the dynamic response of the &ldquo
plate assembly&rdquo
are also presented.

Les lisses des avions de nouvelle génération seront collées (collage ou co-cuisson) à la peau de fuselage. L'objectif de ce travail est d'étudier les liaisons entre les raidisseurs et la peau afin d'assurer le transfert des efforts durablement en particulier dans les phases de post-flambement des structures. Dans un premier temps, une campagne d'essais est menée sur des assemblages " peau/semelles de raidisseurs ". L'influence de la température, du vieillissement ainsi que le comportement en fatigue sont étudiés. Des modèles numériques industrialisables sont mis en place et montrent qu'ils permettent de corréler de façon satisfaisante les résultats d'essais à la fois en statique et en fatigue. Dans un deuxième temps, des essais de flexion trois et quatre points portant sur des assemblages " peau/raidisseurs complets " permettent de dégager les spécificités relatives aux raidisseurs en forme d'oméga ainsi que l'influence des coins de résine liés au mode de fabrication. Des essais spécifiques de flexion sept points montrent les effets de la flexion bi-axiale rencontrée dans les phases de post-flambement des structures. Des études numériques complètent l'analyse et valident les critères de décollement au niveau structural. Une approche " global/local " permettant de diminuer de façon considérable les temps de calcul est ensuite proposée et validée. Pour finir, deux études portant sur des essais de post-flambement sur des panneaux raidis en compression et cisaillement sont effectuées. L'analyse de ces essais permet de valider l'ensemble des méthodes développées au cours de l'étude et de tirer les principales conclusions
For composite fuselage applications, stringers will be bonded to the skin. The aim of this work is to study skin to stringers attachments in order to ensure load transfers mainly during post-buckling phases and taking into account environmental effects and fatigue. Firstly, a test campaign was launched on "skin/stringer flanges". Influences of temperature, ageing and fatigue behaviour have been studied. Numerical models adapted to industrial requirements have been set up and allow a good correlation between tests and calculations in fatigue and static. Secondly, three and four points bending tests on "skin/full stringers" assemblies were performed and demonstrated the particularities of omega stringers and the influence of resin fillets linked to the manufacturing process. Furthermore, specific seven points bending tests show effects of biaxial bending observed during post-buckling phases. For all the tests performed, numerical studies allow the validation of the debonding criterion previously developed and a global/local approach is proposed at structural level. Finally, two stiffened panels were tested in compression and shear. The post-buckling behaviour is studied and allows the validation of the methods developed during the study and provides the main conclusions

The use of the stiffened panels in the aircraft/aeronautical industry has been growing in the last decades. On the other hand, the exponential growth in the use of composite materials in the last years has had a strong 2/2 influence in these structural components and in the industry in general. In consequence, with this new material unknown characteristics appear, for example new failure mechanisms, producing high complexity when simulation, analysis and testing are performed. For this reason, thanks to the increment in the power of the computers, the use of virtual tests with finite element method has become crucial in the simulation of the components with high structural responsibility. In the same way, the general spread of computational resources has made possible the use of optimization methods in the design process of stiffened panels. Optimization methods are able to find the best design according to some criteria, by modifying different parameters
L'ús de panells rigiditzats a la indústria aeronàutica i aeroespacial ha anat creixent les darreres dècades. Per altra banda, el creixement exponencial de l'ús dels materials compòsits en els últims anys també ha tingut una forta incidència en aquests components estructurals i en la indústria en general. Aquest nous materials fan aparèixer comportaments desconeguts fins al moment, com per exemple l'aparició de nous mecanismes de fallada. Aquests fets provoquen que el càlcul, anàlisi i assaig d'estructures de material compòsit sigui complex. Per aquest motiu, sumat a l'augment de potència de càlcul dels ordinadors, l'assaig virtual amb el mètode dels elements finits ha anat agafant una importància cabdal en el càlcul de components d'alta responsabilitat estructural. De la mateixa manera, l'intent de millorar els panells rigiditzats ha portat a utilitzar mètodes d'optimització. Modificant diferents paràmetres es busca dissenyar panells rigiditzats per realitzar una tasca desitjada de manera òptima

碩士
大葉大學
工業工程與科技管理學系
94
The main object of this research was constructed a Stiffened Composite Flat-Panel Speaker of Nano-Carbon Tub and utilized the heuristic algorithms to smooth Sound Pressure Level (SPL) curves by different design parameters. The panel speaker was developed by computer aided engineering, optimization design, measuring and making. The composite panel , spring system and exciting panel was discussed in this investigation. Besides, the model of speaker was simulated by the commercial finite element package, ANSYS. The volume of speaker was defined in 30mm*18mm*7mm, and the frequency was during 300 to 20000Hz. The goal value of SPL was defined as the first nature frequency over 75dB and the mediant valley below 6dB. The average value of SPL during 10000 to 20000Hz was over 70dB. The distance between speaker and the SPL receiver was 30cm. It is shown that the heuristic algorithms can achieve the target efficiently.

Since the Second World War fiber reinforced polymer (FRP) composites have become more attractive as a structural material in a variety of engineering practices, such as infrastructure construction, automobile industry and aerospace engineering, due to high specific strength and stiffness as well as flexibility in tailoring the structural performance. On the other hand, a stiffened panel always performs better in resisting loads compared to an unstiffened panel of same weight. Thus a combination of lightweight composite materials and stiffened structural forms can efficiently enhance the load resisting capability that can be buckling strength of a structure. Stiffened composite panels are subjected to any combination of in-plane, out-of- plane and shear load conditions during service life. These types of thin-walled structures are vulnerable to lose global and local stability under compression loadings. Consequently, buckling-resistant design is one of the most critical issues of stiffened composite panels applied in real practices. Moreover, the buckling optimization design of composite panels is usually a typical multimodal optimization problem, in which there exist multiple global optimal solutions with identical or closely comparable optima of structural performance. Recently, with the development of manufacturing techniques, advanced grid stiffened (AGS) composite panels have increasingly emerged and gained more attention as these grid-stiffening configurations help to enhance the structural efficiency in a more effective way in complex loading conditions compared with conventional unidirectionally-stiffened composite panels. These grid-stiffening configurations provide more available options to select outstanding concept designs of AGS composite panels against buckling for the final appropriate design development at the final construction stage. In this PhD thesis, a novel multi-level optimization methodology for concept design of advanced grid stiffened composite panels against buckling has been developed. Furthermore, an efficient finite element (FE) modelling component for buckling analysis and a robust particle swarm optimization (PSO) algorithms for multimodal optimization have been presented, in order to further consolidate the performance of the proposed methodology. The thesis is divided into six chapters, which are briefly described below: In Chapter 1, a general background along with the objective and originality of the present research is presented. An efficient FE modelling technique is presented in Chapter 2, for the prediction of buckling response of grid stiffened composite panels having different stiffening arrangements. The laminated skin of the stiffened structure is modelled with a triangular degenerated curved shell element. An efficient curved beam element compatible with the shell element is developed for the modelling of stiffeners which may have different lamination schemes. The deformation of the beam element is completely defined in terms of the degrees of freedom of shell elements and it does not require any additional degrees of freedom. Chapter 3 aims to extend conventional unimodal optimization to challenging multimodal optimization of composite structures, by means of newly emerged multimodal PSO using niching techniques. It has shown that the ring topology based PSO without any niching parameter is more robust and efficient for multimodal optimization of composite structures, compared with the species-based PSO (SPSO) and the fitness Euclidean-distance ratio based PSO (FERPSO). In Chapter 4, a random reflection boundary is proposed to replace the conventional fixed absorption boundary for the range-exceeding particles, in order to eliminate/reduce the significance and sensitivity of an empirical parameter of particles’ maximum velocity in PSO. Based on the results obtained from the experimentation on the abovementioned test functions, empirical guidelines for appropriately using the half-range/full-range random reflection boundary are further proposed. Chapter 5 presents an efficient methodology to conduct concept design of AGS composite panels, based on a multi-level approach where an inner 3- stage optimization process is nested within an outer 3-step optimization process. The proposed methodology is applied to a design optimization problem of an AGS composite plate against its buckling resistance, by incorporating a ring topology based multimodal PSO algorithm with an improved FE buckling analysis model. Finally, the conclusions of the present research are summarized in Chapter 6. The limitations and the future development directions of the present study are also described in this chapter.
Thesis (Ph.D.) -- University of Adelaide, School of Civil, Environmental and Mining Engineering, 2015

碩士
國立雲林科技大學
機械工程系
102
In the optimization of fiber angles for composite laminates, the fiber angles of some layers have major effects on the objective, while some just have minor effects. To solve this problem, a genetic algorithm with the function of elite comparison is proposed in this study. Also in this algorithm the unnecessary repeat execution of the finite element analysis is avoided. In order to validate this proposed algorithm, several examples are investigated to have comparison with the results from the literature. For example, the fiber angle optimization to fit the known frequencies、maximization of the fundamental frequency of laminated composites、the fiber angle optimization for composite laminated with a central hole suffered uniaxial load, etc. The results indicate that the proposed algorithm can accelerate the convergence speed and improve the accuracy of the results. In order to explore the fiber angle for the stringer-stiffened composite panel to resist buckling, the problem is divided into two types. One type has the same fiber angle for both the skin and the stinger, and they are the design variables. Form the results ,although the error of one type 128 training group is more than 200 training, but the result that predict the angle is equal. In the other problem, the fiber angle of the skin is fixed, while the fiber angles of the stringer are set as the design variables. That can find maximum buckling load. To save the calculation time of finite element analysis, a back-propagation neural network is established and used with the proposed genetic algorithm to search the optimal solution.

碩士
國立交通大學
機械工程系所
92
The effects on sound radiation of composite panels with different stiffened designs are studied in this thesis. This study is used to design the stiffened composite panels with enough stiffness which can avoid the fluctuation of sound pressure. By investigating the differences among the behaviors of sound pressure responses of four types of stiffened plates and an unstiffened one, we obtain a structural design which can effectively modify the stiffness of a plate structure and its sound radiation behavior. For theoretical analysis, the simulation models of radiating panels with different designs of stiffeners were constructed using the software ANSYS. The amplitudes and the phases of these panels were obtained from the analyses of the simulation models using given system parameters such as spring constants of suspension and damper, the thrust of the voice coil and the damping ratio of the loudspeaker system, which were measured from experiments. Then, we use the amplitudes and the phases in the program of calculation of sound pressure level to simulate the behaviors of sound radiation of the loudspeakers. Through the comparison between the curves of sound pressure response measured by the sound measurement system MLSSA and the ones simulated by computer, we can find out the reason causing the drops in sound pressure curves and the direction to improve the loudspeaker system further. Finally, the discrepancies between the results of experiment and these of simulation were discussed.

碩士
國立交通大學
機械工程系所
96
Optimum Design of Stiffened Composite Sound Radiating Panels Student : Cheng-Lin Yu Advisor : Dr. Tai-Yan Kam Department of Mechanical Engineering National Chiao Tung University ABSTRACT In this thesis, the first part studies the effects of elastic suspensions and panel rigidity on the sound pressure level (SPL) curves of composite sound radiating panels. The SPL curves measured by the used of wave-type suspensions and L-type suspensions in loudspeakers have shown that these effects are minimal. But, the effects of panel rigidity are too significant to be neglected. In order to improve the smoothness of SPL curve, composite stiffeners are added to the radiating plate to suppress the mid-frequency dip of the SPL curve. In the theoretical study, some mode shapes of the radiating plate that may have adverse influences on the sound radiation in certain frequency ranges are determined using a verified ANSYS FEA model. An optimal design method is proposed to determine the dimensions of the stiffeners for different stiffening patterns. For each stiffening pattern, the optimal stiffener sizes are determined to suppress the detrimental modes which are harmful to sound radiation and make the SPL curve smoother and the speaker possess higher sensitivity. Experiments are performed to verify the feasibility of the proposed design of the loudspeaker.

碩士
國立交通大學
機械工程學系
100
This paper is focused on the sound radiation behavior of stiffened sound radiation panels for flat-panel loudspeaker. If the sound radiation panel is not properly design, it is easy for a flat-panel speaker to have a significant sound pressure drop in the mid-frequency range, which may affect the sound quality of the speaker. To suppress or even eliminate the sound pressure drop, this thesis attempts to enhance the rigidity of the plate by using an appropriate stiffening method. There are two parts in this thesis. In the first part, a Rayleigh-Ritz Method is constructed to analyze the sound radiation behavior of stiffened sound radiation panels. The results obtained using the proposed Rayleigh-Ritz Method are compared with those obtained from the finite element software ANSYS. It has been shown that the natural frequencies and the sound pressure level (SPL) curves produced by the proposed method are in good agreement with those produced by ANSYS or experiments. In the second part, ANSYS is used to find the appropriate stiffening pattern for designing the sound radiation panel and the ideal material constants of the panel so that the major sound pressure drop in the mid-frequency range can be suppressed.

No
In this paper, a new three-dimensional spline finite strip method (spline FSM) is introduced. This is done by combining the classical spline finite strip method [1] and the state space approach. According to the traditional spline FSM, a laminated plate is divided into strips. Within each strip, the spline FSM calls for the use of simple polynomials and a continuously differentiable spline function, respectively, in the transverse and in-plane directions. In the through-thickness direction, the state space method is used to compute the distribution of displacements and stresses. The combination of the in-plane spline FSM and the out-of-plane state space formulations results in a global state space equation that is solved numerically by the precise time step integration method [2,3]. Apart from obtaining a three-dimensional solution, the new method has a unique feature that the final algebra equation system is independent of the number of material layers of a laminate. The main aim of this work is to establish the new solution procedure and validate the method. To this end, the work reported in the paper focus on laminated plates with arbitrary boundary conditions. Thus, the spline FSM is more flexible than the FSM in imposing boundary conditions. Future development is expected to extend the solution to more practical applications. From the numerical validation included, it can be seen clearly that the newly developed method can provide accurate three dimensional solutions for laminated composites, particularly, with continuous transverse stress distributions across material interfaces. This is normally difficult to obtain if a traditional three dimensional finite element is used, where only continuity of displacements across material boundaries are guaranteed. Apart from the above new feature, the new three-dimensional formulation always ends up with a global matrix whose dimension depends only on the number of strips and knots that a plate has been divided into, and is completely independent of the number of material layers of the plate.

abstract: Composite materials are increasingly being used in aircraft, automobiles, and other applications due to their high strength to weight and stiffness to weight ratios. However, the presence of damage, such as delamination or matrix cracks, can significantly compromise the performance of these materials and result in premature failure. Structural components are often manually inspected to detect the presence of damage. This technique, known as schedule based maintenance, however, is expensive, time-consuming, and often limited to easily accessible structural elements. Therefore, there is an increased demand for robust and efficient Structural Health Monitoring (SHM) techniques that can be used for Condition Based Monitoring, which is the method in which structural components are inspected based upon damage metrics as opposed to flight hours. SHM relies on in situ frameworks for detecting early signs of damage in exposed and unexposed structural elements, offering not only reduced number of schedule based inspections, but also providing better useful life estimates. SHM frameworks require the development of different sensing technologies, algorithms, and procedures to detect, localize, quantify, characterize, as well as assess overall damage in aerospace structures so that strong estimations in the remaining useful life can be determined. The use of piezoelectric transducers along with guided Lamb waves is a method that has received considerable attention due to the weight, cost, and function of the systems based on these elements. The research in this thesis investigates the ability of Lamb waves to detect damage in feature dense anisotropic composite panels. Most current research negates the effects of experimental variability by performing tests on structurally simple isotropic plates that are used as a baseline and damaged specimen. However, in actual applications, variability cannot be negated, and therefore there is a need to research the effects of complex sample geometries, environmental operating conditions, and the effects of variability in material properties. This research is based on experiments conducted on a single blade-stiffened anisotropic composite panel that localizes delamination damage caused by impact. The overall goal was to utilize a correlative approach that used only the damage feature produced by the delamination as the damage index. This approach was adopted because it offered a simplistic way to determine the existence and location of damage without having to conduct a more complex wave propagation analysis or having to take into account the geometric complexities of the test specimen. Results showed that even in a complex structure, if the damage feature can be extracted and measured, then an appropriate damage index can be associated to it and the location of the damage can be inferred using a dense sensor array. The second experiment presented in this research studies the effects of temperature on damage detection when using one test specimen for a benchmark data set and another for damage data collection. This expands the previous experiment into exploring not only the effects of variable temperature, but also the effects of high experimental variability. Results from this work show that the damage feature in the data is not only extractable at higher temperatures, but that the data from one panel at one temperature can be directly compared to another panel at another temperature for baseline comparison due to linearity of the collected data.
Dissertation/Thesis
M.S. Aerospace Engineering 2012

碩士
大葉大學
工業工程與科技管理學系
98
The main object of this paper is designed and developed a double flat-panel speaker stiffened by nano-carbon tube composites which it had thin thicknesses, broad frequency and acoustic fidelity vigorous not distorted it. Two types of vibrating plate, namely, a high audio speaker and medium-low audio speaker constructed on the basis of the manufacture technique and sound pressure theory are developed for the design and analysis of double flat-panels. The study is analyzed the frequency and sound pressure value of double flat-panel speakers with different design parameters such as stiffness and weight of composite panels, boundary condition and spring constant of suspension system and vibration area which are constructed using a finite element constructed on the basis of the software ANSYS. The double flat-panel speakers can be applied to the general plane video and music electronic products loudspeaker system, achieves nowadays pursues the monitor more and more thin tendency. The double rectangular flat-panel speaker can be used in dual-channel flat-panel speakers for portable DVD players and notebook computers ... and other products. According to developed the analytical method of a set of simulation and optimal design is proceed to optimal design of multiple objective function for stiffened composite double flat-panels in 100Hz~20KHz frequency zones. The 100Hz~20KHz frequency zones had divide into 4 zones, every zones variable values multiply by weight was the sum that was the objective function of multiple optimal design. In the limit small thrust of low power, is used the optimal method to find the best manufacturing parameters (includes the lengths of flat-panel speaker, spring constant of suspension system, stiffened types and vibration lengths) made the sound pressure value curve had smooth and get the best sound pressure value. Therefore, the optimal manufacturing parameters would manufacture double flat-panels to measured sound pressure curve that compared experimental values and theory values.

碩士
大葉大學
工業工程與科技管理學系
95
The main object of this paper is developed a miniature flat-panel speakers stiffened by nano-carbon tube composites which are low frequency sound quality, smooth curve of sound pressure, and reduce the decay rate of high frequency sounds. The several standard of flat-panel speakers stiffened including the 30mm×18mm×7mm、40mm×14mm×7mm and 50mm×14mm×7mm. The paper is used the ANSYS software to solved the sound curve in flat-panel speakers stiffened and used optimal theory to solved optimal manufacture parameters (including the thickness ratio of flat-panel speaker and nano-carbon tubes that in the same weight, boundary condition and spring constant of suspension system, vibration area and location) which make the sound pressure value curve is smooth in global frequency. The flat-panel speakers stiffened which is developed by this project can reach the goals of economize electric power, maximum bearing, the low-frequency had powerful voices and the high-frequency had a better clarity. According to the best results of manufacture parameters, the materials and molds of suspension system are choused to manufacture suspension systems and fabricate miniature flat-panel speakers stiffened. The experimental and optimal methods are presented to study the optimal sound pressure curve of flat-panel speaker. The optimal methods proved to be accuracy.

O σχεδιασμός και η ανάπτυξη μιας σύγχρονης αεροναυπηγικής κατασκευής περιλαμβάνει ως επιμέρους φάσεις (μεταξύ άλλων) τον αρχικό και τον προκαταρκτικό σχεδιασμό. Οι φάσεις αυτές έχουν ιδιαίτερη σημασία διότι εκεί δίνεται η αρχική μορφή και οι διαστάσεις της κατασκευής. Είναι γεγονός ότι η συμβατική σχεδίαση των βασικών δομικών στοιχείων των αεροσκαφών έχει φτάσει σε πολύ υψηλό επίπεδο βελτιστοποίησης που επιδέχεται πλέον μόνο μικρά περιθώρια περαιτέρω βελτίωσης. Οι σύγχρονες όμως απαιτήσεις των ελαφρών κατασκευών, όπως δραστική μείωση του βάρους, αύξηση του ωφέλιμου φορτίου κτλ. ωθεί τις αεροναυπηγικές βιομηχανίες στη δημιουργία δομών που ξεφεύγουν από τις παραδοσιακές (μη-συμβατικές δομές). Παράλληλα με τα παραπάνω γίνεται προσπάθεια για μερική αντικατάσταση μεταλλικών υλικών από σύνθετα υλικά στις πρωτεύουσες δομές αεροναυπηγικών κατασκευών. Για να σχεδιαστούν και να εξελιχθούν μη-συμβατικές αεροναυπηγικές δομές χωρίς να καταφύγει κάποιος σε εκτενείς πειραματικές δοκιμές, η σύγχρονη τάση είναι η ανάπτυξη και ο συνδυασμός προτύπων συμπεριφοράς στη λογική της εξομοίωσης των πειραματικών δοκιμών. Η εξομοίωση αυτή επιτυγχάνεται με τη βοήθεια ηλεκτρονικών υπολογιστών και κατάλληλων μεθόδων βασισμένων στη θεωρία των πινάκων (Πεπερασμένα Στοιχεία, Συνοριακά Στοιχεία κλπ.). Στη φάση του αρχικού και προκαταρκτικού σχεδιασμού η εφαρμογή των μεθοδολογιών προσομοίωσης δεν είναι πάντοτε εύκολη και απλή, λόγω των πολλαπλών αλλαγών στη γεωμετρία, το υλικό και τις κατασκευαστικές λεπτομέρειες που πραγματοποιούνται στη δομή κατά την επαναληπτική διαδικασία του σχεδιασμού. Επομένως, η αποκλειστική χρήση αριθμητικών μεθόδων ανάλυσης καθίσταται αναποτελεσματική από άποψη χρονικών απαιτήσεων, αν δεν συνοδεύεται από αναλυτικές ή ημιαναλυτικές προσεγγίσεις επιμέρους προβλημάτων του σχεδιασμού. Βασικό μέρος του προκαταρκτικού σχεδιασμού μιας πτέρυγας μη συμβατικής δομής αποτελεί η αποφυγή της αστοχίας του άνω τμήματός της, διότι οι λεπτότοιχες ενισχυμένες με δοκούς πλάκες που χρησιμοποιούνται στην κατασκευή υφίστανται λυγισμό λόγω των θλιπτικών φορτίσεων που κυρίως παραλαμβάνουν. Η διαστασιολόγηση των σύνθετων πλακών που φέρουν δοκούς ενίσχυσης στις κατασκευές αυτές απαιτούν συνήθως πλήθος επαναληπτικών υπολογισμών για διαφορετικές γεωμετρίες, φορτίσεις, οριακές συνθήκες κλπ. Η εξέταση της κάθε περίπτωσης μεμονωμένα με τη χρήση αριθμητικών μεθόδων καθιστά την επίλυση ολόκληρης της κατασκευής εξαιρετικά χρονοβόρα. Για το λόγο αυτό, στη φάση της αρχικής θεωρητικής μελέτης και της αρχικής διαστασιολόγησης η χρησιμοποίηση αναλυτικών μεθόδων για την εύρεση του κρίσιμου φορτίου λυγισμού πλακών με δοκούς ενίσχυσης οδηγεί στην εξοικονόμηση υπολογιστικού κόστους. Επομένως, η ανάπτυξη αναλυτικών ή ημιαναλυτικών μεθόδων προσδιορισμού των φορτίων λυγισμού ενισχυμένων με δοκούς συνθέτων πλακών και κελυφών θεωρείται πολύ σημαντική. Για τον σκοπό αυτό, στο πλαίσιο αυτής της διατριβής, αναπτύσσονται αναλυτικές και ημιαναλυτικές λύσεις για το λυγισμό πολύστρωτων πλακών ενισχυμένων με ενισχυτικές διαμήκεις δοκούς, οι οποίες ενσωματώνονται σαν κριτήρια στη μέθοδο διαστασιολόγησης της δομής. Η μεθοδολογία συμπληρώνεται με πλήθος άλλων κατάλληλων κριτηρίων για τον έλεγχο της αντοχής των δομικών στοιχείων της πτέρυγας καθώς και με κριτήρια για την επαναδιαστασιολόγηση των στοιχείων κατά την επαναληπτική διαδικασία της βελτιστοποίησης. Με τη μεθοδολογία που αναπτύσσεται διερευνούνται διατάξεις δομής πτερύγων από σύνθετα υλικά με πολυάριθμες κύριες δοκούς. Πιο συγκεκριμένα, αναπτύσσονται αναλυτικές/ημιαναλυτικές λύσεις ολικού και τοπικού λυγισμού πλακών που φέρουν δοκούς ενίσχυσης. Όσον αφορά τον ολικό λυγισμό αναπτύσσεται μια μεθοδολογία που βασίζεται στη μαθηματική μετατροπή μιας πλάκας που φέρει δοκούς ενίσχυσης σε μια ισοδύναμη ομογενή πλάκα. Η αναπτυχθείσα μεθοδολογία ομογενοποίησης των ενισχυμένων πλακών εμφανίζει σημαντικά πλεονεκτήματα σε σύγκριση με τις αντίστοιχες ήδη υπάρχουσες. Παράλληλα, η ενεργειακή μέθοδος Rayleigh-Ritz εφαρμόζεται για τη λύση προβλημάτων λυγισμού μερικώς ανισότροπων πλακών με ενισχυτικές δοκούς από σύνθετα υλικά, λαμβάνοντας διακριτά υπόψη τις ενισχυτικές δοκούς. Όσον αφορά το πρόβλημα του τοπικού λυγισμού, αναπτύσσεται μια νέα μεθοδολογία για την εύρεση των κρίσιμων φορτίων τοπικού λυγισμού λεπτότοιχης πλάκας με χρήση ενεργειακών μεθόδων. Το μαθηματικό μοντέλο που χρησιμοποιείται για την περίπτωση του τοπικού λυγισμού της επικάλυψης είναι η απομόνωση του τμήματος της επικάλυψης μεταξύ δυο ενισχυτικών δοκών και η αντικατάσταση της δυσκαμψίας της υπόλοιπης πλάκας με ελατήρια μεταβλητής δυσκαμψίας. Η μεθοδολογία αυτή επεκτείνεται και στον προσδιορισμό της μεταλυγισμικής συμπεριφοράς μιας πλάκας ενισχυμένης με διαμήκεις δοκούς. Οι παραπάνω μεθοδολογίες υπολογισμού του κρίσιμου φορτίου λυγισμού που αναπτύσσονται, στα πλαίσια αυτής της διατριβής, εφαρμόζονται στη διαστασιολόγηση πτέρυγας μη συμβατικής δομής από σύνθετα υλικά με πολυάριθμες κύριες δοκούς, σε αντίθεση με τις συμβατικές πτέρυγες (με δύο κύριες δοκούς). Η ανάλυση τάσεων της πτέρυγας πραγματοποιείται με τη βοήθεια της μεθόδου των πεπερασμένων στοιχείων. Η τελική διαστασιολόγηση επιτυγχάνεται με επαναληπτική διαδικασία βελτιστοποίησης βασισμένη σε αναλυτικές και ημιαναλυτικές σχέσεις. Με τον τρόπο αυτό, συγκρίνεται λεπτομερώς η συμβατική δομή πτέρυγας με 2 κύριες δοκούς και οι αντίστοιχες πτέρυγες με 4, 5 και 6 κύριες δοκούς. Για την περαιτέρω βελτιστοποίηση της συμπεριφοράς της πτέρυγας, διερευνάται η επίδραση που έχει η αλλαγή των μηχανικών ιδιοτήτων του υλικού και των επιτρεπόμενων ορίων παραμόρφωσης στη δυνατότητα ελαχιστοποίησης της μάζας της πτέρυγας. Υπολογίστηκε ότι κάτω από συγκεκριμένες συνθήκες η χρήση της μη συμβατικής πτέρυγας μπορεί να οδηγήσει σε μείωση μάζας μέχρι και 12%.
The design and development of a modern aerospace structure consists of many design stages. The most important stages are the conceptual and the preliminary where the initial sizing of the structure is obtained. It is known that the conventional design of the aircraft’s main structural members has reached a high optimization level, where margins for further improvement are small. The current demands of the lightweight structures such as weight reduction, payload increase etc. have led the aerospace industries develop unconventional structures and partially substitute the metallic materials of the primary structures with composites. The current trend of designing and evolving unconventional aerospace structures, without performing extended experimental tests, leads to the development of behavior models. The simulation of the experimental tests (through the behavior models) is achieved using high performance computers and numerical methods (Finite Element Method, Boundary Element Method etc). To apply simulation methods during the conceptual and preliminary stage is not an easy task. Most of the difficulties are the numerous geometrical, material parameters and the structural details that alter during the iterative process of the design. So, the exclusive usage of numerical analysis methods becomes very time consuming, if it is not accompanied by analytical or semi analytical methods of the sub-problems of the design. Part of the preliminary design of an unconventional wing structure is to prevent upper skin from failure. The stiffened panels that comprise the upper skin of the wing suffer from buckling due to the applied compressive loads. The sizing of the composite stiffened panels usually requires numerous of iterative calculations for various geometries, loading and boundary conditions etc. The examination of each case separately, with the use of numerical methods, results to time consuming analyses of the entire structure. Therefore, the development of appropriate analytical or semi analytical methods for estimating stiffened panels’ critical buckling load is of great importance. For this purpose, in the present thesis, analytical and semi analytical methodologies are developed for estimating the critical buckling load of stiffened panels. The developed methodologies are incorporated as design criteria in the sizing routine of the entire structure. The sizing routine comprises additional sizing criteria for checking the strength of wing’s structural members at each phase of the iterative process. Applying the developed sizing routine in various wing configurations made of composite materials, multispar wing designs are studied. Specifically, analytical and semi analytical methods for global and local buckling problems of stiffened panels are developed. The methodology of global buckling problems is based on the mathematical conversion of a stiffened panel to an equivalent homogeneous panel. The developed method of homogenization of stiffened panels appears to have significant advantages over the already existed homogenization methods. Additionally, the energy method Rayleigh-Ritz is applied for solving global buckling problems of stiffened panels with partial anisotropy considering discrete stiffeners. Regarding local buckling problems of stiffened panels, a new methodology is developed for estimating the critical local buckling load with the use of energy methods. The approach considers the stiffened panel segment located between two stiffeners, while the remaining panel is replaced by equivalent transverse and rotational springs of varying stiffness, which act as elastic edge supports. The buckling analysis of the segment provides an accurate and conservative prediction of the panel local buckling behavior. Consequently, the developed methodology is extended in the prediction of post-buckling response of stiffened panels where skin has undergone local buckling. The developed methodologies for calculating the critical buckling load are applied for sizing the wing members of an unconventional wing (multispar configuration) from composite materials. An efficient methodology based on fast Finite Element (FE) stress analysis combined to analytically formulated design criteria is presented for the initial sizing of a large scale composite component. A detailed comparison between optimized designs of conventional (2-spar) and three alternative wing configurations which comprise 4-, 5-, and 6-spars for the wing construction is performed. In order to understand the effect of different material properties, as well as the variation of maximum strain level allowed in the total wing mass, parametric analyses are performed for all wing configurations considered. It arises that under certain conditions the multispar configuration demonstrates significant advantages over the conventional design. This would lead to a mass reduction of 12%.