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.

In this study a methodology is developed for the fatigue and fracture analysis of helicopter fuselage structures, which are considered as the stiffened panels. The damage tolerance behavior of the stiffened panels multiaxially loaded is investigated by implementing virtual crack closure technique (VCCT). Validation of VCCT is done through comparison between numerical analysis and the studies from literature, which consists of stiffened panels uniaxially loaded and the panel with an inclined crack. A program based on Fortran programming language is developed to automate the crack growth analysis under mixed mode conditions. The program integrates the prediction of the change in crack propagation direction by maximum circumferential stress criterion and the computation of energy release rate by VCCT. It allows reducing the computation time for damage tolerance evaluation for mixed mode cases through finite element analysis and runs the procedure file of MSC.Marc/Mentat for numerical analysis and the program generated by Patran Command Language (PCL) of MSC.Patran for remeshing. The developed code is verified by comparing the crack growth trajectories obtained by numerical analysis with the experimental studies from literature. A submodeling technique is utilized to analyze a particular fuselage portion of helicopter tail boom. Effects of different skin/stringer configurations of the helicopter fuselage structure on stress intensity factor are studied by means of the developed program. Fatigue crack growth analysis is performed by using stress intensity factors obtained from numerical analysis and fatigue propagation models proposed in literature.

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.

A global/local framework for multidisciplinary optimization of generalized aircraft wing structure has been developed. The concept of curvilinear stiffening members (spars, ribs and stiffeners) has been applied in the optimization of a wing structure. A global wing optimization framework EBF3WingOpt, which integrates the static aeroelastic, flutter and buckling analysis, has been implemented for exploiting the optimal design at the wing level. The wing internal structure is optimized using curvilinear spars and ribs (SpaRibs). A two-step optimization approach, which consists of topology optimization with shape design variables and size optimization with thickness design variables, is implemented in EBF3WingOpt. A local panel optimization EBF3PanelOpt, which includes stress and buckling evaluation criteria, is performed to optimize the local panels bordered by spars and ribs for further structural weight saving. The local panel model is extracted from the global finite element model. The boundary conditions are defined on the edges of local panels using the displacement fields obtained from the global model analysis. The local panels are optimized to satisfy stress and buckling constraints. Stiffened panel with curvilinear stiffeners is implemented in EBF3PanelOpt to improve the buckling resistance of the local panels. The optimization of stiffened panels has been studied and integrated in the local panel optimization. EBF3WingOpt has been applied for the optimization of the wing structure of the Boeing N+2 supersonic transport wing and NASA common research model (CRM). The optimization results have shown the advantage of curvilinear spars and ribs concept. The local panel optimization EBF3PanelOpt is performed for the NASA CRM wing. The global-local optimization framework EBF3GLWingOpt, which incorporates global wing optimization module EBF3WingOpt and local panel optimization module EBF3PanelOpt, is developed using MATLAB and Python programming to integrate several commercial software: MSC.PATRAN for pre and post processing, MSC.NASTRAN for finite element analysis. An approximate optimization method is developed for the stiffened panel optimization so as to reduce the computational cost. 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. / Ph. D.

This thesis presents the results of 67 ABAQUS elasto-plastic Riks ultimate strength analyses of cross-stiffened panels. These panels cover a wide range of typical geometries. Uniaxial compression is applied to the panels, and in some cases combined with lateral pressure. For eight of the panels full-scale experimental results are available, and these verified the accuracy of the ABAQUS results. The 67 ABAQUS results were then compared to the ultimate strength predictions from the computer program ULSAP. In all but 10 cases the ULSAP predicted strength is within 30% of the ABAQUS value, and in all but 4 cases the predicted failure mode also agrees with that of ABAQUS. In one case the ULSAP predicted ultimate strength is 51% below the experimental value, and so this case is studied in detail. The discrepancy is found to be caused by the method which ULSAP uses for panels that experience overall collapse initiated by beam-column-type failure. The beam-column method program ULTBEAM is used to predict the ultimate strength of the 61 panels that ULSAP predicts to fail due to overall collapse of the stiffeners and plating which may or may not be triggered by yielding of the plate-stiffener combination at the midspan (Mode III or III-1). ULTBEAM is found to give more accurate results than ULSAP for Mode III or III-1 failure. Future work is recommended to incorporate ULTBEAM into ULSAP to predict the ultimate strength of panels that fail in Mode III or III-1. / Master of Science

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