The high stiffness to weight ratio and fatigue resistance make carbon fibre composites suitable for both military and large civil aircraft. The limited ability of current numerical methods to capture the complex growth of damage in laminated composites leads to a conservative design approach applied in today??s composite aircraft structures. The aim of the presented research was to develop an improved methodology for the failure prediction of laminated composites containing delaminations located between arbitrary layers in the laminate, and to extend the investigations to composite structures subjected to barely visible impact damage (BVID). The advantages of fracture mechanics-based methodologies to predict interlaminar failure in composite structures were identified, from which the crack tip element (CTE) approach and the virtual crack closure technique (VCCT) were selected for assessment. Extensive validation of these fracture mechanics methods is presented on a number of composite structures ranging from coupons to large stiffened panels. It was shown that the VCCT was relatively insensitive to the crack front mesh size, whilst predictions using the CTE methodology were significantly influenced by the element size. Based on the obtained results modelling guidelines for the VCCT and CTE were established. Significant contribution of this research to the field of the analysis of composite structures was the development of a novel test method for the evaluation of embedded single and multi-level delaminations. The test procedure of the single delamination specimen was proposed as an analogous test to conventional compression experiments. The transverse test overcame the inherent problems of in-plane compression testing and produced less scatter of experimental measurements. Quantitative analysis of numerical results employing the validated finite element modelling approaches showed that the failure load and location were in agreement with experiments. Furthermore, new modelling techniques for composite structures containing BVID proposed in this research produced good correlation with test data from the compression after impact (CAI) test. The study of BVID provided a significant contribution toward the knowledge of the applicability of implicit FE solvers to predict failure of CAI specimens as well as the criticality of centrally impacted specimens.

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 high stiffness to weight ratio and fatigue resistance make carbon fibre composites suitable for both military and large civil aircraft. The limited ability of current numerical methods to capture the complex growth of damage in laminated composites leads to a conservative design approach applied in today??s composite aircraft structures. The aim of the presented research was to develop an improved methodology for the failure prediction of laminated composites containing delaminations located between arbitrary layers in the laminate, and to extend the investigations to composite structures subjected to barely visible impact damage (BVID). The advantages of fracture mechanics-based methodologies to predict interlaminar failure in composite structures were identified, from which the crack tip element (CTE) approach and the virtual crack closure technique (VCCT) were selected for assessment. Extensive validation of these fracture mechanics methods is presented on a number of composite structures ranging from coupons to large stiffened panels. It was shown that the VCCT was relatively insensitive to the crack front mesh size, whilst predictions using the CTE methodology were significantly influenced by the element size. Based on the obtained results modelling guidelines for the VCCT and CTE were established. Significant contribution of this research to the field of the analysis of composite structures was the development of a novel test method for the evaluation of embedded single and multi-level delaminations. The test procedure of the single delamination specimen was proposed as an analogous test to conventional compression experiments. The transverse test overcame the inherent problems of in-plane compression testing and produced less scatter of experimental measurements. Quantitative analysis of numerical results employing the validated finite element modelling approaches showed that the failure load and location were in agreement with experiments. Furthermore, new modelling techniques for composite structures containing BVID proposed in this research produced good correlation with test data from the compression after impact (CAI) test. The study of BVID provided a significant contribution toward the knowledge of the applicability of implicit FE solvers to predict failure of CAI specimens as well as the criticality of centrally impacted specimens.

De nos jours, la durée de vie des modules d’électronique de puissance est désormais limitée par les technologies standards de conditionnement, telles que le câblage par fils et le brasage. Ainsi une optimisation des technologies actuellement employées n’est pas suffisante pour satisfaire les futures exigences de fiabilité. Pour dépasser ces limites, un nouveau module de puissance remplaçant les fils de connexion par des clips en cuivre a été développé. Ce design innovant vise à améliorer la fiabilité du module puisqu’il empêche la dégradation des fils de connexion, constituant bien souvent la principale source de défaillance. La contrepartie de ce gain de fiabilité réside dans la complexification de la structure interne du module. En effet, l’emploi d’un clip en cuivre nécessite une brasure supplémentaire fixant le clip à la puce. Ainsi, le comportement thermomécanique et les différents modes de rupture auxquels le composant est soumis lors de son utilisation doivent être caractérisés. Cette étude utilise la simulation numérique pour analyser avec précision le comportement de chaque couche de matériaux lors des cycles actifs de puissance. De plus, une étude de sensibilité à la fois expérimentale et numérique concernant les paramètres de tests est réalisée. Les zones critiques du module ainsi que les combinaisons critiques des paramètres de tests pour les différents modes de rupture sont mis en évidence. Par ailleurs, une analyse en mécanique de la rupture est conduite et la propagation des fissures à différentes zones clés est analysée en fonction des différents paramètres de tests. Les résultats obtenus permettent la définition de modèles de prédiction de durée de vie. / Today a point has been reached where safe operation areas and lifetimes of power modules are limited by the standard packaging technologies, such as wire bonding and soft soldering. As a result, further optimization of used technologies will no longer be sufficient to meet future reliability requirements. To surpass these limits, a new power module was designed using Cu clips as interconnects instead of Al wire bonds. This new design should improve the reliability of the module as it avoids wire bond fatigue failures, often the root cause of device failures. The counterpart for an improved reliability is a quite complicated internal structure. Indeed, the use of a Cu clip implies an additional solder layer in order to fix the clip to the die. The thermo-mechanical behavior and failure mechanisms of such a package under application have to be characterized. The present study takes advantage of numerical simulations to precisely analyze the behavior of each material layer under power cycling. Furthermore an experimental and numerical sensitivity study on tests parameters is conducted. Critical regions of the module are pointed out and critical combinations of tests parameters for different failure mechanisms are highlighted. Then a fracture mechanics analysis is performed and the crack growth at different locations is analyzed in function of different tests parameters. Results obtained enable the definition of lifetime prediction models.

Global-local finite element analyses were used to study the damage tolerance of curvilinearly stiffened panels; fabricated using the modern additive manufacturing process, the so-called unitized structures, and that of adhesive joints. A damage tolerance study of the unitized structures requires cracks to be defined in the vicinity of the critical stress zone. With the damage tolerance study of unitized structures as the focus, responses of curvilinearly stiffened panels to the combined shear and compression loadings were studied for different stiffeners' height. It was observed that the magnitude of the minimum principal stress in the panel was larger than the magnitudes of the maximum principal and von Mises stresses. It was also observed that the critical buckling load factor increased significantly with the increase of stiffeners' height. To study the damage tolerance of curvilinearly stiffened panels, in the first step, buckling analysis of panels was performed to determine whether panels satisfied the buckling constraint. In the second step, stress distributions of the panel were analyzed to determine the location of the critical stress under the combined shear and compression loadings. Then, the fracture analysis of the curvilinearly stiffened panel with a crack of size 1.45 mm defined at the location of the critical stress, which was the common location with the maximum magnitude of the principal stresses and von Mises stress, was performed under combined shear and tensile loadings. This crack size was used because of the requirement of a sufficiently small crack, if the crack is in the vicinity of any stress raiser. A mesh sensitivity analysis was performed to validate the choice of the mesh density near the crack tip. All analyses were performed using global-local finite element method using MSC. Marc, and global finite element methods using MSC. Marc and ABAQUS. Negligible difference in results and 94% saving in the CPU time was achieved using the global-local finite element method over the global finite element method by using a mesh density of 8.4 element/mm ahead of the crack tip. To study the influence of different loads on basic modes of fracture, the shear and normal (tensile) loads were varied differently. It was observed that the case with the fixed shear load but variable normal loads and the case with the fixed normal load but variable shear loads were Mode-I. Under the maximum combined loading condition, the largest effective stress intensity factor was very smaller than the critical stress intensity factor. Therefore, considering the critical stress intensity factor of the panel with the crack of size 1.45 mm, the design of the stiffened panel was an optimum design satisfying damage tolerance constraints. To acquire the trends in stress intensity factors for different crack lengths under different loadings, fracture analyses of curvilinearly stiffened panels with different crack lengths were performed by using a global-local finite element method under three different load cases: a) a shear load, b) a normal load, and c) a combined shear and normal loads. It was observed that 85% data storage space and the same amount in CPU time requirement could be saved using global-local finite element method compared to the standard global finite element analysis. It was also observed that the fracture mode in panels with different crack lengths was essentially Mode-I under the normal load case; Mode-II under the shear load case; and again Mode-I under the combined load case. Under the combined loading condition, the largest effective stress intensity factor of the panel with a crack of recommended size, if the crack is not in the vicinity of any stress raiser, was very smaller than the critical stress intensity factor. This work also includes the performance evaluation of adhesive joints of two different materials. This research was motivated by our experience of an adhesive joint failure on a test-fixture that we used to experimentally validate the design of stiffened panels under a compression-shear load. In the test-fixture, steel tabs were adhesively bonded to an aluminum panel and this adhesive joint debonded before design loads on the test panel were fully applied. Therefore, the requirement of studying behavior of adhesive joints for assembling dissimilar materials was found to be necessary. To determine the failure load responsible for debonding of adhesive joints of two dissimilar materials, stress distributions in adhesive joints of the nonlinear finite element model of the test-fixture were studied under a gradually increasing compression-shear load. Since the design of the combined load test fixture was for transferring the in-plane shear and compression loads to the panel, in-plane loads might have been responsible for the debonding of the steel tabs, which was similar to the results obtained from the nonlinear finite element analysis of the combined load test fixture. Then, fundamental studies were performed on the three-dimensional finite element models of adhesive lap joints and the Asymmetric Double Cantilever Beam (ADCB) joints for shear and peel deformations subjected to a loading similar to the in-plane loading conditions in the test-fixtures. The analysis was performed using ABAQUS, and the cohesive zone modeling was used to study the debonding growth. It was observed that the stronger adhesive joints could be obtained using the tougher adhesive and thicker adherends. The effect of end constraints on the fracture resistance of the ADCB specimen under compression was also investigated. The numerical observations showed that the delamination for the fixed end ADCB joints was more gradual than for the free end ADCB joints. Finally, both the crack propagation and the characteristics of adhesive joints were studied using a global-local finite element method. Three cases were studied using the proposed global-local finite element method: a) adhesively bonded Double Cantilever Beam (DCB), b) an adhesive lap joint, and c) a three-point bending test specimen. Using global-local methods, in a crack propagation problem of an adhesively bonded DCB, more than 80% data storage space and more than 65% CPU time requirement could be saved. In the adhesive lap joints, around 70% data storage space and 70% CPU time requirement could be saved using the global-local method. For the three-point bending test specimen case, more than 90% for both data storage space and CPU time requirement could be saved using the global-local method. / Ph. D.