In recent years, use of composites is increasing in most fields of engineering such as aerospace, automotive, civil construction, marine, prosthetics, etc., because of its light weight, very high specific strength and stiffness, corrosion resistance, high thermal resistance etc. It can be seen that the specific strength of fibers are many orders more compared to metals. Thus, laminated fiber reinforced plastics have emerged to be attractive materials for many engineering applications. Though the uses of composites are enormous, there is always an element of fuzziness in the design of composites. Composite structures are required to be designed to resist high stresses. For this, one requires a reliable failure criterion. The anisotropic behaviour of composites makes it very difficult to formulate failure criteria and experimentally verify it, which require one to perform necessary bi-axial tests and plot the failure envelopes. Failure criteria are usually based on certain assumption, which are some times questionable. This is because, the failure process in composites is quite complex. The failure in a composite is normally based on initiating failure mechanisms such as fiber breaks, fiber compressive failure, matrix cracks, matrix crushing, delamination, disbonds or a combination of these. The initiating failure mechanism is the one, which is/are responsible for initiating failure in a laminated composites. Initiating failure mechanisms are generally dependant on the type of loading, geometry, material properties, condition of manufacture, boundary conditions, weather conditions etc. Since, composite materials exhibit directional properties, their applications and failure conditions should be properly examined and in addition to this, robust computational tools have to be exploited for the design of structural components for efficient utilisation of these materials. Design of structural components requires reliable failure criteria for the safe design of the components. Several failure criteria are available for the design of composite laminates. None of the available anisotropic strength criteria represents observed results sufficiently accurate to be employed confidently by itself in design. Most of the failure criteria are validated based on the available uniaxial test data, whereas, in practical situations, laminates are subjected to at least biaxial states of stresses. Since, the generation of biaxial test data are very difficult and time consuming to obtain, it is indeed a necessity to develop computational tools for modelling the biaxial behavior of the composite laminates. Understanding of the initiating failure mechanisms and the development of reliable failure criteria is an essential prerequisite for effective utilization of composite materials. Most of the failure criteria, considers the uniaxial test data with constant shear stress to develop failure envelopes, but in reality, structures are subjected to biaxial normal stresses as well as shear stresses. Hence, one can develop different failure envelopes depending upon the percentage of the shear stress content. As mentioned earlier, safe design of the composite structural components require reliable failure criterion. Currently two broad approaches, namely, (1) Damage Tolerance Based Design and (2)Failure Criteria Based Design are in use for the design of laminated structures in aerospace industry. Both approaches have some limitations. The damage tolerance based design suffers from a lack of proper definition of damage and the inability of analytical tools to handle realistic damage. The failure criteria based design, although relatively, more attractive in view of the simplicity, it forces the designer to use unverified design points in stress space, resulting in unpredictable failure conditions. Generally, failure envelopes are constructed using 4 or 5 experimental constants. In this type of approach, small experimental errors in these constants lead to large shift in the failure boundaries raising doubts about the reliability of the boundary in some segments. Further, they contain segments which have no experimental support and so can lead to either conservative or nonconservative designs. Conservative design leads to extra weight, a situation not acceptable in aerospace industry. Whereas, a nonconservative design, is obviously prohibitive, as it implies failure. Hence, both the damage tolerance based design and failure criteria based design have limitations. A new method, which combines the advantages of both the approaches is desirable. This issue is also thoroughly debated in many international conference on composites. Several pioneers in the composite industry indicated the need for further research work in the development of reliable failure criteria. Hence, this is motivated to carry out research work for the development of new failure criterion for the design of composite structures. Several experts meetings held world wide towards the assessment of existing failure theories and computer codes for the design of composite structures. One such meeting is the experts meeting held at United Kingdom in 1991.This meeting held at St. Albans(UK) on ’Failure of Polymeric Composites and Structures: Mechanisms and Criteria for the Prediction of Performance’, in 1991 by UK Science & Engineering Council and UK Institute of Mechanical Engineers. After thorough deliberations it was concluded that 1. There is no universal definition of failure of composites. 2. There is little or lack of faith in the failure criteria that are in current use and 3. There is a need to carry out World Wide Failure Exercise(WWFE) Based on the experts suggestions, Hinton and Soden initiated the WWFE in consultation with Prof.Bryan Harris (Editor, Journal of Composite Science and Tech-nology)to have a program to get comparative assessment of existing failure criteria and codes with following aims 1. Establish the current level of maturity of theories for predicting the failure response of fiber reinforced plastic(FRP)laminates. 2. Closing the knowledge gap between theoreticians and design practitioners in this field. 3. Stimulating the composites’ community into providing design engineers with more robust and accurate failure prediction methods, and the confidence to use them. The organisers invited pioneers in the composite industry for the program of WWFE. Among the pioneer in the composite industry Professor Hashin declined to participate in the program and had written a letter to the organisers saying that, My only work in this subject relates to failure criteria of unidirectional fiber composites, not to laminates. I do not believe that even the most complete information about failure of single plies is sufficient to predict the failure of a laminate, consisting of such plies. A laminate is a structure which undergoes a complex damage process (mostly of cracking) until it finally fails. The analysis of such a process is a prerequisite for failure analysis. ”While significant advances have been made in this direction we have not yet arrived at the practical goal of failure prediction”. Another important conference held in France in 1999, Composites for the next Millennium (Proceedingof Symposium in honor of S.W.Tsaion his 70th Birth Day Torus, France, July 2-3, 1999, pp.19.) also concludedon similar line to the meeting held at UK in 1991. Paul A Lagace and S. Mark Spearing, have pointed out that, by referring to the article on ’Predicting Failure in Composite Laminates: the background to the exercise’, by M.J.Hinton & P.D.Soden, Composites Science and Technology, Vol.58, No.7(1998), pp.1005. ”After Over thirty years of work ’The’ composite failure criterion is still an elusive entity”. Numerous researchers have produced dozens of approaches. Hundreds of papers, manuscripts and reports were written and presentations made to address the latest thoughts, add data to accumulated knowledge bases and continue the scholarly debate. Thus, the out come of these experts meeting, is that, there is a need to develop new failure theories and due to complexities associated with experimentation, especially getting bi-axial data, computational methods are the only viable alternative. Currently, biaxial data on composites is very limited as the biaxial testing of laminates is very difficult and standardization of biaxial data is yet to be done. All these experts comments and suggestions motivated us to carry out research work towards the development of new failure criterion called ’Failure Mechanism Based Failure Criterion’ based on initiating failure mechanisms. The objectives of the thesis are 1. Identification of the failure mechanism based failure criteria for the specific initiating failure mechanism and to assign the specific failure criteria for specific initiating failure mechanism, 2. Use of the ’failure mechanism based design’ method for composite pressurant tanks and to evaluate it, by comparing it with some of the standard ’failure criteria’ based designs from the point of view of overall weight of the pressurant tank, 3. Development of new failure criterion called ’Failure Mechanism Based Failure Criterion’ without shear stress content and the corresponding failure envelope, 4. Development of different failure envelopes with the effect of shear stress depending upon the percentage of shear stress content and 5. Design of composite laminates with the Failure Mechanism Based Failure Criterion using optimization techniques such as Genetic Algorithms(GA) and Vector Evaluated Particle Swarm Optimization(VEPSO) and the comparison of design with other failure criteria such as Tsai-Wu and Maximum Stress failure criteria. The following paragraphs describe about the achievement of these objectives. In chapter 2, a rectangular panel subjected to boundary displacements is used as an example to illustrate the concept of failure mechanism based design. Composite Laminates are generally designed using a failure criteria, based on a set of standard experimental strength values. Failure of composite laminates involves different failure mechanisms depending upon the stress state and so different failure mechanisms become dominant at different points on the failure envelope. Use of a single failure criteria, as is normally done in designing laminates, is unlikely to be satisfactory for all combination of stresses. As an alternate use of a simple failure criteria to identify the dominant failure mechanism and the design of the laminate using appropriate failure mechanism based criteria is suggested in this thesis. A complete 3-D stress analysis has been carried out using a general purpose NISA Finite Element Software. Comparison of results using standard failure criteria such as Maximum Stress, Maximum Strain, Tsai-Wu, Yamada-Sun, Maximum Fiber Strain, Grumman, O’brien, & Lagace, indicate substantial differences in predicting the first ply failure. Results for Failure Load Factors, based on the failure mechanism based approach are included. Identification of the failure mechanism at highly stressed regions and the design of the component, to withstand an artificial defect, representative this failure mechanism, provides a realistic approach to achieve necessary strength without adding unnecessary weight to the structure. It is indicated that the failure mechanism based design approach offers a reliable way of assessing critically stressed regions to eliminate the uncertainties associated with the failure criteria. In chapter 3, the failure mechanism based design approach has been applied to a composite pressurant tanks of upper stages of launch vehicles and propulsion systems of space crafts. The problem is studied using the failure mechanism based design approach, by introducing an artificial matrix crack representative of the initiating failure mechanism in the highly stressed regions and the strain energy release rate (SERR) are calculated. The total SERR value is obtained as 3330.23 J/m2, which is very high compared to the Gc(135 J/m2) value, which means the crack will grow further. The failure load fraction at which the crack has a tendency to grow is estimated to be 0.04054.Results indicates that there are significant differences in the failure load fraction for different failure criteria.Comparison with Failure Mechanism Based Criterion (FMBC) clearly indicates matrix cracks occur at loads much below the design load yet fibers are able to carrythe design load. In chapter 4, a Failure Mechanism Based Failure Criterion(FMBFC)has been proposed for the development of failure envelope for unidirectional composite plies. A representative volume element of the laminate under local loading is micromechanically modelled to predict the experimentally determined strengths and this model is then used to predict points on the failure envelope in the neighborhood of the experimental points. The NISA finite element software has been used to determine the stresses in the representative volume element. From these micro-stresses, the strength of the lamina is predicted. A correction factor is used to match the prediction of the present model with the experimentally determined strength so that, the model can be expected to provide accurate prediction of the strength in the neighborhood of the experimental points. A procedure for the construction of the failure envelope in the stress space has been outlined and the results are compared with the some of the standard and widely used failure criteria in the composite industry. Comparison of results with the Tsai-Wu failure criterion shows that there are significant differences, particularly in the third quadrant, when the ply is under bi-axial compressive loading. Comparison with maximum stress criterion indicates better correlation. The present failure mechanism based failure criterion approach opens a new possibility of constructing reliable failure envelopes for bi-axial loading applications, using the standard uniaxialtest data. In chapter 5, the new failure criterion for laminated composites developed based on initiating failure mechanism as mentioned in chapter 4 for without shear stress condition is extended to obtain the failure envelopes with the shear stress condition. The approach is based on Micromechanical analysis of composites, wherein a representative volume consists of a fiber surrounded by matrix in appropriate volume fraction and modeled using 3-D finite elements to predict the strengths.In this chapter, different failure envelopes are developed by varying shear stress say from 0% of shear strength to 100% of shear strength in steps of 25% of shear strength. Results obtained from this approach are compared with Tsai-Wu and Maximum stress failure criteria. The results show that the predicted strengths match more closely with maximum stress criterion. Hence, it can be concluded that influence of shear stress on the failure of the lamina is of little consequence as far as the prediction of strengths in laminates. In chapter 6, the failure mechanism based failure criterion, developed by the authors is used for the design optimization of the laminates and the percentage savings in total weight of the laminate is presented. The design optimization of composite laminates are performed using Genetic Algorithms. The genetic algorithm is one of the robust tools available for the optimum design of composite laminates. The genetic algorithms employ techniques originated from biology and dependon the application of Darwin’s principle of survival of the fittest. When a population of biological creatures is permitted to evolve over generations, individual characteristics that are beneficial for survival have a tendency to be passed on to future generations, since individuals carrying them get more chances to breed. In biological populations, these characteristics are stored in chromosomal strings. The mechanics of natural genetics is derived from operations that result in arranged yet randomized exchange of genetic information between the chromosomal strings of the reproducing parents and consists of reproduction, cross over, mutation, and inversion of the chromosomal strings. Here, optimization of the weight of the composite laminates for given loading and material properties is considered. The genetic algorithms have the capability of selecting choice of orientation, thickness of single ply, number of plies and stacking sequence of the layers. In this chapter, minimum weight design of composite laminates is presented using the Failure Mechanism Based(FMB), Maximum Stress and Tsai-Wu failure criteria. The objective is to demonstrate the effectiveness of the newly proposed FMB Failure Criterion(FMBFC) in composite design. The FMBFC considers different failure mechanisms such as fiber breaks, matrix cracks, fiber compressive failure, and matrix crushing which are relevant for different loadin gconditions. FMB and Maximum Stress failure criteria predicts byupto 43 percent savings in weight of the laminates compared to Tsai-Wu failure criterion in some quadrants of the failure envelope. The Tsai-Wu failure criterion over predicts the weight of the laminate by up to 86 percent in the third quadrant of the failure envelope compared to FMB and Maximum Stress failure criteria, when the laminate is subjected to biaxial compressive loading. It is found that the FMB and Maximum Stress failure criteria give comparable weight estimates. The FMBFC can be considered for use in the strength design of composite structures. In chapter 7, Particle swarm optimization is used for design optimization of composite laminates. Particle swarm optimization(PSO)is a novel meta-heuristic inspired by the flocking behaviour of birds. The application of PSO to composite design optimization problems has not yet been extensively explored. Composite laminate optimization typically consists in determining the number of layers, stacking sequence and thickness of ply that gives the desired properties. This chapter details the use of Vector Evaluated Particle Swarm Optimization(VEPSO) algorithm, a multi-objective variant of PSO for composite laminate design optimization. VEPSO is a modern coevolutionary algorithm which employs multiple swarms to handle the multiple objectives and the information migration between these swarms ensures that a global optimum solution is reached. The current problem has been formulated as a classical multi-objective optimization problem, with objectives of minimizing weight of the component for a required strength and minimizing the totalcost incurred, such that the component does not fail. In this chapter, an optimum configuration for a multi-layered unidirectional carbon/epoxy laminate is determined using VEPSO. The results are presented for various loading configurations of the composite structures. The VEPSO predicts the same minimum weight optimization and percentage savings in weight of the laminate when compared to GA for all loading conditions.There is small difference in results predicted by VEPSO and GA for some loading and stacking sequence configurations, which is mainly due to random selection of swarm particles and generation of populations re-spectively.The difference can be prevented by running the same programme repeatedly. The Thesis is concluded by highlighting the future scope of several potential applications based on the developments reported in the thesis.

Diversos métodos de ensaios não-destrutivos (Ultrasom, Radiografia, Termografia e Shearografia) foram empregados na inspeção de peças laminadas estruturais compósitas sólidas de matriz polimérica manufaturadas pela indústria aeronáutica. Concluiu-se que Ultrasom pulso-eco convencional de contato foi a técnica mais abrangente na indicação de descontinuidades (danos/defeitos de manufatura) considerando-se peças planas e curvas, não obstante tenha falhado na detecção de trincas longas e bem definidas localizada na alma de componentes co-curados. Radiografia convencional por filmes exibiu um potencial interessante como método alternativo, ou complementar ao de Ultrasom. Termografia infravermelha foi a técnica mais veloz na indicação de descontinuidades, sendo uma valiosa alternativa para um mapeamento rápido em inspeções preliminares seguidas pela aplicação de técnicas complementares. Shearografia realizada com equipamento da empresa Photonita detectou inclusões em peça plana compósita, porém a interpretação dos resultados obtidos em geometrias mais complexas se mostrou duvidosa. O uso de equipamento shearográfico da empresa Dantec indicou posições em regiões curvas que podem estar associadas à presença de defeitos/danos locais, porém uma confirmação cabal desta relação não foi efetivada. Evidências visuais da profusa presença de danos/defeitos de manufatura, além da existência de detalhes construtivos nas peças compósitas mais complexas, sugerem que estas descontinuidades podem ter sensibilizado, em maior ou menor extensão, os equipamentos END, juntamente com a eventual presença de delaminações/falhas de adesão nas respectivas peças avaliadas. / Several nondestructive methods (Ultrasonics, Radiography, Thermography and Shearography) were employed for inspection of structural polymer matrix composite laminated parts manufactured by aircraft industry. Is has been concluded that conventional pulse-echo contact Ultrasonics was the most comprehensive technique to indicate discontinuities (manufacture damages/defects) in flat and curved parts, although it has failed to detect long and well-defined cracks in co-cured components. Conventional film radiography exhibited good potential as alternative or complementary method to Ultrasonics. Infrared thermography was the fastest technique to indicate discontinuities, so that it is a valuable option for rapid mapping in preliminary inspection followed by the application of complementary techniques. Sherography by means of Photonita equipment detected inclusions in flat panels, but the interpretation of results from pieces with more complex geometries was dubious. A shearographic of the Dantec device indicated positions in curved regions which might possibly be associated to the presence of damages/defects, however this relationship could not be definitively established. Visual evidences of the profuse existence of manufacture damages/defects, besides constructive details in more complex composite parts, suggest that theses discontinuities might have affected to some extent the NDT equipments, along with the eventual presence of delaminations/lack of adhesion on the respective evaluated parts.

Estudou-se o comportamento de desgaste sob deslizamento em múltiplos passes de pinos fixos de liga metálica friccionados contra discos rotativos de laminados compósitos de matriz polimérica fortalecidos com fibras contínuas de carbono (C), para distintas cargas aplicadas e temperaturas de ensaio. Determinou-se a perda de volume por desgaste dos diversos pares tribológicos, e avaliaram-se os principais aspectos das superfícies e dos produtos gerados durante o processo de desgaste. Concluiu-se que o par formado pelo aço inoxidável martensítico PH 15-5 e o laminado compósito termoplástico C-PPS (poli-sulfeto de fenileno) apresentou o menor desgaste na maioria das circunstâncias avaliadas. A liga Ti6Al4V foi a pior opção dentre os pinos metálicos, independentemente do laminado compósito empregado, devido à baixa resistência ao cisalhamento e alta rugosidade superficial do metal, a qual prejudicou o ancoramento de filmes poliméricos de transferência. O ótimo desempenho do aço PH resultou de sua alta dureza e da adequada rugosidade superficial para o efetivo ancoramento do filme de transferência. O laminado compósito C-PPS se mostrou mais susceptível às variações na carga de contato, enquanto que o C-EPX (Epóxi) foi mais sensível à temperatura de ensaio. Filmes poliméricos de transferência atuaram como lubrificante do sistema tribológico, preservando ambos o pino metálico e o disco compósito de um desgaste excessivo. A natureza dos produtos de desgaste, sua geometria, seu tamanho médio e sua distribuição de tamanhos dependeram fortemente do disco rotativo de laminado compósito e das condições de ensaio de desgaste, e afetaram sobremaneira, como terceiro corpo abrasivo, o subseqüente processo de desgaste dos pares de fricção metal-compósito. Os resultados obtidos são potencialmente úteis na especificação de juntas mecânicas aeronáuticas mais apropriadas de laminados compósitos de matriz polimérica unidos por intermédio de prendedores metálicos. / Multipass sliding wear behavior of static metal alloy pins rubbing against rotating carbon fiber reinforcing polymer matrix composite laminate disks has been studied for different applied loads and test temperatures. The volume loss due wear has been determined for several tribological metal / composite systems, and the main aspects of wear surfaces and debris have been evaluated. It has been concluded that the best performance was achieved in very most cases by the PH15-5 steel pin contacting the C-PPS (poly-phenylene sulphide) disk. The Ti6Al4V alloy displayed the worst performance among the metal pins, regardless the employed composite laminate, as a result of the metal´s low shear strength and high surface roughness, which impaired the formation of polymer transfer films. The outstanding performance of PH steel pins derived from its high hardness and proper surface roughness that favored the transfer film formation. The C-PPS composite laminate was more sensitive to contact load variations, whereas the C-EPX laminate was more susceptible to the test temperature. Polymer transfer films acted as lubricant in the tribological system, preserving both the metallic pin and the composite disk from wearing in excess. The nature of wear debris, their geometry, mean size, and size distribution strongly depended on the composite laminate disk as well as on testing conditions, and affected overwhelmingly, as third body abrasive particles, the subsequent wear process of metal / composite friction pairs. The obtained results are potentially useful for the specification of more equalized aircraft mechanical joints comprising polymer matrix composite laminates and metallic fasteners.

With the development of the design and manufacture technology, composite materials are widely used in the aeronautical industry. But, one of the main concerns which affects the application of composites is foreign object impact. The damages induced by the Low Velocity Impact (LVI), which can significantly reduce the strength of the structures, can’t be easily inspected routinely. The so-called Barely Visible Impact Damages (BVID) due to LVI typically includes interlaminar delamination, matrix cracks and fibre fracture at the back face. Previous researches have shown that the results of LVI test are similar to that of the Quasi-Static Load (QSL) test. The initiation and propagation of delamination can be detected more easily in the QSL test and the displacement and reaction force of the impactor can be controlled and measured much more accurately. Moreover, it is easier to model QSL tests than dynamic impacts. To investigate the impact damage induced by LVI, a Finite Element (FE) model employing cohesive elements was used. At the same time, the ply clustering effect, when several plies of the same orientation were stack together, was modelled in the FE model in terms of damage resistance and damage size. A bilinear traction-separation law was introduced in the cohesive elements employed to simulate the initiation and propagation of the impact damage and delamination. Firstly, a 2D FE model of the Double Cantilever Beam (DCB) and End Notched Flexure (ENF) specimens were built using the commercial FEM software ABAQUS. The results have shown that the cohesive elements can be used to simulate mode I and mode II delamination sufficiently and correctly. Secondly, an FE model of a composite plate under QSL but without simulating damage was built using the continuum shell elements. Agreement between the FEA results with published test results is good enough to validate the capability of continuum shell elements and cohesive elements in modelling the composite laminate under the transverse load condition (QSL). Thirdly, an FE model containing discrete interface delamination and matrix cracks at the back face of the composite plate was built by pre-setting the cohesive failure elements at potential damage locations according to the experimental observation. A cross-ply laminate was modelled first where fewer interfaces could be delaminated. Good agreement was found in terms of the delamination area and impactor’s displacement-force curve. Finally, the effect of ply clustering on impact damage resistance was studied using Quasi-Isotropic (QI) layup laminates. Because of the limited time available for calculation, the simulation was only partly completed for the quasi-isotropic laminates (L2 configuration) which have more delaminated interfaces. The results showed that cohesive elements obeying the bilinear traction-separation law were capable of predicting the reaction force in quasi-isotropic laminates. However, discrepancies with the test results in terms of delamination area were observed for quasi-isotropic laminates. These discrepancies are mainly attributed to the simplification of matrix cracks simulation and compressive load at the interface in the thickness direction which is not taken into account.

Estudou-se o comportamento de desgaste sob deslizamento em múltiplos passes de pinos fixos de liga metálica friccionados contra discos rotativos de laminados compósitos de matriz polimérica fortalecidos com fibras contínuas de carbono (C), para distintas cargas aplicadas e temperaturas de ensaio. Determinou-se a perda de volume por desgaste dos diversos pares tribológicos, e avaliaram-se os principais aspectos das superfícies e dos produtos gerados durante o processo de desgaste. Concluiu-se que o par formado pelo aço inoxidável martensítico PH 15-5 e o laminado compósito termoplástico C-PPS (poli-sulfeto de fenileno) apresentou o menor desgaste na maioria das circunstâncias avaliadas. A liga Ti6Al4V foi a pior opção dentre os pinos metálicos, independentemente do laminado compósito empregado, devido à baixa resistência ao cisalhamento e alta rugosidade superficial do metal, a qual prejudicou o ancoramento de filmes poliméricos de transferência. O ótimo desempenho do aço PH resultou de sua alta dureza e da adequada rugosidade superficial para o efetivo ancoramento do filme de transferência. O laminado compósito C-PPS se mostrou mais susceptível às variações na carga de contato, enquanto que o C-EPX (Epóxi) foi mais sensível à temperatura de ensaio. Filmes poliméricos de transferência atuaram como lubrificante do sistema tribológico, preservando ambos o pino metálico e o disco compósito de um desgaste excessivo. A natureza dos produtos de desgaste, sua geometria, seu tamanho médio e sua distribuição de tamanhos dependeram fortemente do disco rotativo de laminado compósito e das condições de ensaio de desgaste, e afetaram sobremaneira, como terceiro corpo abrasivo, o subseqüente processo de desgaste dos pares de fricção metal-compósito. Os resultados obtidos são potencialmente úteis na especificação de juntas mecânicas aeronáuticas mais apropriadas de laminados compósitos de matriz polimérica unidos por intermédio de prendedores metálicos. / Multipass sliding wear behavior of static metal alloy pins rubbing against rotating carbon fiber reinforcing polymer matrix composite laminate disks has been studied for different applied loads and test temperatures. The volume loss due wear has been determined for several tribological metal / composite systems, and the main aspects of wear surfaces and debris have been evaluated. It has been concluded that the best performance was achieved in very most cases by the PH15-5 steel pin contacting the C-PPS (poly-phenylene sulphide) disk. The Ti6Al4V alloy displayed the worst performance among the metal pins, regardless the employed composite laminate, as a result of the metal´s low shear strength and high surface roughness, which impaired the formation of polymer transfer films. The outstanding performance of PH steel pins derived from its high hardness and proper surface roughness that favored the transfer film formation. The C-PPS composite laminate was more sensitive to contact load variations, whereas the C-EPX laminate was more susceptible to the test temperature. Polymer transfer films acted as lubricant in the tribological system, preserving both the metallic pin and the composite disk from wearing in excess. The nature of wear debris, their geometry, mean size, and size distribution strongly depended on the composite laminate disk as well as on testing conditions, and affected overwhelmingly, as third body abrasive particles, the subsequent wear process of metal / composite friction pairs. The obtained results are potentially useful for the specification of more equalized aircraft mechanical joints comprising polymer matrix composite laminates and metallic fasteners.

L’assemblage des matériaux composites par collage présente des avantages incontestés par rapport à d’autres méthodes telles que le boulonnage ou le rivetage. Cependant, la principale difficulté que rencontrent les concepteurs est celle de la prévision du niveau et du mode de rupture de ces collages. Dans le cas des composites stratifiés, un des facteurs influents sur le comportement des joints collés est la séquence d'empilement, mais les travaux présentés dans la littérature ne séparent pas les effets globaux (modification des rigidités de membrane et de flexion) et les effets locaux (orientation des plis au contact de la colle). La présente étude s'intéresse à la caractérisation de ces effets dans le cas de joints de type simple recouvrement de stratifiés carbone/epoxy. Pour isoler les effets locaux, des séquences d'empilement spécifiques quasi isotropes quasi homogènes sont utilisées. A propriétés de raideur globale égales, des différences de résistance de plus de 30% sont observées selon les séquences considérées. Les essais réalisés avec un stratifié symétrique anisotrope en flexion montrent également que la raideur en flexion joue un rôle important dans le comportement des joints. Les modèles analytiques utilisés prédisent les effets globaux avec une bonne précision mais sont inappropriés lorsque des phénomènes locaux se produisent. Une approche par éléments finis permet de prendre en compte ces phénomènes, en modélisant explicitement les plis au contact de la colle et en rendant possible le décollement interlaminaire de ces plis à l'aide d'un modèle de zone cohésive. Cette modélisation est mise en œuvre pour réaliser une étude paramétrique de la géométrie du joint et pour produire une enveloppe de rupture en fonction de la direction de sollicitation. / Adhesive bonding of composite materials has undeniable advantages over other methods such as bolting or riveting. However, one of the difficulties encountered by designers is the prediction of the failure level and failure mode of these adhesively bonded assemblies. In the case of composite laminates, one of the factors acting on the bonded joint behaviour is the stacking sequence, but works presented in the literature do not separate global effects (membrane and bending stiffness modification) and local effect (ply orientation near the adhesive layer). This study deals with the characterization of such effects in the case of single lap joints of carbon/epoxy laminates. In order to isolate the local effects, specific quasi isotropic quasi homogeneous stacking sequences are used. When stiffness properties are maintained constant, strength variations of more than 30 % are observed. Tests performed with a symmetric laminate with bending anisotropy show that the bending stiffness plays also an important role in the joint behaviour. Closed form models are able to predict global effects with a good accuracy but are inappropriate when local effects occur. The use of finite element models can account for these phenomena, by explicitly modelling the laminate plies near the adhesive layer and introducing delamination between these plies with a cohesive zone model. This model is used to perform a parametric study of the joint geometry and to produce a failure envelope according to the orientation of the loading.

Diversos métodos de ensaios não-destrutivos (Ultrasom, Radiografia, Termografia e Shearografia) foram empregados na inspeção de peças laminadas estruturais compósitas sólidas de matriz polimérica manufaturadas pela indústria aeronáutica. Concluiu-se que Ultrasom pulso-eco convencional de contato foi a técnica mais abrangente na indicação de descontinuidades (danos/defeitos de manufatura) considerando-se peças planas e curvas, não obstante tenha falhado na detecção de trincas longas e bem definidas localizada na alma de componentes co-curados. Radiografia convencional por filmes exibiu um potencial interessante como método alternativo, ou complementar ao de Ultrasom. Termografia infravermelha foi a técnica mais veloz na indicação de descontinuidades, sendo uma valiosa alternativa para um mapeamento rápido em inspeções preliminares seguidas pela aplicação de técnicas complementares. Shearografia realizada com equipamento da empresa Photonita detectou inclusões em peça plana compósita, porém a interpretação dos resultados obtidos em geometrias mais complexas se mostrou duvidosa. O uso de equipamento shearográfico da empresa Dantec indicou posições em regiões curvas que podem estar associadas à presença de defeitos/danos locais, porém uma confirmação cabal desta relação não foi efetivada. Evidências visuais da profusa presença de danos/defeitos de manufatura, além da existência de detalhes construtivos nas peças compósitas mais complexas, sugerem que estas descontinuidades podem ter sensibilizado, em maior ou menor extensão, os equipamentos END, juntamente com a eventual presença de delaminações/falhas de adesão nas respectivas peças avaliadas. / Several nondestructive methods (Ultrasonics, Radiography, Thermography and Shearography) were employed for inspection of structural polymer matrix composite laminated parts manufactured by aircraft industry. Is has been concluded that conventional pulse-echo contact Ultrasonics was the most comprehensive technique to indicate discontinuities (manufacture damages/defects) in flat and curved parts, although it has failed to detect long and well-defined cracks in co-cured components. Conventional film radiography exhibited good potential as alternative or complementary method to Ultrasonics. Infrared thermography was the fastest technique to indicate discontinuities, so that it is a valuable option for rapid mapping in preliminary inspection followed by the application of complementary techniques. Sherography by means of Photonita equipment detected inclusions in flat panels, but the interpretation of results from pieces with more complex geometries was dubious. A shearographic of the Dantec device indicated positions in curved regions which might possibly be associated to the presence of damages/defects, however this relationship could not be definitively established. Visual evidences of the profuse existence of manufacture damages/defects, besides constructive details in more complex composite parts, suggest that theses discontinuities might have affected to some extent the NDT equipments, along with the eventual presence of delaminations/lack of adhesion on the respective evaluated parts.

Diversos critérios de tenacidade à fratura translaminar dinâmica foram determinados para o laminado híbrido metal-fibra TiGra, empregando-se conceitos e metodologias da Mecânica da Fratura Elástica Linear MFEL (fator-K) e da Mecânica da Fratura Elasto-Plástica MFEP (integral-J). Verificou-se que as tenacidades de iniciação elasto-plástica, Jid, e de carga máxima, Jmd, do TiGra são controladas pelo desenvolvimento, ou supressão de delaminações. Os resultados indicaram que o emprego deste material se justifica mais pela sua resistência à propagação de danos (caracterizada por Jmd) do que à iniciação da fratura dinâmica (por Jid). De modo geral, os requisitos de validade de Jid como verdadeira propriedade do material (JId) foram satisfeitos, embora para Jmd boa parte das restrições quanto ao tamanho mínimo do corpo-de-prova tenha sido violada. Mais freqüentemente, velocidades mais rápidas de impacto beneficiaram as tenacidades-J do TiGra, enquanto que temperaturas mais elevadas afetaram negativamente estas propriedades. Quanto à MFEL, a tenacidade KJd do TiGra foi beneficiada pelo incremento na taxa de carregamento sob temperaturas mais elevadas, enquanto que a tenacidade Kid foi negativamente afetada pela taxa de deformação em todas as temperaturas avaliadas. Temperaturas mais altas também degradaram as propriedades de tenacidade-K do TiGra. Em oposição às tenacidades-J, os critérios KJd e Kid não satisfizeram em absoluto os mais exigentes critérios de contenção de plasticidade estabelecidos pela MFEL, se comparados aos propostos pela MFEP. Por fim, o desempenho mecânico do laminado TiGra foi severamente comprometido quando do cômputo da densidade específica para a determinação das tenacidades J e K por unidade de massa, sendo nesta ocasião o laminado híbrido facilmente superado por vários laminados convencionais da classe dos Carbono-Epóxi. / Several dynamic translaminar fracture toughness criteria have been determined for TiGr hybrid fiber-metal laminate through Linear Elastic (K-factor) and Elastic-Plastic (J-integral) Fracture Mechanics (LEFM and EPFM, respectively) concepts and methodologies. Instrumented Charpy impact testing was carried out over a wide range of temperatures under two loading rates. It has been discovered that the elastic-plastic initiation toughness, Jid, and the toughness at maximum load, Jmd, of TiGr are controlled by either delamination favoring or suppression. Impact tests revealed that the in-service use of TiGr must rely on its resistance to dynamic fracture propagation (as characterized by Jmd) rather than on fracture initiation (by Jid). In a broad sense, the requirements for Jid data validity as a material property (JId) were fulfilled, whereas many restrictive demands in regard to the minimum testpiece size were violated by the Jmd criterion. Generally, higher impact velocities were beneficial to TiGrs J-toughnesses, inasmuch as higher temperatures impaired these properties. Regarding the LEFM approach, KJd toughness of TiGr laminate was imparted by faster impacts at higher temperatures, whilst the strain rate negatively influenced the Kid toughness over the whole temperature range tested. Higher temperatures also degraded the K-toughness properties of TiGr hybrid laminate. Differently from J-toughnesses values, the KJd e Kid criteria did not satisfy at all the more stringent criteria set forth by the LEFM approach with regard to plastic constraint, as compared to those established by EPFM. Finally, the mechanical performance of TiGr laminate was overwhelmingly compromised as the materials specific gravity was taken in account to obtain K and J toughness values by unit weight, so that TiGr was by far exceeded in this regard by conventional Carbon/Epoxy composite laminates.

Composite laminates are an important subject of modern technology and engineering. The most common mode of failure in these materials is probably interlaminar fracture (delamination). Delamination growth under applied fatigue loads usually leads to structural integrity loss of the composite laminate, and hence its catastrophic failure. It is known that several parameters can affect the fatigue fracture performance of laminates. These include the constituent material properties, composite geometry, fatigue load variables or environmental factors. The knowledge about effects of these parameters on fatigue delamination growth can lead to a better understanding of composite fatigue fracture behaviour. Effects of some of these parameters can be elucidated by undertaking appropriate sensitivity analysis combined with the finite element method (FEM) and related software. The purpose of this work was three-fold. The first goal was the elaboration and computational implementation of FEM-based numerical strategies for the sensitivity analysis of interface fatigue crack propagation in elastic composite laminates. The second goal of this work was the numerical determination and investigation of displacement and stress fields near the crack tip, contact pressures along crack surfaces, mixed mode angle, energy release rate and the number of cumulative fatigue cycles. The third aim of the present study was to use the developed strategies to evaluate numerically the sensitivity gradients of the total energy release rate and fatigue life with respect to design variables of the curved boron/epoxy-aluminium (B/Ep-Al) composite laminate in two different material configurations under cyclic shear of constant amplitude. This study provided novel strategies for undertaking sensitivity analysis of the delamination growth under fatigue loads for elastic composite laminates using the package ANSYS. The numerical results of the work shed more light on mechanisms of interfacial crack propagation under cyclic shear in the case of a curved B/Ep-Al composite laminate. Moreover, the outcome of the sensitivity gradients demonstrated some advantages for using the sensitivity analysis to pinpoint directions for the optimisation of fatigue fracture performance of elastic laminates. The strategies proposed in this work can be used to study the sensitivity of the interface fatigue crack propagation in other elastic laminates, if the crack propagates at the interface between the elastic and isotropic components. However, the strategies can be potentially extended to composites with interfacial cracks propagating between two non-isotropic constituents under a constant amplitude fatigue load. Finally, the strategies can also be used to undertake the sensitivity analysis of composite fatigue life with respect to variables of fatigue load.

Diversos critérios de tenacidade à fratura translaminar dinâmica foram determinados para o laminado híbrido metal-fibra TiGra, empregando-se conceitos e metodologias da Mecânica da Fratura Elástica Linear MFEL (fator-K) e da Mecânica da Fratura Elasto-Plástica MFEP (integral-J). Verificou-se que as tenacidades de iniciação elasto-plástica, Jid, e de carga máxima, Jmd, do TiGra são controladas pelo desenvolvimento, ou supressão de delaminações. Os resultados indicaram que o emprego deste material se justifica mais pela sua resistência à propagação de danos (caracterizada por Jmd) do que à iniciação da fratura dinâmica (por Jid). De modo geral, os requisitos de validade de Jid como verdadeira propriedade do material (JId) foram satisfeitos, embora para Jmd boa parte das restrições quanto ao tamanho mínimo do corpo-de-prova tenha sido violada. Mais freqüentemente, velocidades mais rápidas de impacto beneficiaram as tenacidades-J do TiGra, enquanto que temperaturas mais elevadas afetaram negativamente estas propriedades. Quanto à MFEL, a tenacidade KJd do TiGra foi beneficiada pelo incremento na taxa de carregamento sob temperaturas mais elevadas, enquanto que a tenacidade Kid foi negativamente afetada pela taxa de deformação em todas as temperaturas avaliadas. Temperaturas mais altas também degradaram as propriedades de tenacidade-K do TiGra. Em oposição às tenacidades-J, os critérios KJd e Kid não satisfizeram em absoluto os mais exigentes critérios de contenção de plasticidade estabelecidos pela MFEL, se comparados aos propostos pela MFEP. Por fim, o desempenho mecânico do laminado TiGra foi severamente comprometido quando do cômputo da densidade específica para a determinação das tenacidades J e K por unidade de massa, sendo nesta ocasião o laminado híbrido facilmente superado por vários laminados convencionais da classe dos Carbono-Epóxi. / Several dynamic translaminar fracture toughness criteria have been determined for TiGr hybrid fiber-metal laminate through Linear Elastic (K-factor) and Elastic-Plastic (J-integral) Fracture Mechanics (LEFM and EPFM, respectively) concepts and methodologies. Instrumented Charpy impact testing was carried out over a wide range of temperatures under two loading rates. It has been discovered that the elastic-plastic initiation toughness, Jid, and the toughness at maximum load, Jmd, of TiGr are controlled by either delamination favoring or suppression. Impact tests revealed that the in-service use of TiGr must rely on its resistance to dynamic fracture propagation (as characterized by Jmd) rather than on fracture initiation (by Jid). In a broad sense, the requirements for Jid data validity as a material property (JId) were fulfilled, whereas many restrictive demands in regard to the minimum testpiece size were violated by the Jmd criterion. Generally, higher impact velocities were beneficial to TiGrs J-toughnesses, inasmuch as higher temperatures impaired these properties. Regarding the LEFM approach, KJd toughness of TiGr laminate was imparted by faster impacts at higher temperatures, whilst the strain rate negatively influenced the Kid toughness over the whole temperature range tested. Higher temperatures also degraded the K-toughness properties of TiGr hybrid laminate. Differently from J-toughnesses values, the KJd e Kid criteria did not satisfy at all the more stringent criteria set forth by the LEFM approach with regard to plastic constraint, as compared to those established by EPFM. Finally, the mechanical performance of TiGr laminate was overwhelmingly compromised as the materials specific gravity was taken in account to obtain K and J toughness values by unit weight, so that TiGr was by far exceeded in this regard by conventional Carbon/Epoxy composite laminates.