Cette étude présente deux objectifs : industriellement elle contribue à la certification d’unaéronef léger de quatre places (MCR 4S) suivant la norme CS 23 régie par l’EASA(European Aviation Safety Agency) et scientifiquement, elle cherche à mieux appréhenderle lien entre la variabilité des matériaux et du comportement des structures. L’étude devariabilité montre une grande dispersion des caractéristiques mécaniques telle qu’il nousest impossible d’optimiser le coefficient de sécurité. A partir de ce constat, une étude plusapprofondie est effectuée et a permis de déterminer les origines de la variabilité à savoir, lanon homogénéité de l’épaisseur et donc de la teneur volumique de fibre dus au processusde fabrication utilisé (moulage au contact).A partir d’une instrumentation par jauges de déformation sur des zones ciblées de lavoilure d’un aéronef prototype, un essai en vol est effectué dans le but d’extraire un spectred’utilisation en cours de service et ce dans les différentes configurations de vol (roulage,décollage, montée, vol en palier, descente, virage et atterrissage). Ce spectre est utilisé parla suite dans un essai de fatigue sur l’aile afin de comprendre les mécanismesd’endommagement dans ce cas de figure. Conjointement, un modèle numérique de lavoilure validé par des essais statiques sur la voilure a permis de mieux comprendre lecomportement de la voilure grâce à une cartographie détaillée des contraintes. / This study has two objectives: industrially it contributes to the certification of a four-seaterlight aircraft (MCR 4S) following the CS 23 standard governed by EASA (EuropeanAviation Safety Agency) and scientifically, it searches to better understand the linkbetween the variability of materials and structural behavior. The study of variability showsa wide dispersion of mechanical characteristics such that is impossible to optimize thesafety factor. From this, further study is conducted and identified the origins of variabilityi.e., non homogeneity of the thickness and therefore the fiber volume content due to themanufacturing process used (contact molding).From instrumentation by strain gages on targeted areas of the wing, a flight test isperformed in order to extract a spectrum of use in service on different configurations flight(taxiing, takeoff, climb, level flight, descent, turn and landing). This spectrum issubsequently used in a fatigue test on the wing in order to understand the damagemechanisms in this case. Together, a numerical model of the wing validated by static testson the wing helped to better understand the behavior of the wing through a detailed stressdistribution.

Fibre Reinforced Polymer Composites are extensively used in aircraft structures because of its high specific stiffness, high specific strength and tailorability. Though Fibre Reinforced Polymers offer many advantages, they are not free from problems. The damage of different nature, e.g., service mechanical damages, fatigue damage or environmental damage can be observed during operating conditions. Among all the damages, manufacturing or service induced, delamination related damage is the most important failure mechanisms of aircraft-composite structures and can be detrimental for safety. Delamination growth under fatigue loading may take place due to local buckling, growth from free edges and notches such as holes, growth from ply-drops and impact damaged composites containing considerable delamination. Delamination growth can also occur due to interlaminar stresses, which can arise in complex structures due to unanticipated loading. The complex nature of composite failure, involving different failure modes and their interactions, makes it necessary to characterize/identify the relevant parameters for fatigue damage resistance, accumulation and life prediction. An effort has been made in this thesis to understand the fatigue behavior of carbon fibre reinforced epoxy laminates under aircraft wing service loading conditions. The study was made on laminates with different lay-up sequences (quasi-isotropic and fibre dominated) and different geometries (plain specimen, specimen with a hole and ply-drop specimen). The fatigue behaviour of the composite was analyzed by following methods: . Ultrasonic C-Scan was used to characterize the delamination growth. . Dynamic Mechanical Analysis (DMA) was done to study the interfacial degradation due to fatigue loading. In this analysis, the interfacial strength indicator and interfacial damping were calculated. The DMA also provides the storage modulus degradation under fatigue loading. . Scanning electron microscope examination was carried out to understand the fatigue damage mechanisms. . A semi-empirical phenomenological model was also used to estimate the residual fatigue life. This research work reveals that the Carbon Fibre Reinforced Polymer laminates are in the safe limit under service loading conditions, except the specimen with a hole. The specimen with a hole showed delaminations around the hole due to stress concentration and higher interlaminar stresses at the hole edges and this delamination is found to be associated with fibre breakage and fibre pullout. The quasi-isotropic laminate is found to show poorer fatigue behaviour when compared to fibre dominated laminate and ply-drop also shows poor performance due to high stress concentration in the ply-drop region.