DAMAGE SIMULATION IN COMPOSITE STRUCTURES UNDER TENSILE LOAD CONSIDERING FRICTION AND SLIDING EFFECT Gregório Felipe Oliveira Ferreira Ricardo de Medeiros José Luis Ogea Volnei Tita Department of Aeronautic Engineering, University of São Paulo – São Carlos School of Engineering, São Carlos, Brazil gregferr@sc. usp.br; medeiros@sc.usp.br ; joseluisogea@gmail.com ; voltita@sc.usp.br. Abstract. Recent improvements in manufacturing processes and material properties associated to excellent mechanical characteristics and low weight have become composite materials very attractive for application on civil aircraft structures. However, even new designs are still very conservative, because the composite structure failure phenomena are very complex. So, it is strategic to know better and to predict these complex failure mechanisms, developing more accurate material models, which reduce the number of experimental tests and be able to simulate the most of characteristics of physic phenomena. Thus, this work focuses on the development and application of a material model based on Continuum Damage Mechanics to simulate the intralaminar progressive failure of a composite laminate structure under tensile test, considering the sliding effect and friction between the metallic clamping system and the composite structure. The proposed damage model was implemented as a UMAT (User Material Subroutine), which was linked to ABAQUS TM . Several numerical analyses were performed via Finite Element Method in order to predict the damage in the laminate structures under those condition by using the implemented UMAT. Moreover, some experiments were carried out in order to identify the material model parameters, and other tests were performed to evaluate the potentialities and limitation of the proposed material model, as well. Keywords: composite materials, material model, progressive failure analyses, modelling via Finite Element Method. 1. INTRODUCTION During the last decades, fiber reinforced polymer composite materials have aroused great interest within the aircraft and aerospace industry, where the design of structures requires optimal stiffness/weight and strength/weight ratio. Composite materials enable the reduction of the structural weight of aircrafts, as well as they increases payload and aircraft range. In spite of the designers’ efforts to use composites within the aircraft industry, the application of those materials on civil aircraft structures is still limited due to the certification process and the service life (Zhou and Gao, 2012). The difficulty of developing material models to study the behavior of laminate composite structures consists on estimating their complex failure mechanisms. Due to heterogeneity and anisotropy, different failure mechanisms are present in a micro structural level (fiber/matrix), and they occur simultaneously before structure collapse. In general terms, failure/damage on laminate composites can be classified into two main types: • Intralaminar damage/failure: damage occurs inside the plies in fibers, matrixes and/or fiber- matrix interfaces (Figure 1-a); • Interlaminar failure: delamination occurs between adjacent plies, i.e. separation of layers (Figure 1-b). According to Anderson (1995), composites, which show weak fiber-matrix interaction, presents interface failure, creating fiber-matrix debonding as shown in Figure 1-a (mechanism (3)). However, composites, which shows strong fiber-matrix interaction, presents probably fiber breakage, causing pull- out mechanism, as seen in Figure 1-a (mechanism (1)). This mechanism is characterized by a process