7th ECCOMAS Thematic Conference on the Mechanical Response of Composites COMPOSITES 2019 A. Turon, P. Maimí & M. Fagerström (Editors) FATIGUE FRACTURE BEHAVIOR OF DISSIMILAR METAL- COMPOSITE ADHESIVE JOINTS FOR AEROSPACE APPLICATIONS: AN EXPERIMENTAL STUDY Theodoros Loutas 1* , Lucas Adamos 1 , Panayiotis Tsokanas 1 , Federico Martin de la Escalera 2 , Yasser Essa 2 1 Laboratory of Applied Mechanics and Vibrations, Department of Mechanical Engineering and Aeronautics, University of Patras, Patras University Campus, GR-26504 Rio-Patras, Greece 2 AERNNOVA Engineering Division, Research and Technology Dept., E-28034 Madrid, Spain * thloutas@upatras.gr Key words: Metal-composite adhesive joint, Fatigue fracture behavior, Aerospace applications. Summary: In the present work, the mode I and II fatigue fracture performance of titanium- CFRP adhesive joints is investigated. The adhesive joint under study is composed from two thin adherents, one titanium sheet and one CFRP laminate, and is reinforced from both sides with two thick aluminum beams to increase its flexural stiffness and ensure the non-yielding of the metallic parts. The manufacturing process as well as the intended aerospace application of the joint are presented in our previous works. Here, the dynamic interfacial fracture resistance is in the focus and aspects such as the loading that defines the crack propagation threshold as well as the determination of an appropriate Paris law are under investigation. The fatigue crack growth rate (FCG) / is determined with two methods; i) through visual inspections with a special camera in mode I, and ii) through an effective crack length which utilizes the changes in experimental compliance in mode II. The equation of compliance over crack length is predetermined by a procedure known as the compliance calibration method (CCM). Detailed results are given, and useful conclusions are obtained for the fatigue crack propagation rates in such complex metal-composite adhesive joints. 1 INTRODUCTION The use of composites in primary aviation structures has increased the need for a more reliable structural design. As it is known, composite materials are inherent in various types of damage such as fiber breakage, delamination as well as microcracks in the matrix. Since damage in composite materials cannot be avoided, the structures must be designed to function properly despite a failure, a design philosophy known as damage tolerance. Delamination is one of the most common and serious types of damage in composites. The strength and stiffness properties of a composite laminate with cracks are progressively degraded by increasing the crack length, possibly leading to a macroscopic/large-scale failure. Some common causes for delamination damage are poor manufacturing of the composite, slow cutting tool speeds, impact loading and fatigue. Among these, fatigue loading is a major reason that increases the length of the delamination and thus is taken into account considerably during design. For the