Composite materials together with adhesive-bonding have been increasingly used in the aviation industry. Delamination is among the critical failure modes in fiber-reinforced laminated composite structures including adhesive-bonded assemblies. This thesis presents an analytical approach by taking into account the first-ply failure in adhesive-bonded composite joints subjected to axial tension. The ASTM D3165 standard test specimen geometry is followed for model development derivations. The field equations, in terms of displacements within the joint, are formulated by using the first-order, shear-deformable, laminated plate theory together with kinematics relations and force equilibrium conditions. The stress distributions for the adherends and adhesive are determined after the appropriate boundary and loading conditions are applied and the equations for the field displacements are solved. The equivalent forces at the tip of the prescribed interlaminar crack are obtained based on interlaminar stress distributions. The strain energy release rate of the crack is then determined by using the virtual crack closure technique (VCCT). The system of second-order differential field equations is solved to provide the adherend and adhesive stresses using the symbolic computation tool, Maple 9.52. Finite element analyses using the J-integral as well as the VCCT are performed to verify the developed analytical model. Finite element analyses are conducted using the commercial finite element analysis software ABAQUS 6.5-1. Results determined using the analytical method are shown to correlate well with the results from the finite element analyses.