Integrated and bonded unitized stiffened composite structures are increasingly being considered for application in aircraft wing and fuselage panels design aimed to increase the local panel's stiffness and to reduce the total structural weight. Panels in the wing skins and fuselage are generally subjected to in-plane normal axial and shear loads. This paper presents an efficient finite element approach to study the vibration response of a stiffened composite panel attached with curved composite stiffeners in the presence of the in-plane axial and shear loads. A first-order shear-deformable theory is employed for modeling the response of both the panel and the stiffeners. Displacement compatibility conditions are imposed at the panel-stiffener interfaces. First, both the convergence and validation studies related to the free vibration modes of the curvilinearly stiffened, isotropic panels, and stiffened composite panels have been conducted. Next, we study the vibration responses of the stiffened composite panels with arbitrarily shaped composite stiffeners in the presence of the in-plane uniform biaxial and shear loads. Several parametric studies are conducted to investigate the influence of the stiffener shape including the stiffener geometric curvature and the stiffener placement, the stiffener depth ratio (height-to-width ratio) and the in-plane load factor on the panel vibrational response.
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