This study dealt with the flutter and biaxial buckling of composite sandwich panels based on a higher order theory. The formulation was based on an enhanced higher order sandwich panel theory that the vertical displacement component of the face sheets were assumed as quadratic one while a cubic pattern was used for the in-plane displacement components of the face sheets and the all displacement components of the core. The transverse normal stress in the face sheets and the in-plane stresses in the core were considered. For the first time, the continuity conditions of the displacements, transverse shear and normal stress at the layer interfaces, as well as the conditions of zero transverse shear stresses on the upper and lower surfaces of the sandwich panel are simultaneously satisfied. The aerodynamic loading was obtained by the first-order piston theory. The equations of motion and boundary conditions were derived via the Hamilton principle. Moreover, effects of some important parameters like lay-up of the face sheets, length to width ratio, length to panel thickness ratio, thickness ratio of the face sheets to panel, fiber angle, elastic modulus ratio and thickness ratio of the face sheets on the stability boundaries were investigated. The results were validated by those published in the literature. The results revealed that by increasing length to width ratio, length to panel thickness ratio and elastic modulus ratio of the face sheets, the stability boundaries were decreased and the largest nondimensional buckling load was occurred at the angle ply sandwich panel.