When the wing is treated as a composite sandwich plate and it is assumed that the wing chordwise section is rigid, mathematical formulations for the stiffened composite multicell wing structures are provided, and associated governing equations for the aeroelastic divergence are also derived by a direct approach. When the matrix notation is used, the system of equations is written into an explicit and simple mathematical expression that can then be solved explicitly by using the technique of Laplace integral transform. In this modeling, the composite wing skins and stringers (including the spar flanges) are simulated as the faces, whereas the spar webs and ribs are simulated as the core of the sandwiches. Because the wing cross section must have a streamline shape, unlike the usual uniform thickness modeling, it is more appropriate to simulate the wings as variable thickness sandwiches where the thickness is a function of the airfoil. Moreover, as is usual for assumptions for the sandwich plates, the effects of the transverse shear deformation are also considered. Because several special conditions have been studied in the literature, we first compare our solutions with some existing solutions. To show the generality, several illustrative examples are given to consider the effects of spars, skins (including the ply orientation and stacking sequence), stringers, swept angles, aspect ratio, shape of airfoil, and the warping restraints on the divergence dynamic pressures and the lift loads redistribution.
All Science Journal Classification (ASJC) codes
- Aerospace Engineering