Aeronautical industry has primarily relied on the p-k method for aeroelastic damping extrapolation. The p-k method can be numerically unstable in regions where modes are close to one another, and repeated iterations of the eigenproblem are required making the approach expensive. In this paper, an entirely new approach for aeroelastic damping estimation is proposed. It uses the Dynamic Eigen Decomposition to find flutter points adjusted by an artificially imposed structural damping. Based on the new aeroelastic equations of motion modified with the artificial structural damping, neutrally stable solutions are found. The amount of the negative structural damping added is interpreted as the true aeroelastic damping at the subcritical point. This analysis is carried out using the Nyquist stability criterion applied to the perturbed system from a nominal stable condition. For demonstration, Goland wing model with six structural modes at Mach=.7 is examined. It is shown that the proposed method can yield aeroelastic damping of critical modes, i.e., lightly damped modes, accurately close to the results of the p-k iterations without causing the mode tracking issue.

An approximate method is developed to investigate the interlaminar stresses near the free edges of beam-type composite laminate structures subjected to out-of-plane shear/bending. The method is based upon admissible function representations for in-plane stresses that contain a linear variation in the longitudinal direction. Closed-form solutions of all stress components are sought by minimizing the complementary energy with respect to the unknown functions. The resulting solutions satisfy the stress equilibrium, compatibility, and all of the boundary conditions. Numerical examples are given for both cross-ply and angle-ply laminates. It is found that interlaminar stresses under the shear/bending, particularly those for angle-ply laminates, may exhibit substan-tially different characteristics than under uniaxial loading or under pure bending. Nomenclature b = half-laminate width /i, /2 = unknown functions for the ply stresses cr22 and o\\2 g(k) = unknown function for the ply stress an in the fcth ply h = laminate thickness h{k \\ h2k} = unknown functions for the strain en in the &th ply L = laminate length AW 1,2,3 = eigenvalue solutions to the characteristic equation (64) n = total number of plies P(XI)- arbitrary loading distributed in the longitudinal direction x \\ (force/unit area) Sjj = ply compliances /(Ar) = thickness of the fcth ply V- applied out-of-plane shear per unit width or laminate volume (force/unit width) X [ = longitudinal coordinate x2 = transverse coordinate x2 = normalized transverse coordinate AT3 = out-of-plane coordinate with origin at the bottom of each ply ci i » €22 * €33 = normal strains 723, To, 712 = shear strains ^23, v\\3, ^12 = poisson ratios Hc = complementary energy HC = complementary energy defined by Eq. (58) a/y = far-field stresses at = companion stresses £ = longitudinal location of shear load