The increasing need for automatic mesh generation has led to the development of efficient triangulation algorithms that are able to discretize any 2D or 3D domain. Modern finite element formulations based on strain smoothing techniques (SFEM) provide enhanced convergence properties, preventing yet the stiffening behavior of triangular meshes. Recent research has shown that meshless methods based on triangular mapping of the integration domain can be used to produce even better convergence properties than SFEM. The present study explores the Edge-based Smoothed Point Interpolation Method (ES-PIM) as a meshless solution to investigate linear buckling on variable angle tow (VAT) laminates. Such advanced composite structures show a heterogeneous distribution of constitutive properties and thickness, presenting additional challenges to the numerical solution. Important aspects related to the transverse shear correction herein adopted are investigated, leading to interesting conclusions regarding the possibility to use the ES-PIM for conservative estimates of the critical buckling load of VAT laminates.@en

Flutter in aeronautical panels is a type of self-excited oscillation which can occur during supersonic flights. At the flutter point the vibrations of the panel become unstable and increase significantly in time. This manuscript presents a semi-analytical model taking into account the stiffener's base effects, in order to predict the aeroelastic response of laminated composite stiffened panels under supersonic flow. Krumhaar's modified supersonic piston theory, which considers the radius effect, is adopted to model the aerodynamic loading. The proposed model has been validated against results available in the literature for various configurations. A parametric study considering different panels and stiffener configurations is also presented. The numerical results indicate that the stiffener base significantly affects the panel aeroelastic behavior. Preliminary studies also indicate that redistributing the laminate plies from the stiffener's flange to its base significantly increases the torsion stiffness of the panel locally, opening new design possibilities that may lead to higher critical flutter speeds and therefore to better designs. The results also indicate that designs with plies distributed on the base may lead to a better flutter performance when the airflow is transverse to the longitudinal stiffener direction.@en

Flutter in aeronautical panels is a type of self-excited oscillation which can occur during supersonic flights. At the flutter point the vibrations of the panel become unstable and increase significantly in time. This manuscript presents a semi-analytical model taking into account the stiffener's base effects, in order to predict the aeroelastic response of laminated composite stiffened panels under supersonic flow. Krumhaar's modified supersonic piston theory, which considers the radius effect, is adopted to model the aerodynamic loading. The proposed model has been validated against results available in the literature for various configurations. A parametric study considering different panels and stiffener configurations is also presented. The numerical results indicate that the stiffener base significantly affects the panel aeroelastic behavior. Preliminary studies also indicate that redistributing the laminate plies from the stiffener's flange to its base significantly increases the torsion stiffness of the panel locally, opening new design possibilities that may lead to higher critical flutter speeds and therefore to better designs. The results also indicate that designs with plies distributed on the base may lead to a better flutter performance when the airflow is transverse to the longitudinal stiffener direction.@en

The supersonic aeroelastic stability of tow-steered carbon reinforced composite panels, in each layer of which the fibers follow curvilinear paths, is assessed.Astructural model based on the Rayleigh–Ritz method, combined with the aerodynamic piston theory, is derived to represent the aeroelastic behavior of rectangular plates under different boundary conditions. In this model, the classical lamination theory, considering a symmetric stacking sequence and fiber trajectories described by Lagrange polynomials of different orders, is used. In addition, manufacturing constraints, which impose limitations to the feasible fiber trajectories, and the effect of in-plane loads are considered in the model. Using a multicriteria differential evolution algorithm, numerical optimization is performed for a variety of scenarios and aimed at increasing the flutter and linear buckling stability margins of tow-steered plates, considering the geometrical parameters defining the fiber trajectories on the layers as design variables. The results obtained for the different optimization scenarios are compared, having a composite plate with unidirectional fibers as the baseline and aimed at evaluating the benefits achieved by the optimum tow-steered plates. The results enable quantification of the stability improvements by exploring fiber steering, which has been shown to be beneficial, even in situations in which manufacturing constraints are accounted for. @en

The utter behavior of tow steered composite panels, in which the fiber placement follow curvilinear trajectory, is evaluated. A simple structural model based on Ritz method combined with supersonic aerodynamic piston theory is used to analyze the aeroelastic behavior. Classical lamination plate theory and symmetric stacking sequence are used and the fiber trajectories are defined by Lagrange interpolation functions. The utter stability boundaries for optimal conventional (constant stiffness laminates) layups and non- conventional (variable stiffness laminates) steered panels are numerically compared. The effect of in-plane loads is also accounted for in the aeroelastic analyses.@en