This paper presents a novel damage mechanics based failure model enabling the prediction of low cycle fatigue life and residual strength of isotropic structures under multiaxial loading. The approach herein proposed does not discretize every load cycle but instead takes an envelope loading whereby the numerical load remains constant at a maximum load level and the number of cycles is obtained from a given elapsed time defined within a pseudo-time framework. The proposed formulation is based on the smeared cracking approach accounting for damage propagation due to static and fatigue loadings, where the static component is based on the Von-Mises yield criterion and Prandtl-Reuss stress flow rule; whereas the crack propagation in cyclic loading component is based on the Paris-law. Furthermore, the formulation combines damage mechanics and fracture mechanics within a unified approach enabling the control of the energy dissipated in each loading cycle.

The use of morphing structures aims to increase aerodynamic efficiency, decreasing fuel consumption and aircraft overall weight. Within this context, piezoelectric materials are of great interest in the design of smart structures because the piezoelectric effect is reversible, allowing them to work both as sensors and as actuators. The present work proposes a geometrically nonlinear finite element formulation for composite beams with embedded piezoelectric layers for application in morphing aerostructures. The formulation uses the complete Green strain tensor to account for geometric nonlinearities, and linear piezoelectricity to model the electromechanical behavior. A set of nonlinear equilibrium equations results from the application of variational principles, and is finally solved by means of an iterative-incremental arc length method. The sensitivity of the element to stacking sequence and number of actuators are investigated.

This work investigates the effects of temperature in the shape memory alloy hybrid composites (SMAHC) cylindrical stiffened panels’ aeroelastic stability. The SMAHC is modelled using a micromechanical formulation embedding carbon fiber, SMA wire and resin to the same lamina and taking into account the martensite/austenite phases of transformation in the material response. Virtual work principle formulation is implemented with classical laminate plate theory (CLPT) panel formulation and one-dimensional Euler-Bernoulli beam theory formulation for the stiffener. Numerical results are obtained by using an energy based semi-analytical method applying hierarchical polynomials to approximate the membrane and out of plane displacement fields. Different geometric configurations, laminate stacking sequences, boundary conditions and radii of curvature are investigated. The study shows that the variation of temperature induce stiffening due to changes in the martensite/austenite fractions of the SMA, increasing the critical flutter dynamic pressure. Therefore, it can be achieved certain control in the flutter critical boundary by increasing the temperature of the shape memory alloy (SMA) wire. The effects due to the SMA wire stiffening with the temperature are more pronounced for cross-ply stiffened cylindrical panels with unitary aspect ratio and for angle-ply panels with aspect ratio higher than one.

The substitution of conventional mechanical fasteners by adhesive joints has been advocated by the aircraft and aerospace industries due to the weight saving potential. Flaws such as debonding of the adhesive layer between the skin and the stiffener may greatly affect the structural behavior of composite panels. Within this context, this work presents a semi-analytical approach for the numerical investigation on the effects of skin-stiffener bonding flaw size on the vibration and linear buckling behavior of T-stiffened composite panels. Skin and stiffener have been modeled using an assembly of curved and flat panel components, with each domain approximated using a set of hierarchical polynomial functions. A penalty-based approach has been used to assemble the various domains and to model the debonded region between the stiffener flange base and the plate. This approach ensures full compatibility in terms of displacements and rotations between the stiffener's base top face and the panel bottom face allowing to model different skin/stiffener debonding lengths. The results obtained using the proposed semi-analytical models have been compared and verified against numerical predictions based on finite element analyses.

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.

Nondestructive experimental methods to calculate the buckling load of imperfection sensitive thin-walled structures are one of the most important techniques for the validation of new structures and numerical models of large scale aerospace structures. Vibration correlation technique (VCT) allows determining equivalent boundary conditions and buckling load for several types of structures without reaching the instability point. VCT is already widely used for beam structures, but the technique is still under development for thin-walled plates and shells. This paper intends to explain the capabilities and current limitations of this technique applied to two types of structures under buckling conditions: flat plates and cylindrical shells prone to buckling. Experimental results for a flat plate and a cylindrical shell are presented together with reliable finite element models for both cases. Preliminary results showed that the VCT can be used to determine the realistic boundary conditions of a given test setup, providing valuable data for the estimation of the buckling load by finite element models. Also numerical results herein presented show that VCT can be used as a nondestructive tool to estimate the buckling load of unstiffened cylindrical shells. Experimental tests are currently under development to further validate the approach proposed herein.