Purpose: This study investigates how curvilinear fibre paths in variable stiffness composite laminates (VSCLs) influence large-amplitude, non-linear aeroelastic oscillations—particularly limit-cycle oscillations (LCOs) and chaotic responses—of circular cylindrical shells under supersonic flow. The aim is to assess whether curvilinear fibre reinforcements offer stability and performance advantages over traditional straight-fibre laminates in post-flutter regimes. Methods: A new geometrically non-linear model for circular cylindrical shells reinforced by curvilinear fibres is developed. It uses Kirchhoff’s hypothesis and von Kármán strain–displacement relations, with the curvilinear fibre paths influencing the stiffness related terms. A single-element computational model incorporating polynomial and trigonometric basis functions enables efficient dynamic analysis. Linear stability is assessed via eigensolution routines, and post-flutter non-linear responses are obtained by time domain integration, where advantage is taken of the naturally reduced-order model. Results: The study finds that while the circumferential component of membrane inertia significantly affects certain vibration and flutter modes, longitudinal inertia can be neglected. Curvilinear fibre configurations delay flutter onset and modify the post-flutter response. Both LCOs and chaotic oscillations are observed, with curvilinear fibres shown to reduce oscillation amplitudes and lower frequency content in LCOs. Longitudinally travelling waves are identified and it is found that non-linear modal interaction is connected with chaotic behaviour. Conclusion: Curvilinear fibre orientations enhance the aeroelastic performance of cylindrical shells by extending the stable operating range before flutter occurs and by reducing the severity of post-flutter oscillations. These findings suggest that the use of curvilinear reinforcement fibres enables improvements in the non-linear dynamic behaviour of aerospace circular cylindrical thin structures.

Modeling open-hole failure of composites is a complex task, consisting of a highly nonlinear response with interacting failure modes. Numerical modeling of this phenomenon has traditionally been based on the finite element method, but requires to tradeoff between high fidelity and computational cost. To mitigate this shortcoming, recent work has leveraged machine learning to predict the strength of open-hole composite specimens. Here, we also propose using data-based models to tackle open-hole composite failure from a classification point of view. More specifically, we show how to train surrogate models to learn the ultimate failure envelope of an open-hole composite plate under in-plane loading. To achieve this, we solve the classification problem via support vector machine (SVM) and test different classifiers by changing the SVM kernel function. The flexibility of kernel-based SVM also allows us to integrate the recently developed quantum kernels in our algorithm and compare them with the standard radial basis function kernel. Finally, thanks to kernel-target alignment optimization, we tune the free parameters of all kernels to best separate safe and failure-inducing loading states. The results show classification accuracies higher than 90% for RBF, especially after alignment, followed closely by the quantum kernel classifiers.

This article presents a novel camber-twist morphing flap concept with two chordwise degrees-of-freedom. The flap is capable of reflexed airfoil morphing, thereby decoupling lift from the aerodynamic moment with respect to the aerodynamic centre. The theoretical potential of such a flap is calculated via XFOIL for arbitrary trailing edge shapes, revealing ellipse-like clusters in the lift-moment plane for each value of angle of attack. A conceptual design is proposed, capable of the above functionality. Key features include two spanwise slits along the pressure side skin joined by a flexible structure, with a spar placed between them and two pairs of linear electric motors. The design is validated numerically using a nonlinear aeroelastic analysis toolchain, iterating between the finite element model of the flap and XFOIL. The attainable range of lift-moment combinations is calculated, forming an ellipse-like cluster determined by actuator stroke and force limits. The morphing flap achieves a lift-to-drag ratio of over 104.3 over a range of angles of attack. A high degree of twist morphing range is demonstrated by fixing one pair of actuators and varying the strokes on the other. The range of attainable shapes on the free end is coupled to the fixed end strokes.

A global optimization framework with global–local verification is developed for composite wings by coupling a NASTRAN-based global finite element model with Carrera Unified Formulation (CUF)-based high-order local analyses. The proposed methodology enables the minimization of structural weight under stress and buckling constraints, while identifying critical zones of interest for multi-fidelity and hierarchical strategy. The global model based on first-order shear deformation theory identifies the critical regions where the CUF Equivalent Single Layer and Layer-Wise refinements should be introduced to provide detailed local buckling and stress analyses. This integrated global-local framework enables adaptive local fidelity global-local analysis, improving margins of accuracy while mapping the accessible design space and maintaining computational efficiency.

Correction notice The CL in the title of Fig. 7(b) was corrected from 0.4 in the original version to CL=1.0. (a) Climb local lift spanwise distribution at CL=1 0 0.2 0.4 0.6 0.8 1 0 5 10 10-3 0 0.2 0.4 0.6 0.8 1 0 5 10 10-3 (b) Solid line (suction side)-dashed line (pressure side) Fig. 7 Local lift coefficient distribution with a selected friction coefficient of one section.

High aspect ratio strut-braced wing aircraft can significantly reduce the induced drag while limiting the weight penalty of increasing the wingspan. As part of the Hybrid Electric Regional Wing Integration Novel Green Technologies (HERWINGT) project, a multifunctional morphing strut is being investigated. In this study, an optimization framework is proposed to define the thickness distribution of the morphing trailing edge of the strut to achieve the desired operational shapes while considering laminate manufacturing guidelines and material allowables. The optimizer finds designs capable of achieving the objective shapes and provides load and mass estimations that can be used to make design decisions.

This paper aims to develop a reduced-order modelling methodology for nonlinear, unsteady, aerodynamic loads for active control transonic aeroelastic instabilities. To this end, a NACA0012 airfoil equipped with a flap is chosen as the test configuration. The aim here is to understand the interaction between the transonic shock dynamics and flap actuation at various amplitudes and frequencies. The high-fidelity simulations are carried out for two angles of attack, i.e. a = 0.0°, 4.0°. It is found that transonic buffet characteristics significantly change with airfoil geometry. Additionally, the flap is seen to be ineffective in the separated flow regions, thereby making the Ci-fi slopes highly nonlinear. However, increasing the frequencies of flap oscillations, increases flap effectiveness, increases control over buffet motion and moves towards linear lift responses. Furthermore, we also evaluate the performance of several Bayesian Filters that are crucial in the state-estimation process of the active control of nonlinear systems. It is observed that nonlinear filters such as Unscented Karman Filter perform better than the traditional linear Kalman Filter as system response to flap actuation becomes nonlinear in the presence of separated boundary layer.

Cantilevers find a wide range of applications in the design of scientific equipment and large-scale engineering structures such as aircraft wings. Analysis techniques based on linearization approximations are unable to capture the large amplitude oscillation behaviour of such structures and thus, necessitates development of dedicated nonlinear methods. In this work, the recent developments in the Koiter-Newton model reduction method are utilized to obtain nonlinear reduced order models (ROMs) from full finite element structural models in order to simulate large amplitude dynamics of cantilevers. The method describes a nonlinear system of governing equations comprising quadratic and cubic terms which are obtained as higher order derivatives of the in-plane strain energy. To ensure that the large rotations in cantilevers and the resultant foreshortening effect is also accounted for, a ROM updating algorithm is adopted where the ROM parameters are varied with the structural deflections. Linear eigenmodes of the structure are utilized to form the reduction subspace. To validate the methodology, the ROM solution is compared against experimental results and a convergence study is conducted to identify the number of modes needed to replicate the nonlinear response. Finally, a composite wingbox structure is considered for which time domain simulations are conducted and frequency response curves, obtained through a frequency sweep, are presented.

In the development of electric aircraft, due to the use of Distributed Electric Propulsion (DEP), not only the classic wing flutter but also the propeller whirl flutter needs to be considered for wing structural design. To this end, this paper proposes an aeroelastic optimization method within the framework of an in-house tool named PROTEUS, which enables the preliminary design of DEP wing laminates including propeller whirl flutter effect. In this method, a new aeroelastic model is developed for the coupled propeller-wing system, based on a classic whirl flutter analysis model and the wing aeroelastic model implemented in PROTEUS. Further, the required sensitivities of aeroelastic stability constraints are derived and implemented by making use of these implemented in PROTEUS for conventional wing design. The objective of the optimization is to minimize wing mass by aeroelastically tailoring the lamination parameters and thickness of wing laminates, subject to given aerostructural design constraints. The features and usefulness of the proposed optimization approach are demonstrated through two numerical case studies (with and without whirl flutter constraints) focused on sizing the wing structure of a reference DEP aircraft. The necessary inputs regarding propeller mounting stiffness and damping for the case studies are determined through parametric studies of isolated propellers. The results indicate that including whirl flutter effect in wing sizing slightly increases wing mass, and introducing a flexible-mount-propeller leads to the decrease in wing flutter speed. Additionally, a parametric study of investigating propeller mounting stiffness is conducted, which confirms that the propeller mounting properties have a large influence on aeroelastic instability of the coupled propeller-wing system.