Current advances in the structural optimization of aircraft structures have led to the introduction of sandwich panels into the optimization process. This study attempts to extend the possibilities of sandwich optimization by proposing an analytical model which predicts the homogenized properties of a sandwich panel with a honeycomb core and CFRP skins. The model is based on a combination of Classical laminate theory and a 1-D beam model of the honeycomb core. The finite-element equivalent of tensile and shear tests is used to validate the proposed model on a broad range of core geometries with different combinations of core thickness, wall angle, cell elongation, and cell wall thickness. The results of 425 different geometries showed the overall precision of the proposed model, highlighted effects in the behavior of the core that drive the sandwich properties further from the predicted values, and suggested which parts of the model are suitable for optimization and where are their limits of applicability.

Actuator nonlinearities can significantly affect control systems, leading to performance degradation and even loss of stability. Physical constraints such as rate and acceleration limits are particularly detrimental in applications where rapid actuation is required, yet their combined effects remain largely unexplored. This paper investigates the nonlinear dynamic behaviour induced by rate and acceleration limits in closed-loop systems, focusing on their steady-state response to sinusoidal excitation. The saturation regimes associated with these nonlinearities are fully characterised, and their analytical boundaries are represented in a two-dimensional parameter space defined by normalised rate and acceleration limits. Sinusoidal describing functions are derived for each regime, providing a unified frequency-domain representation of the actuator dynamics. These formulations are employed to analyse the impact of actuator nonlinearities on closed-loop dynamics, including the onset of nonlinear behaviour, phase lag and gain reduction. Analytical conditions for the occurrence of jump resonance are derived, along with the lowest frequency where multiple steady-state solutions appear, leading to potential abrupt changes in system response. The applicability of the proposed framework is demonstrated through both an illustrative first-order system and a realistic high-order aeroservoelastic model for gust load alleviation, where the interaction between actuator nonlinearities and closed-loop dynamics is shown to produce multiple jump resonance scenarios and isolated nonlinear response branches. The results highlight the critical role of actuator rate and acceleration limits in high-bandwidth control applications and provide practical insights for frequency-domain stability assessment and preliminary feedback control system design.

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.

Accurate modeling of atmospheric turbulence is critical for the design and operation of next-generation large-scale wind turbines, particularly those exceeding 15 MW rated capacity and spanning well above the atmospheric surface layer (typically 10 − 20% of the atmospheric boundary layer (ABL)). In this study, Large Eddy Simulations (LES) were performed to investigate turbulence characteristics at high altitudes, up to 300 m above ground level — a region increasingly relevant for large turbine rotors. Turbulence coherence was analyzed and compared with field measurements to assess the fidelity of numerical predictions. Coherence estimates from LES were validated against lidar-based measurements obtained under stable, neutral, and unstable atmospheric conditions. Results show good agreement in the coherence decay rates and cross-spectral characteristics, with notable discrepancies only at very low frequencies (on the order of several 10 ⁻⁴ Hz) and large spatial separations (on the order of several 10 ² m). Consequently, a LES-tuned empirical lateral coherence model is proposed, featuring distinct coherence decay rates for each atmospheric stability regime (stable, neutral, and unstable ABL), offering improved representation of turbulence structures across a range of operating conditions. These findings provide a valuable reference for refining turbulence models for improving load estimation methodologies for next-generation wind turbines operating at hub heights above 200 m.

Large wind turbines face more intricate atmospheric conditions with turbulent coherent structures sized similarly to the rotor diameter, posing loading challenges. The present study assesses twelve distinct wind fields using the Large Eddy Simulations (LES) and International Electrotechnical Commission (IEC) Kaimal model scaled to their LES counterpart. The hub height wind speed in the different cases was set to 8.5 m/s (below-rated), 11.5 m/s (at-rated), and 14.5 m/s (above-rated). In a previous study, it was found that the unscaled IEC model-based wind field is conservative and scaled IEC model-based wind fields were found to yield different loads than upon use of LES-based wind fields in different atmospheric stability conditions. The present study aims to understand these differences. Utilizing Spectral Proper Orthogonal Decomposition (SPOD), the original wind fields were decomposed and reconstructed to study the influence of large and small coherent structures represented by their distinct frequencies. SPOD analysis was complemented by wind field spectral analysis considering atmospheric surface layer height, integral length scales, and co-coherence estimates. Integral length scales in the scaled IEC Kaimal model were found to be half of those in unstable atmosphere LES wind fields. The aero-elastic impact on the IEA 22 MW reference wind turbine with a 280 m rotor diameter was evaluated. The analysis reveals that large coherent structures, particularly low-frequency (≤0.06 Hz) ones, significantly impact wind turbine loads, contingent upon atmospheric stratification. Compared to the scaled IEC Kaimal model wind field, the maximum tower fore–aft bending moment and the maximum blade root flap-wise bending moment were found to be higher, for example, by 10% and 5% respectively in an unstable atmosphere during below-rated wind turbine operation. In the same scenario, standard deviation of the tower fore–aft bending moment was found to be higher by up to 50% while standard deviation of the blade root flap-wise bending moment was found to be lower by up to 25%. These findings underscore the critical importance of accurately modeling atmospheric turbulence and its coherent structures for more reliable design and operation of large wind turbines.

Shape sensing techniques allow for the time-efficient reconstruction of displacements based on measured strain data. There are technical applications, where the structure of interest is deformed in the geometrically non-linear domain. In aeronautics, this is the case for high-aspect-ratio wings, which are more frequently found in future designs. Only shape sensing methods that specifically take the non-linearity into account, can deliver appropriate displacement estimates for such application. A shape sensing method based on the linear modal approach can be utilised incrementally to capture the geometric non-linearity; it has therefore been denoted incremental modal method (IMM). This paper presents analytical relations for the uncertainty propagation for the various input quantities of the method, specifically strain mode shapes, displacement mode shapes, and measured strain. Deterministic shape sensing and uncertainty propagation are demonstrated using data obtained with a finite element model of a high-aspect-ratio wing experiencing geometric non-linear deflections in flapwise bending. Virtual strain and acceleration sensors are assumed for this setup, imitating the instrumentation conceivable for experimental work. The results obtained by analytical propagation are compared to Monte Carlo simulations for the purpose of validation. The derived propagation formulas make it possible to follow the evolution of the uncertainties over the number of increments. Given that the variability of the input quantities is known, the number of increments that minimise uncertainties can be determined for a model-free application of the shape sensing. Together with the deterministic estimates provided by an FE model, it is possible to determine the ideal number of increments for a specific shape sensing application in the geometrically non-linear domain.

This paper investigates the impact of introducing a switchable vortex generator (SVG), acting as a minitab, on the aerodynamic performance of a high-aspect-ratio wing’s outer section in transonic regime. A parametric study is conducted employing computational fluid dynamics two-dimensional simulations, focusing on the aerodynamic effects of changing the chordwise position and height of the vane of a SVG located on the airfoil upper surface in both nominal cruise conditions and for varying angles of attack. The analysis reveals that minitabs can strongly affect the aerodynamic forces produced by the wing section, showing great potential for load alleviation and control, but also emphasizing the need for a careful parameter selection to reduce undesirable effects such as the generation of shock waves. In cruise conditions, lift reduction increases with the vane height and has its maximum for chordwise positions at 60% of the chord length. However, SVGs located in the first half of the chord length yield more robust performance for varying angle of attack, without sharp lift variations or generated shock waves, and a delayed stall onset. High SVGs (greater than or equal to 3% chord length) can also lead to strong shock waves on the airfoil lower surface at small or negative angle of attack, while small SVGs (less than 1% chord length) can generate normal shock waves on the upper surface, with limited lift reduction in cruise conditions and at higher incidence.

There are technical applications where structures undergo deformation in the geometrically non-linear domain. This is the case for high-aspect-ratio wings, which may play a more important role in the future aircraft designs. Shape sensing methods can estimate the deflection of these structures during operation, if a direct measurement of the displacements is inconvenient or not possible. For the geometrically nonlinear range, the modal rotation method has been proposed as a candidate suitable for slender structures. The method superposes modal rotation increments of segments along the length of the structure, typically obtained from a finite element model. If the method is applied model-free, based on modal rotations identified from test data, the variability of the modal rotations leads to uncertainty in the displacement estimates. The present study illustrates how displacement output uncertainty can be expressed using linearised propagation formulae, relying on the prerequisite that the modal rotations exhibit a normally distributed and independent scatter around their mean. This uncertainty propagation is investigated in the shape sensing of a high-aspect-ratio wing model, and verification through Monte Carlo simulations demonstrates that the derived expressions accurately propagate the uncertainty from variable modal rotations. Consequently, these expressions can be applied to specific shape sensing tasks in experiments where this variability can be recorded.

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.

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.