Driven by a need to improve the efficiency of aircraft and reduce the fuel consumption, composite materials are applied extensively in the design of aircraft. A dynamic aeroelastic framework for the conceptual design of a generic composite wing structure is presented. The wing is discretized in several spanwise sections, where each section has a number of laminates throughout the cross-section, each having their own stiffness and thickness. The model uses a geometrically nonlinear beam model linearized around the nonlinear static aeroelastic equilibrium position coupled to a continuous-time state-space unsteady aerodynamic model to obtain the dynamic aeroelastic response, making the model suitable for dynamic aeroelastic analysis of generic aircraft wings under the assumption of small disturbances with respect to the static aeroelastic equilibrium position. Two optimisations are run for a generic aircraft wing under manoeuvre load conditions and aeroelastic, structural, and aerodynamic constraints: one, a quasi-isotropic wing to serve as a reference solution and two, a fully tailored wing clearly showing the benefit of aeroelastic tailoring and the use of the present framework for conceptual wing design.
This paper summarises the design of a gust generator and the comparison between high fidelity numerical results and experimental results. The gust generator has been designed for a low subsonic wind tunnel in order to perform gust response experiments on wings and assess load alleviation. Special attention has been given to the different design parameters that influence the shape of the gust velocity profile by means of CFD simulations. Design parameters include frequency of actuation, flow speed, maximum deflection, chord length and gust vane spacing. The numerical results are compared to experimental results obtained using a hot-wire anemometer and using flow visualisation by means of smoke. Discrepancies have been noticed between CFD and flow measurements but trends compare well and the system is fully functional.
to a design of the tailored composite wing for a flying demonstrator is presented in this paper. In the design process the structural mass of the wing is minimised including a cruise shape constraint. Introduction of the cruise shape constraint is explained as well as a number of other important design requirements which were imposed in order to obtain a feasible and flight worthy design. The effect of the cruise shape constraint is investigated by performing a comparison study. For this purpose two wing types were defined: the reference wing and the tailored wing. The difference between the two wings is in the laminate definition comprising each wing. The reference wing was designed with symmetric-balanced laminates, while symmetric-only laminates were used for the tailored wing. The comparison was performed in terms of laminate stiffness and thickness distribution along the span, jig twist, and the aeroelastic response. Elastic deformations, aerodynamic load distribution and wing root loads are compared within the scope of aeroelastic response.
The design of composite structures based on fibre angle optimisation rapidly becomes intractable as the number of design variables increases. Lamination parameters can be used instead as intermediate design variables in order to overcome this issue. An extra step is then required in order to convert the optimal design expressed in lamination parameter into feasible blended stacking sequences. However, disparities between the lamination parameter and stacking sequence designs performance are generally observed due to discrepancies between both design spaces. In this paper, the lamination parameter blending constraints recently proposed by the authors are applied to the aeroelastic optimisation of the common research wing model in order to diminish these discrepancies and achieve more realistic lamination parameter designs. A comparison between the optimised designs achieved with and without the proposed blending constraints is carried out to evaluate our approach. Results demonstrate that the application of blending constraints greatly increases the matching quality between lamination parameter and stacking sequence designs, consequently facilitating the retrieval of equivalent blended stacking sequences.
Werter et al. 1 introduced a novel approach to the conceptual design of morphing aircraft by means of a two-level design approach. This paper presents the extension of this model with an updated camber and span morphing model, a parasitic drag model, and a dynamic aeroelastic model. The first level consist of a morphing wing model specific to the concept that is investigated and is used as an input to the generic morphing aeroelastic optimisation framework in the second level. The second level returns the optimum morphing configuration and the energy required to overcome the external forces to reach this configuration. This information can be used to assess the feasibility of the design and, if required, the feasible design limits defined in the first level can be redefined. The framework has been applied to the optimisation of a UAV at three different flight conditions for minimum drag in a trimmed flight condition under aeroelastic and morphing constraints. The results show the trade-off between parasitic and induced drag depending on the flight speed and show that the framework presented can be used successfully for the design and optimisation of morphing wings.
The goal of the present paper is to provide experimental validation data for the aeroelastic analysis of composite aeroelastically tailored wings with a closed-cell cross-sectional structure. Several rectangular wings with different skin thicknesses and composite layups are designed in order to minimise root bending moment under manoeuvre loading using an aeroelastic anal-
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