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N.P.M. Werter
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1
Application of the aeroelastic analysis and design framework developed at Delft University of Technology to a design of two aeroelastically tailored composite wings for a flying demonstrator is presented. The objective of the design process is to minimize structural mass of the wing while maintaining a target cruise shape. For this purpose, the jig shape of the wing is parameterized and becomes an integral part of the optimization, while the cruise shape is maintained by means of a constraint. Additionally strength, buckling, aeroelastic stability constraints, and a number of other design requirements have been introduced to obtain a feasible and flight-worthy design. Two wing types were designed: the reference wing and the tailored wing. The difference between the two wings is in the definition of the laminates comprising each wing. The reference wing was designed with symmetric-balanced laminates, while symmetric-unbalanced laminates were used for the tailored wing. The comparison is performed in terms of laminate stiffness and thickness distribution along the span, jig twist, and the aeroelastic response covering elastic deformations, aerodynamic load distribution, and wing root loads, showing a significant mass reduction for the tailored wing compared to the reference wing.
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Application of the aeroelastic analysis and design framework developed at Delft University of Technology to a design of two aeroelastically tailored composite wings for a flying demonstrator is presented. The objective of the design process is to minimize structural mass of the wing while maintaining a target cruise shape. For this purpose, the jig shape of the wing is parameterized and becomes an integral part of the optimization, while the cruise shape is maintained by means of a constraint. Additionally strength, buckling, aeroelastic stability constraints, and a number of other design requirements have been introduced to obtain a feasible and flight-worthy design. Two wing types were designed: the reference wing and the tailored wing. The difference between the two wings is in the definition of the laminates comprising each wing. The reference wing was designed with symmetric-balanced laminates, while symmetric-unbalanced laminates were used for the tailored wing. The comparison is performed in terms of laminate stiffness and thickness distribution along the span, jig twist, and the aeroelastic response covering elastic deformations, aerodynamic load distribution, and wing root loads, showing a significant mass reduction for the tailored wing compared to the reference wing.
Application of the aeroelastic framework developed at Delft University of Technology 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.
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Application of the aeroelastic framework developed at Delft University of Technology 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 present paper proposesa continuous-time state-space formulationofthe unsteady vortexlattice method, which is derived through a discretization of the governing advection equation for transport of vorticity in the wake. A continuous-time system isobtained byonly discretizing the advection equationinspace, while retaining the derivative with respect to time. The discretization in space is based on the discontinuous Galerkin method. The present method can be applied to any arbitrary nonuniform wake discretization and can be extended to higher-order panel methods or a nonflat wake shape. The method is extended to compressible flows by applying the Prandtl-Glauert transformation. The time-dependent terms in the small disturbance potential equation are neglected. Thus, incompressible flow solution procedures are applied with minimum modifications to unsteady compressible problems.The benefits are demonstratedby applying the modelto the gust analysisofageneral aircraft wing, varying the time step, and introducing a nonuniform wake discretization, resulting in a reduced model size for a given accuracy. The resulting continuous-time state-space model can be used for efficient loads analysis of general aircraft wings including the effects of compressibility and allows for easy integration with structural or flight dynamic models for efficient aero(servo)elastic analyses.
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The present paper proposesa continuous-time state-space formulationofthe unsteady vortexlattice method, which is derived through a discretization of the governing advection equation for transport of vorticity in the wake. A continuous-time system isobtained byonly discretizing the advection equationinspace, while retaining the derivative with respect to time. The discretization in space is based on the discontinuous Galerkin method. The present method can be applied to any arbitrary nonuniform wake discretization and can be extended to higher-order panel methods or a nonflat wake shape. The method is extended to compressible flows by applying the Prandtl-Glauert transformation. The time-dependent terms in the small disturbance potential equation are neglected. Thus, incompressible flow solution procedures are applied with minimum modifications to unsteady compressible problems.The benefits are demonstratedby applying the modelto the gust analysisofageneral aircraft wing, varying the time step, and introducing a nonuniform wake discretization, resulting in a reduced model size for a given accuracy. The resulting continuous-time state-space model can be used for efficient loads analysis of general aircraft wings including the effects of compressibility and allows for easy integration with structural or flight dynamic models for efficient aero(servo)elastic analyses.
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 flow visualisation by means of a tuft and smoke. The first assessment of the performance of the gust generator showed proper operation of the gust generator across the entire range of interest.
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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 flow visualisation by means of a tuft and smoke. The first assessment of the performance of the gust generator showed proper operation of the gust generator across the entire range of interest.
The design of composite structures based on _bre 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 param- eter into feasible blended stacking sequences. However, disparities between the lamination parameter and stacking sequence designs performance are generally observed due to dis- crepancies 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 in- creases the matching quality between lamination parameter and stacking sequence designs, consequently facilitating the retrieval of equivalent blended stacking sequences.
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The design of composite structures based on _bre 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 param- eter into feasible blended stacking sequences. However, disparities between the lamination parameter and stacking sequence designs performance are generally observed due to dis- crepancies 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 in- creases the matching quality between lamination parameter and stacking sequence designs, consequently facilitating the retrieval of equivalent blended stacking sequences.
In order to accommodate the growth in air traffic whilst reducing the impact on the environment, operational efficiency is becoming more and more important in the design of the aircraft of the future. A possible approach to increase the operational efficiency of aircraft wings is the use of aeroelastic tailoring, by taking advantage of the directional stiffness properties of composite materials to control the aeroelastic deformations of the wing in a beneficial way or morphing, by actively changing the wing shape in flight to optimise performance across a range of flight conditions.
In order to investigate the benefits of aeroelastic tailoring and morphing, this dissertation presents a dynamic aeroelastic analysis and optimisation framework suitable for the design of aeroelastically tailored and morphing wings. The framework is sufficiently efficient to explore the design space, as well as sufficiently comprehensive to account for all factors relevant in the design of aircraft wings.
In order to illustrate the advantages of the framework, it has been applied to two design studies: the optimisation of a morphing wing designed for a 25 kg UAV and the optimisation of the NASA Common Research Model, a contemporary transonic supercritical wing, resulting in wing designs that take advantage of the aeroelastic response of the wing, ensuring optimal performance at cruise flight conditions, while showing significant improvements at off-cruise conditions.
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In order to accommodate the growth in air traffic whilst reducing the impact on the environment, operational efficiency is becoming more and more important in the design of the aircraft of the future. A possible approach to increase the operational efficiency of aircraft wings is the use of aeroelastic tailoring, by taking advantage of the directional stiffness properties of composite materials to control the aeroelastic deformations of the wing in a beneficial way or morphing, by actively changing the wing shape in flight to optimise performance across a range of flight conditions.
In order to investigate the benefits of aeroelastic tailoring and morphing, this dissertation presents a dynamic aeroelastic analysis and optimisation framework suitable for the design of aeroelastically tailored and morphing wings. The framework is sufficiently efficient to explore the design space, as well as sufficiently comprehensive to account for all factors relevant in the design of aircraft wings.
In order to illustrate the advantages of the framework, it has been applied to two design studies: the optimisation of a morphing wing designed for a 25 kg UAV and the optimisation of the NASA Common Research Model, a contemporary transonic supercritical wing, resulting in wing designs that take advantage of the aeroelastic response of the wing, ensuring optimal performance at cruise flight conditions, while showing significant improvements at off-cruise conditions.
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 discretised 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.
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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 discretised 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.
Conference paper
(2016)
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M. Radestock, J. Riemenschneider, H.P. Monner, O. Huxdorf, Noud Werter, Roeland De Breuker
In a wind tunnel experiment a morphing wing with span extension and camber morphing was investigated. The considered aircraft is an unmanned aerial vehicle (UAV) with a span of 4 m. During the investigations a half wing model was analysed with pressure and structural measurement. The half wing model has three different morphing mechanisms. The focus is on the morphing leading edge concept, which will be shortly described with the main objectives. First of all the morphing leading edge will be described. Afterwards the results of the wind tunnel measurement for pressure distribution, balance and deformation measurement will be presented and discussed.
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In a wind tunnel experiment a morphing wing with span extension and camber morphing was investigated. The considered aircraft is an unmanned aerial vehicle (UAV) with a span of 4 m. During the investigations a half wing model was analysed with pressure and structural measurement. The half wing model has three different morphing mechanisms. The focus is on the morphing leading edge concept, which will be shortly described with the main objectives. First of all the morphing leading edge will be described. Afterwards the results of the wind tunnel measurement for pressure distribution, balance and deformation measurement will be presented and discussed.
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 differ- ent skin thicknesses and composite layups are designed in order to minimise root bending moment under manoeuvre loading using an aeroelastic anal- ysis framework that closely couples a geometrically nonlinear beam model to a vortex lattice aerodynamic model. The globally convergent method of moving asymptotes is used to derive an optimised layup for the tailored wings. In addition a quasi-isotropic wing is analysed to serve as a ref- erence. Both the tailored wings and the quasi-isotropic wing have been manufactured and tested structurally and in the wind tunnel. In the wind tunnel, aerodynamic forces and moments and wing deformation have been measured to provide experimental validation data.
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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 differ- ent skin thicknesses and composite layups are designed in order to minimise root bending moment under manoeuvre loading using an aeroelastic anal- ysis framework that closely couples a geometrically nonlinear beam model to a vortex lattice aerodynamic model. The globally convergent method of moving asymptotes is used to derive an optimised layup for the tailored wings. In addition a quasi-isotropic wing is analysed to serve as a ref- erence. Both the tailored wings and the quasi-isotropic wing have been manufactured and tested structurally and in the wind tunnel. In the wind tunnel, aerodynamic forces and moments and wing deformation have been measured to provide experimental validation data.
Driven by the need to improve the performance and energy-efficiency of aircraft, current research in the field of morphing wings is growing in significance. The most recently developed concepts typically adjust only one characteristic of the wing. Within this paper a new concept for morphing wings is developed and tested, enabling large changes of the chord length and camber simultaneously. By changing two characteristics of the wing instead of one, the range of flight missions can be extended more effectively. To achieve these large shape changes, highly adaptable leading and trailing sections are mounted onto a rigid wingbox. A thin polymer film is encompassing these three sections. By adjusting the length of this film, the outline of the wing can be changed significantly. A prototype has been designed and manufactured for wind tunnel tests. The leading and trailing sections are made of polyurethane foam, which can be compressed to 10 percent of its original volume. Different polyurethane foams are tested for optimal stiffness to withstand the aerodynamic loads acting on the wing, while being soft enough to accommodate the large deformations. The film encompassing the sections is retracted into the wingbox to achieve the desired shapes of the airfoil. On the one hand, this film needs to be very thin to fit into the wingbox, while on the other hand it needs to be stiff enough to withstand the pressure exerted by the foam. The manufactured prototype enables changes of the chord length by 30 percent and the camber by 10 percent of the chord length. These deformations were achieved without any significant kinking or buckling. Three different asymmetric airfoil shapes and two symmetric shapes with different chord lengths were tested in a series of wind tunnel tests. All five airfoil shapes deformed very little under wind loads. The lift and drag results were compared to generate values and matched very closely. The prototype fulfils all predefined requirements and performs very well over a wide range of airfoil shapes and wind speeds.
...
Driven by the need to improve the performance and energy-efficiency of aircraft, current research in the field of morphing wings is growing in significance. The most recently developed concepts typically adjust only one characteristic of the wing. Within this paper a new concept for morphing wings is developed and tested, enabling large changes of the chord length and camber simultaneously. By changing two characteristics of the wing instead of one, the range of flight missions can be extended more effectively. To achieve these large shape changes, highly adaptable leading and trailing sections are mounted onto a rigid wingbox. A thin polymer film is encompassing these three sections. By adjusting the length of this film, the outline of the wing can be changed significantly. A prototype has been designed and manufactured for wind tunnel tests. The leading and trailing sections are made of polyurethane foam, which can be compressed to 10 percent of its original volume. Different polyurethane foams are tested for optimal stiffness to withstand the aerodynamic loads acting on the wing, while being soft enough to accommodate the large deformations. The film encompassing the sections is retracted into the wingbox to achieve the desired shapes of the airfoil. On the one hand, this film needs to be very thin to fit into the wingbox, while on the other hand it needs to be stiff enough to withstand the pressure exerted by the foam. The manufactured prototype enables changes of the chord length by 30 percent and the camber by 10 percent of the chord length. These deformations were achieved without any significant kinking or buckling. Three different asymmetric airfoil shapes and two symmetric shapes with different chord lengths were tested in a series of wind tunnel tests. All five airfoil shapes deformed very little under wind loads. The lift and drag results were compared to generate values and matched very closely. The prototype fulfils all predefined requirements and performs very well over a wide range of airfoil shapes and wind speeds.