This paper investigates the influence of the spanwise-distributed trailing-edge camber morphing on the dynamic stall characteristics of a finite-span wing at Re=2×10^(5).The mathematical model of the spanwise-dis...This paper investigates the influence of the spanwise-distributed trailing-edge camber morphing on the dynamic stall characteristics of a finite-span wing at Re=2×10^(5).The mathematical model of the spanwise-distributed trailing-edge camber morphing is established based on Chebyshev polynomials,and the deformed wing surface is modeled by a spline surface according to the rib's morphing in the chordwise direction.The Computational Fluid Dynamics(CFD)method is adopted to obtain flow-field results and aerodynamic forces.The SST-γmodel is introduced and the overset mesh technique is adopted.The numerical results show that the spanwisedistributed trailing-edge morphing obviously changes the aerodynamic and energy transfer characteristics of the dynamic stall.Especially when the phase difference between the trailing-edge motion and the wing pitch is-π/2,the interaction between the three-dimensional(3-D)Leading-Edge Vortex(LEV)and Trailing-Edge Vortex(TEV)is strengthened,and the work done by the aerodynamic force turns negative.This indicates that the trailing-edge deformation has the potential to suppress the oscillation amplitude of stall flutter.We also found that as the trailing-edge camber morphing varies more complexly along the spanwise direction,the suppression effect decreases accordingly.展开更多
Monitoring the shape and deformation of morphing wings is vital for ensuring multi-mission flight and safety operation.During the morphing process,the complex deformation of the flexible skin wing usually involves lar...Monitoring the shape and deformation of morphing wings is vital for ensuring multi-mission flight and safety operation.During the morphing process,the complex deformation of the flexible skin wing usually involves large amounts of movement,shearing,bending,and distortion.This paper proposes an improved stereo-digital image correlation measurement system designed to characterize full-field complex large deformation of flexible skin shear variable-sweep wings(SVSWs).By minimizing reference image updating frequency using the proposed conditional incremental strategy,effectively addressing the computational failures caused by image decorrelation due to complex large deformations.To improve tracking efficiency and accuracy of uncoded targets in complex backgrounds,an automatic subpixel detection method for circular diagonal targets is presented.A series of experiments are performed on a 1200 mm span flexible skin SVSW to verify the proposed methods.The results show that the length and angle measurement accuracies are better than 0.11 mm and 0.05°,respectively.Based on the measured morphing geometry parameters,displacement and strain fields,the structural integrity and morphing performance of the wing under different loads are discussed.During the shear variable-sweep process,the wingtip load dominates the deflection distribution,while its effect on the strain distribution is relatively minor.The proposed method and system can provide reliable data support for the structural optimization design and safety evaluation of such morphing wings.展开更多
Hypersonic morphing vehicle(HMV)can reconfigure aerodynamic geometries in real time,adapting to diverse needs like multi-mission profiles and wide-speed-range flight,spanwise morphing and sweep angle variation are rep...Hypersonic morphing vehicle(HMV)can reconfigure aerodynamic geometries in real time,adapting to diverse needs like multi-mission profiles and wide-speed-range flight,spanwise morphing and sweep angle variation are representative large-scale wing reconfiguration modes.To meet the HMV's need for an increased lift and a lift to drag ratio during hypersonic maneuverability and cruise or reentry equilibrium glide,this paper proposes an innovative single-DOF coupled morphing-wing system.We then systematically analyze its open-loop kinematics and closed-loop connectivity constraints,and the proposed system integrates three functional modules:the preset locking/release mechanism,the coupled morphing-wing mechanism,and the integrated wing locking with active stiffness control mechanism.Experimental validation confirms stable,continuous morphing under simulated aerodynamic loads.The experimental results indicate:(i)SMA actuators exhibit response times ranging from 18 s to 160 s,providing sufficient force output for wing unlocking;(ii)The integrated wing locking with active stiffness control mechanism effectively secures wing positions while eliminating airframe clearance via SMA actuation,improving the first-order natural frequency by more than 17%;(iii)The distributed aerodynamic loading system enables precise multi-stage follow-up loading during morphing,with the coupled morphing wing maintaining stable,continuous operation under 0-3500 N normal loads and 110-140 N axial force.The proposed single-DOF coupled morphing mechanism not only simplifies and improves structural efficiency but also demonstrates superior performance in locking control,stiffness enhancement,and aerodynamic responsiveness.This establishes a foundational framework for the design of future intelligent morphing configurations and the implementation of flight control systems.展开更多
An active control technique utilizing piezoelectric actuators to alleviate gust-response loads of a large-aspect-ratio flexible wing is investigated. Piezoelectric materials have been extensively used for active vibra...An active control technique utilizing piezoelectric actuators to alleviate gust-response loads of a large-aspect-ratio flexible wing is investigated. Piezoelectric materials have been extensively used for active vibration control of engineering structures. In this paper, piezoelectric materials further attempt to suppress the vibration of the aeroelastic wing caused by gust. The motion equation of the flexible wing with piezoelectric patches is obtained by Hamilton＇s principle with the modal approach, and then numerical gust responses are analyzed, based on which a gust load alleviation（GLA） control system is proposed. The gust load alleviation system employs classic propor tional-integral-derivative（PID） controllers which treat piezoelectric patches as control actuators and acceleration as the feedback signal. By a numerical method, the control mechanism that piezoelectric actuators can be used to alleviate gust-response loads is also analyzed qualitatively. Furthermore, through low-speed wind tunnel tests, the effectiveness of the gust load alleviation active control technology is validated. The test results agree well with the numerical results. Test results show that at a certain frequency range, the control scheme can effectively alleviate the z and x wingtip accelerations and the root bending moment of the wing to a certain extent. The control system gives satisfying gust load alleviation efficacy with the reduction rate being generally over 20%.展开更多
A computational and test method for calibrating the flight loads carried by aircraft wings is proposed.The wing load is measured in real-time based on the resistance and fiber Bragg grating strain gauges.The linear st...A computational and test method for calibrating the flight loads carried by aircraft wings is proposed.The wing load is measured in real-time based on the resistance and fiber Bragg grating strain gauges.The linear stepwise regression method is used to construct the load equations.The mean impact value algorithm is employed to select suitable bridges.In the ground calibration experiment,the wing load calculation equations in both forward and reverse installation states are calibrated.The correctness of the load equations was verified through equation error and inspection error analysis.Finally,the actual flight load of the wing was obtained through flight tests.展开更多
Aeroelastic control is a critical technique for high-aspect-ratio flexible wings.A novel aeroelastic control method is introduced,utilizing the internal Moving Mass Control(MMC)technique,which demonstrates the potenti...Aeroelastic control is a critical technique for high-aspect-ratio flexible wings.A novel aeroelastic control method is introduced,utilizing the internal Moving Mass Control(MMC)technique,which demonstrates the potential to fulfill hybrid control demands without incurring a drag penalty.Dynamic equations for a flexible wing equipped with a spanwise moving mass under unsteady aerodynamic loading are derived using mass position as the input variable.Controloriented analyses indicate that intrinsic structural frequencies,flutter characteristics,and gust response can be actively modified by varying the spanwise and chordwise positions of the mass element.Among these,the chordwise position exerts a more significant impact on the structural modes and flutter speed of the wing.A hybrid aeroelastic control system,incorporating motion planning and control law,is proposed to evaluate real-time performance in Active Flutter Suppression(AFS)and Gust Load Alleviation(GLA).Control outcomes suggest that,with a mass ratio of 1/16 and a half-chord installation area for the guide rail,flutter speed increases by about 10%.Additionally,excitation amplitudes across different gust frequencies are substantially mitigated,achieving a maximum reduction of vibration amplitude by about 73%.These findings offer a comprehensive understanding of the MMC technique and its application to flexible aircraft.展开更多
High-aspect-ratio aircraft are widely used in military and civilian fields,such as reconnaissance,surveillance,and attacks,due to their high lift-to-drag ratio,strong payload capability,significant endurance effect,an...High-aspect-ratio aircraft are widely used in military and civilian fields,such as reconnaissance,surveillance,and attacks,due to their high lift-to-drag ratio,strong payload capability,significant endurance effect,and good stealth performance.However,compared to conventional aircraft,high-aspect-ratio aircraft are more susceptible to gust disturbances during flight.In response to this phenomenon,a full-scale dynamic model of a high-aspect-ratio unmanned aerial vehicle was developed.Considering the coupling among control surfaces,structural forces,and aerodynamic forces,along with sensor,actuator,and delay effects,an H_(∞)control law was designed using the principle of singular value energy flow reduction and weighted function,with a PID(Proportional-Integral-Derivative)control law for comparison.The two controllers were then subjected to pulse-response and jury stability tests.Finally,wind tunnel tests were conducted to investigate the gust alleviation principle,in which gust disturbances were generated using gust generators and control surface self-excitation.The results present that the average wing root bending moment and wing tip overload under the PID control law decrease by approximately 30%,while under the H_(∞)control law,both the average wing root bending moment and wing tip overload reduction rate exceed 50%,with peaks reaching 60%.This validates the feasibility and efficiency of the designed H_(∞)controller.展开更多
A theoretical nonlinear aeroelastic response analysis for a flexible high-aspect ratio wing excited by harmonic gust load is presented along with a companion wind tunnel test. A multidisci- plinary coupled numerical c...A theoretical nonlinear aeroelastic response analysis for a flexible high-aspect ratio wing excited by harmonic gust load is presented along with a companion wind tunnel test. A multidisci- plinary coupled numerical calculation is developed to simulate the flexible model wing undergoing gust load in the time domain via discrete nonlinear finite element structural dynamic analysis and nonplanar unsteady vortex lattice aerodynamic computation. A dynamic perturbation analysis about a nonlinear static equilibrium is also used to determine the small perturbation flutter bound- ary. A novel noncontact 3-D camera measurement analysis system is firstly used in the wind tunnel test to obtain the spatial large deformation and responses. The responses of the flexible wing under different static equilibrium states and frequency gust loads are discussed. The fair to good quanti- tative agreements between the theoretical and experimental results demonstrate that the presented analysis method is an acceptable way to predict the geometrically nonlinear gust response for flex- ible wings.展开更多
The aim of this study is to identify the influence of the dip angle of a pre-existing macrocrack on the lifetime and ultimate deformation of rock-like material. Prediction of lifetime has been studied for three groups...The aim of this study is to identify the influence of the dip angle of a pre-existing macrocrack on the lifetime and ultimate deformation of rock-like material. Prediction of lifetime has been studied for three groups of specimens under axial static compressive load levels. The specimens were investigated from 65% to 85% of UCS(uniaxial compressive strength) at an interval of 10% of UCS for the groups of specimens with a single modelled open flaw with a dip angle to the loading direction of 30°(first group), at an interval of 5% of UCS increment for the groups of specimens with single(second group), and double sequential open flaws with a dip angle to the loading direction of 60°(third group). This study shows that crack propagation in specimens with a single flaw follows the same sequences. At first, wing cracks appear, and then shear crack develops from the existing wing cracks. Shear cracking is responsible for specimen failure in all three groups. A slip is expected in specimens from the third group which connects two individual modelled open flaws. The moment of the slip is noticed as a characteristic rise in the axial deformation at a constant load level. It is also observed that axial deformation versus time follows the same pattern, irrespective of local geometry. Specimens from the first group exhibit higher axial deformation under different load levels in comparison with the specimens from the second and third groups.展开更多
The tip vortices and aerodynamics of a NACA0012 wing in the vicinity of the ground were studied in a wind tunnel.The wing tip vortex structures and lift/drag forces were measured by a seven-hole probe and a force bala...The tip vortices and aerodynamics of a NACA0012 wing in the vicinity of the ground were studied in a wind tunnel.The wing tip vortex structures and lift/drag forces were measured by a seven-hole probe and a force balance,respectively.The evolution of the flow structures and aerodynamics with a ground height were analyzed.The vorticity of tip vortices was found to reduce with the decreasing of the ground height,and the position of vortex-core moved gradually to the outboard of the wing tip.Therefore,the down-wash flow induced by the tip vortices was weakened. However,vortex breakdown occurred as the wing lowered to the ground.From the experimental results of aerodynamics,the maximum lift-to-drag ratio was observed when the angle of attack was 2.5°and the ground clearance was 0.2.展开更多
Fixed-wing aircraft cannot maintain optimal aerodynamic performance at different flight speeds. As a type of morphing aircraft, the shear variable-sweep wing(SVSW) can dramatically improve its aerodynamic performance ...Fixed-wing aircraft cannot maintain optimal aerodynamic performance at different flight speeds. As a type of morphing aircraft, the shear variable-sweep wing(SVSW) can dramatically improve its aerodynamic performance by altering its shape to adapt to various flight conditions.In order to achieve smooth continuous shear deformation, SVSW's skin adopts a flexible composite skin design instead of traditional aluminum alloy materials. However, this also brings about the non-linear difficulty in stiffness modeling and calculation. In this research, a new SVSW design and efficient stiffness modeling method are proposed. Based on shear deformation theory, the flexible composite skin is equivalently modeled as diagonally arranged nonlinear springs, simulating the elastic force interaction between the skin and the mechanism. By shear loading tests of flexible composite skin, the accuracy of this flexible composite skin modeling method is verified. The SVSW stiffness model was established, and its accuracy was verified through static loading tests. The effects of root connection, sweep angles, and flexible composite skin on the SVSW stiffness are analyzed. Finally, considering three typical flight conditions of SVSW: low-speed flow(Ma = 0.3,Re = 5.82 × 10^(6)), transonic flow(Ma = 0.9, Re = 3.44 × 10^(6)), and supersonic flow(Ma = 3,Re = 7.51 × 10^(6)), the stiffness characteristics of SVSW under flight conditions were evaluated.The calculated results guide the application of SVSW.展开更多
基金co-supported by the National Natural Science Foundation of China(No.12472332)。
文摘This paper investigates the influence of the spanwise-distributed trailing-edge camber morphing on the dynamic stall characteristics of a finite-span wing at Re=2×10^(5).The mathematical model of the spanwise-distributed trailing-edge camber morphing is established based on Chebyshev polynomials,and the deformed wing surface is modeled by a spline surface according to the rib's morphing in the chordwise direction.The Computational Fluid Dynamics(CFD)method is adopted to obtain flow-field results and aerodynamic forces.The SST-γmodel is introduced and the overset mesh technique is adopted.The numerical results show that the spanwisedistributed trailing-edge morphing obviously changes the aerodynamic and energy transfer characteristics of the dynamic stall.Especially when the phase difference between the trailing-edge motion and the wing pitch is-π/2,the interaction between the three-dimensional(3-D)Leading-Edge Vortex(LEV)and Trailing-Edge Vortex(TEV)is strengthened,and the work done by the aerodynamic force turns negative.This indicates that the trailing-edge deformation has the potential to suppress the oscillation amplitude of stall flutter.We also found that as the trailing-edge camber morphing varies more complexly along the spanwise direction,the suppression effect decreases accordingly.
基金supported by the National Natural Science Foundation of China(Grant Nos.12202282 and 12102267).
文摘Monitoring the shape and deformation of morphing wings is vital for ensuring multi-mission flight and safety operation.During the morphing process,the complex deformation of the flexible skin wing usually involves large amounts of movement,shearing,bending,and distortion.This paper proposes an improved stereo-digital image correlation measurement system designed to characterize full-field complex large deformation of flexible skin shear variable-sweep wings(SVSWs).By minimizing reference image updating frequency using the proposed conditional incremental strategy,effectively addressing the computational failures caused by image decorrelation due to complex large deformations.To improve tracking efficiency and accuracy of uncoded targets in complex backgrounds,an automatic subpixel detection method for circular diagonal targets is presented.A series of experiments are performed on a 1200 mm span flexible skin SVSW to verify the proposed methods.The results show that the length and angle measurement accuracies are better than 0.11 mm and 0.05°,respectively.Based on the measured morphing geometry parameters,displacement and strain fields,the structural integrity and morphing performance of the wing under different loads are discussed.During the shear variable-sweep process,the wingtip load dominates the deflection distribution,while its effect on the strain distribution is relatively minor.The proposed method and system can provide reliable data support for the structural optimization design and safety evaluation of such morphing wings.
基金supported by the National Natural Science Foundation of China(Grant No.52405257)the China Postdoctoral Science Foundation(Grant No.2024M764201).
文摘Hypersonic morphing vehicle(HMV)can reconfigure aerodynamic geometries in real time,adapting to diverse needs like multi-mission profiles and wide-speed-range flight,spanwise morphing and sweep angle variation are representative large-scale wing reconfiguration modes.To meet the HMV's need for an increased lift and a lift to drag ratio during hypersonic maneuverability and cruise or reentry equilibrium glide,this paper proposes an innovative single-DOF coupled morphing-wing system.We then systematically analyze its open-loop kinematics and closed-loop connectivity constraints,and the proposed system integrates three functional modules:the preset locking/release mechanism,the coupled morphing-wing mechanism,and the integrated wing locking with active stiffness control mechanism.Experimental validation confirms stable,continuous morphing under simulated aerodynamic loads.The experimental results indicate:(i)SMA actuators exhibit response times ranging from 18 s to 160 s,providing sufficient force output for wing unlocking;(ii)The integrated wing locking with active stiffness control mechanism effectively secures wing positions while eliminating airframe clearance via SMA actuation,improving the first-order natural frequency by more than 17%;(iii)The distributed aerodynamic loading system enables precise multi-stage follow-up loading during morphing,with the coupled morphing wing maintaining stable,continuous operation under 0-3500 N normal loads and 110-140 N axial force.The proposed single-DOF coupled morphing mechanism not only simplifies and improves structural efficiency but also demonstrates superior performance in locking control,stiffness enhancement,and aerodynamic responsiveness.This establishes a foundational framework for the design of future intelligent morphing configurations and the implementation of flight control systems.
基金supported by the National Key Research and Development Program (2016YFB 0200703)
文摘An active control technique utilizing piezoelectric actuators to alleviate gust-response loads of a large-aspect-ratio flexible wing is investigated. Piezoelectric materials have been extensively used for active vibration control of engineering structures. In this paper, piezoelectric materials further attempt to suppress the vibration of the aeroelastic wing caused by gust. The motion equation of the flexible wing with piezoelectric patches is obtained by Hamilton＇s principle with the modal approach, and then numerical gust responses are analyzed, based on which a gust load alleviation（GLA） control system is proposed. The gust load alleviation system employs classic propor tional-integral-derivative（PID） controllers which treat piezoelectric patches as control actuators and acceleration as the feedback signal. By a numerical method, the control mechanism that piezoelectric actuators can be used to alleviate gust-response loads is also analyzed qualitatively. Furthermore, through low-speed wind tunnel tests, the effectiveness of the gust load alleviation active control technology is validated. The test results agree well with the numerical results. Test results show that at a certain frequency range, the control scheme can effectively alleviate the z and x wingtip accelerations and the root bending moment of the wing to a certain extent. The control system gives satisfying gust load alleviation efficacy with the reduction rate being generally over 20%.
基金supported by the National Natural Science Foundation of China(Grant No.11602237)the Middleaged and Young Teachers’Basic Ability Promotion Project of Guangxi(Grant No.2022KY1070)。
文摘A computational and test method for calibrating the flight loads carried by aircraft wings is proposed.The wing load is measured in real-time based on the resistance and fiber Bragg grating strain gauges.The linear stepwise regression method is used to construct the load equations.The mean impact value algorithm is employed to select suitable bridges.In the ground calibration experiment,the wing load calculation equations in both forward and reverse installation states are calibrated.The correctness of the load equations was verified through equation error and inspection error analysis.Finally,the actual flight load of the wing was obtained through flight tests.
基金supported by the National Natural Science Foundation of China(No.12102096)the Guangdong Basic and Applied Basic Research Foundation,China(No.2022A1515011885)the Research Fund of National Key Laboratory of Aerospace Physics in Fluids,China(No.2024-APF-KFQMJJ-08)。
文摘Aeroelastic control is a critical technique for high-aspect-ratio flexible wings.A novel aeroelastic control method is introduced,utilizing the internal Moving Mass Control(MMC)technique,which demonstrates the potential to fulfill hybrid control demands without incurring a drag penalty.Dynamic equations for a flexible wing equipped with a spanwise moving mass under unsteady aerodynamic loading are derived using mass position as the input variable.Controloriented analyses indicate that intrinsic structural frequencies,flutter characteristics,and gust response can be actively modified by varying the spanwise and chordwise positions of the mass element.Among these,the chordwise position exerts a more significant impact on the structural modes and flutter speed of the wing.A hybrid aeroelastic control system,incorporating motion planning and control law,is proposed to evaluate real-time performance in Active Flutter Suppression(AFS)and Gust Load Alleviation(GLA).Control outcomes suggest that,with a mass ratio of 1/16 and a half-chord installation area for the guide rail,flutter speed increases by about 10%.Additionally,excitation amplitudes across different gust frequencies are substantially mitigated,achieving a maximum reduction of vibration amplitude by about 73%.These findings offer a comprehensive understanding of the MMC technique and its application to flexible aircraft.
基金supported by the National Natural Science Foundation of China(Nos.12272104,U22B2013).
文摘High-aspect-ratio aircraft are widely used in military and civilian fields,such as reconnaissance,surveillance,and attacks,due to their high lift-to-drag ratio,strong payload capability,significant endurance effect,and good stealth performance.However,compared to conventional aircraft,high-aspect-ratio aircraft are more susceptible to gust disturbances during flight.In response to this phenomenon,a full-scale dynamic model of a high-aspect-ratio unmanned aerial vehicle was developed.Considering the coupling among control surfaces,structural forces,and aerodynamic forces,along with sensor,actuator,and delay effects,an H_(∞)control law was designed using the principle of singular value energy flow reduction and weighted function,with a PID(Proportional-Integral-Derivative)control law for comparison.The two controllers were then subjected to pulse-response and jury stability tests.Finally,wind tunnel tests were conducted to investigate the gust alleviation principle,in which gust disturbances were generated using gust generators and control surface self-excitation.The results present that the average wing root bending moment and wing tip overload under the PID control law decrease by approximately 30%,while under the H_(∞)control law,both the average wing root bending moment and wing tip overload reduction rate exceed 50%,with peaks reaching 60%.This validates the feasibility and efficiency of the designed H_(∞)controller.
基金supported by the National Natural Science Foundation of China(Nos.11302011,11172025)the National Natural Science Foundation for Youth of China(No.11402013)
文摘A theoretical nonlinear aeroelastic response analysis for a flexible high-aspect ratio wing excited by harmonic gust load is presented along with a companion wind tunnel test. A multidisci- plinary coupled numerical calculation is developed to simulate the flexible model wing undergoing gust load in the time domain via discrete nonlinear finite element structural dynamic analysis and nonplanar unsteady vortex lattice aerodynamic computation. A dynamic perturbation analysis about a nonlinear static equilibrium is also used to determine the small perturbation flutter bound- ary. A novel noncontact 3-D camera measurement analysis system is firstly used in the wind tunnel test to obtain the spatial large deformation and responses. The responses of the flexible wing under different static equilibrium states and frequency gust loads are discussed. The fair to good quanti- tative agreements between the theoretical and experimental results demonstrate that the presented analysis method is an acceptable way to predict the geometrically nonlinear gust response for flex- ible wings.
文摘The aim of this study is to identify the influence of the dip angle of a pre-existing macrocrack on the lifetime and ultimate deformation of rock-like material. Prediction of lifetime has been studied for three groups of specimens under axial static compressive load levels. The specimens were investigated from 65% to 85% of UCS(uniaxial compressive strength) at an interval of 10% of UCS for the groups of specimens with a single modelled open flaw with a dip angle to the loading direction of 30°(first group), at an interval of 5% of UCS increment for the groups of specimens with single(second group), and double sequential open flaws with a dip angle to the loading direction of 60°(third group). This study shows that crack propagation in specimens with a single flaw follows the same sequences. At first, wing cracks appear, and then shear crack develops from the existing wing cracks. Shear cracking is responsible for specimen failure in all three groups. A slip is expected in specimens from the third group which connects two individual modelled open flaws. The moment of the slip is noticed as a characteristic rise in the axial deformation at a constant load level. It is also observed that axial deformation versus time follows the same pattern, irrespective of local geometry. Specimens from the first group exhibit higher axial deformation under different load levels in comparison with the specimens from the second and third groups.
基金supported by the National Natural Science Foundation of China(11072142)Shanghai Program for Innovative Research Team in Universities
文摘The tip vortices and aerodynamics of a NACA0012 wing in the vicinity of the ground were studied in a wind tunnel.The wing tip vortex structures and lift/drag forces were measured by a seven-hole probe and a force balance,respectively.The evolution of the flow structures and aerodynamics with a ground height were analyzed.The vorticity of tip vortices was found to reduce with the decreasing of the ground height,and the position of vortex-core moved gradually to the outboard of the wing tip.Therefore,the down-wash flow induced by the tip vortices was weakened. However,vortex breakdown occurred as the wing lowered to the ground.From the experimental results of aerodynamics,the maximum lift-to-drag ratio was observed when the angle of attack was 2.5°and the ground clearance was 0.2.
基金Supported by the National Nature Science Foundation of China(Grant No.52192631 and No.52105013).
文摘Fixed-wing aircraft cannot maintain optimal aerodynamic performance at different flight speeds. As a type of morphing aircraft, the shear variable-sweep wing(SVSW) can dramatically improve its aerodynamic performance by altering its shape to adapt to various flight conditions.In order to achieve smooth continuous shear deformation, SVSW's skin adopts a flexible composite skin design instead of traditional aluminum alloy materials. However, this also brings about the non-linear difficulty in stiffness modeling and calculation. In this research, a new SVSW design and efficient stiffness modeling method are proposed. Based on shear deformation theory, the flexible composite skin is equivalently modeled as diagonally arranged nonlinear springs, simulating the elastic force interaction between the skin and the mechanism. By shear loading tests of flexible composite skin, the accuracy of this flexible composite skin modeling method is verified. The SVSW stiffness model was established, and its accuracy was verified through static loading tests. The effects of root connection, sweep angles, and flexible composite skin on the SVSW stiffness are analyzed. Finally, considering three typical flight conditions of SVSW: low-speed flow(Ma = 0.3,Re = 5.82 × 10^(6)), transonic flow(Ma = 0.9, Re = 3.44 × 10^(6)), and supersonic flow(Ma = 3,Re = 7.51 × 10^(6)), the stiffness characteristics of SVSW under flight conditions were evaluated.The calculated results guide the application of SVSW.
