It is a major challenge for the airframe-inlet design of modern combat aircrafts,as the flow and electromagnetic wave propagation in the inlet of stealth aircraft are very complex.In this study,an aerodynamic/stealth ...It is a major challenge for the airframe-inlet design of modern combat aircrafts,as the flow and electromagnetic wave propagation in the inlet of stealth aircraft are very complex.In this study,an aerodynamic/stealth optimization design method for an S-duct inlet is proposed.The upwind scheme is introduced to the aerodynamic adjoint equation to resolve the shock wave and flow separation.The multilevel fast multipole algorithm(MLFMA)is utilized for the stealth adjoint equation.A dorsal S-duct inlet of flying wing layout is optimized to improve the aerodynamic and stealth characteristics.Both the aerodynamic and stealth characteristics of the inlet are effectively improved.Finally,the optimization results are analyzed,and it shows that the main contradiction between aerodynamic characteristics and stealth characteristics is the centerline and crosssectional area.The S-duct is smoothed,and the cross-sectional area is increased to improve the aerodynamic characteristics,while it is completely opposite for the stealth design.The radar cross section(RCS)is reduced by phase cancelation for low frequency conditions.The method is suitable for the aerodynamic/stealth design of the aircraft airframe-inlet system.展开更多
Variable-fidelity optimization (VFO) has emerged as an attractive method of performing, both, high-speed and high-fidelity optimization. VFO uses computationally inexpensive low-fidelity models, complemented by a surr...Variable-fidelity optimization (VFO) has emerged as an attractive method of performing, both, high-speed and high-fidelity optimization. VFO uses computationally inexpensive low-fidelity models, complemented by a surrogate to account for the difference between the high-and low-fidelity models, to obtain the optimum of the function efficiently and accurately. To be effective, however, it is of prime importance that the low fidelity model be selected prudently. This paper outlines the requirements for selecting the low fidelity model and shows pitfalls in case the wrong model is chosen. It then presents an efficient VFO framework and demonstrates it by performing transonic airfoil drag optimization at constant lift, subject to thickness constraints, using several low fidelity solvers. The method is found to be efficient and capable of finding the optimum that closely agrees with the results of high-fidelity optimization alone.展开更多
A new approach for selecting proper discretization schemes and grid size is presented. This method is based on the convection-diffusion equation and can provide insight for the Navier-Stokes equation. The approach mai...A new approach for selecting proper discretization schemes and grid size is presented. This method is based on the convection-diffusion equation and can provide insight for the Navier-Stokes equation. The approach mainly addresses two aspects, i.e., the practical accuracy of diffusion term discretization and the behavior of high wavenum- ber disturbances. Two criteria are included in this approach. First, numerical diffusion should not affect the theoretical diffusion accuracy near the length scales of interest. This is achieved by requiring numerical diffusion to be smaller than the diffusion discretization error. Second, high wavenumber modes that are.much smaller than the length scales of interest should not be amplified. These two criteria provide a range of suitable scheme combinations for convective flux and diffusive flux and an ideal interval for grid spacing. The effects of time discretization on these criteria are briefly discussed.展开更多
Mitigating the sonic boom to an acceptable stage is crucial for the next generation of supersonic transports.The primary way to suppress sonic booms is to develop a low sonic boom aerodynamic shape design.This paper p...Mitigating the sonic boom to an acceptable stage is crucial for the next generation of supersonic transports.The primary way to suppress sonic booms is to develop a low sonic boom aerodynamic shape design.This paper proposes an inverse design approach to optimize the near-field signature of an aircraft,making it close to the shaped ideal ground signature after propagation in the atmosphere.By introducing the Deep Neural Network(DNN)model for the first time,a predicted input of Augmented Burgers equation is inversely achieved.By the K-fold cross-validation method,the pre-dicted ground signature closest to the target ground signature is obtained.Then,the corresponding equivalent area distribution is calculated using the classical Whitham’s F-function theory from the optimal near-field signature.The inversion method is vali-dated using the classic example of the C608 vehicle provided by the Third Sonic Boom Prediction Workshop(SBPW-3).The results show that the design ground signature is consistent with the target signature.The equivalent area distribution of the design result is smoother than the baseline distribution,and it shrinks significantly in the rear section.Finally,the robustness of this method is verified through the inverse design of sonic boom for the non-physical ground signature target.展开更多
Mitigation of sonic boom to an acceptable stage is a key point for the next generation of supersonic transports. Meanwhile, designing a supersonic aircraft with an ideal ground signature is always the focus of researc...Mitigation of sonic boom to an acceptable stage is a key point for the next generation of supersonic transports. Meanwhile, designing a supersonic aircraft with an ideal ground signature is always the focus of research on sonic boom reduction. This paper presents an inverse design approach to optimize the near-field signature of an aircraft, making it close to the shaped ideal ground signature after the propagation in the atmosphere. Using the Proper Orthogonal Decomposition(POD) method, a guessed input of augmented Burgers equation is inversely achieved. By multiple POD iterations, the guessed ground signatures successively approach the target ground signature until the convergence criteria is reached. Finally, the corresponding equivalent area distribution is calculated from the optimal near-field signature through the classical Whitham F-function theory. To validate this method, an optimization example of Lockheed Martin 1021 is demonstrated. The modified configuration has a fully shaped ground signature and achieves a drop of perceived loudness by 7.94 PLdB. This improvement is achieved via shaping the original near-field signature into wiggles and damping it by atmospheric attenuation. At last, a nonphysical ground signature is set as the target to test the robustness of this inverse design method and shows that this method is robust enough for various inputs.展开更多
Common,unsteady aerodynamic modeling methods usually use wind tunnel test data from forced vibration tests to predict stable hysteresis loop.However,these methods ignore the initial unstable process of entering the hy...Common,unsteady aerodynamic modeling methods usually use wind tunnel test data from forced vibration tests to predict stable hysteresis loop.However,these methods ignore the initial unstable process of entering the hysteresis loop that exists in the actual maneuvering process of the aircraft.Here,an excitation input suitable for nonlinear system identification is introduced to model unsteady aerodynamic forces with any motion in the amplitude and frequency ranges based on the Least Squares Support Vector Machines(LS-SVMs).In the selection of the input form,avoiding the use of reduced frequency as a parameter makes the model more universal.After model training is completed,the method is applied to predict the lift coefficient,drag coefficient and pitching moment coefficient of the RAE2822 airfoil,in sine and sweep motions under the conditions of plunging and pitching at Mach number 0.8.The predicted results of the initial unstable process and the final stable process are in close agreement with the Computational Fluid Dynamics(CFD)data,demonstrating the feasibility of the model for nonlinear unsteady aerodynamics modeling and the effectiveness of the input design approach.展开更多
The achievement of laminar flow in the boundary layer at high-speed cruise conditions may further, in addition to shock-wave control, reduce the drag and extend the range of military fighter aircraft. To this end, a f...The achievement of laminar flow in the boundary layer at high-speed cruise conditions may further, in addition to shock-wave control, reduce the drag and extend the range of military fighter aircraft. To this end, a further investigation on transitional boundary-layer flow of fighter wings is needed due to different configurations from the wings used on conventional transport aircraft. In this paper, wind tunnel experiments and numerical simulations were conducted on three-dimensional transition of thin diamond-shaped wings used on advanced fighter aircraft at tran/supersonic design points. A newly proposed correlation of crossflow transition which includes the effect of surface roughness was introduced into the c-Rehttransition model. Predicted results were in good agreement with flow visualizations. Results showed that the strength of the crossflow component grew rapidly around the leading edge because of the severe flow acceleration, just as the same as wings with a large aspect ratio. However, there seemed no regular pattern of instabilitydominance variation in span-wise for a diamond configuration. The dominance of different instability mechanisms strongly depended on the local pressure distribution. Hereby, the research recommended a ‘‘roof-like" shape of pressure distribution to suppress both crossflow and Tollmien-Schlichting(T-S) instabilities. Besides, a sharp suction peak with a serious pressure rise should be cut off to avoid stronger instabilities. Further discussions also revealed an independence of the unit Reynolds number when transition was triggered by T-S instabilities. Aerodynamic force comparisons indicated that further benefit on drag reduction could be expected by including the three-dimensional transition effect into a wing design process.展开更多
This paper develops a low-diffusion robust flux splitting scheme termed TVAP to achieve the simulation of wide-ranging Mach number flows.Based on Toro-Vazquez splitting approach,the new scheme splits inviscid flux int...This paper develops a low-diffusion robust flux splitting scheme termed TVAP to achieve the simulation of wide-ranging Mach number flows.Based on Toro-Vazquez splitting approach,the new scheme splits inviscid flux into convective system and pressure system.This method introduces Mach number splitting function and numerical sound speed to evaluate advection system.Meanwhile,pressure diffusion term,pressure momentum flux,interface pressure and interface velocity are constructed to measure pressure system.Then,typical test problems are utilized to systematically assess the robustness and accuracy of the resulting scheme.Matrix stability analysis and a series of numerical cases,such as double shear-layer problem and hypersonic viscous flow over blunt cone,demonstrate that TVAP scheme achieves excellent low diffusion,shock stability,contact discontinuity and low-speed resolution,and is potentially a good candidate for wide-ranging Mach number flows.展开更多
The aerodynamic layout of the Canard Rotor/Wing(CRW) aircraft in helicopter flight mode differs significantly from that of conventional helicopters. In order to study the flight dynamics characteristics of CRW aircraf...The aerodynamic layout of the Canard Rotor/Wing(CRW) aircraft in helicopter flight mode differs significantly from that of conventional helicopters. In order to study the flight dynamics characteristics of CRW aircraft in helicopter mode, first, the aerodynamic model of the main rotor system is established based on the blade element theory and wind tunnel test results. The aerodynamic forces and moments of the canard wing, horizontal tail, vertical tail and fuselage are obtained via theoretical analysis and empirical formula. The flight dynamics model of the CRW aircraft in helicopter mode is developed and validated by flight test data. Next, a method of model trimming using an optimization algorithm is proposed. The flight dynamics characteristics of the CRW are investigated by the method of linearized small perturbations via Simulink. The trim results are consistent with the conventional helicopter characteristics, and the results show that with increasing forward flight speed, the canard wing and horizontal tail can provide considerable lift,which reflects the unique characteristics of the CRW aircraft. Finally, mode analysis is implemented for the linearized CRW in helicopter mode. The results demonstrate that the stability of majority modes increases with increasing flight speed. However, one mode that diverges monotonously,and the reason is that the CRW helicopter mode has a large vertical tail compared to the conventional helicopter. The results of the dynamic analysis provide optimization guidance and reference for the overall design of the CRW aircraft in helicopter mode, and the model developed can be used for control system design.展开更多
An array of distributed round synthetic jets was used to control a fully developed turbulent boundary layer.The study focused on the related skin friction drag reduction and mechanisms involved.The control effects wer...An array of distributed round synthetic jets was used to control a fully developed turbulent boundary layer.The study focused on the related skin friction drag reduction and mechanisms involved.The control effects were analyzed by measuring the streamwise velocities using a hot-wire anemometer downstream of the array.A reduction in the skin friction was observed both in the regions downstream of the orifices and in the regions between two adjacent orifices.A statistical analysis with the variable-interval time-averaging(VITA)technique demonstrated a weakened bursting intensity with synthetic jet in the near-wall region.The streamwise vortices were lifted by the upwash effect caused by synthetic jet and induced less low-speed streaks.The control mechanism acted in a way to suppress the dynamic interaction between the streamwise vortices and low-speed streaks and to attenuate the turbulence production in the near-wall region.The forcing frequency was found to be a more relevant parameter when synthetic jet was applied in turbulent boundary layer flow control.A higher forcing frequency induced a higher reduction in the skin friction.The power spectral density and autocorrelation of the fluctuating velocities showed that the synthetic jets gradually decayed in the streamwise direction,having an effect as far as 34.5 times the displacement thickness that was on the trailing edge of the distributed synthetic jets array.展开更多
Based on improved multi-objective particle swarm optimization（MOPSO） algorithm with principal component analysis（PCA） methodology, an efficient high-dimension multiobjective optimization method is proposed, which,...Based on improved multi-objective particle swarm optimization（MOPSO） algorithm with principal component analysis（PCA） methodology, an efficient high-dimension multiobjective optimization method is proposed, which, as the purpose of this paper, aims to improve the convergence of Pareto front in multi-objective optimization design. The mathematical efficiency,the physical reasonableness and the reliability in dealing with redundant objectives of PCA are verified by typical DTLZ5 test function and multi-objective correlation analysis of supercritical airfoil,and the proposed method is integrated into aircraft multi-disciplinary design（AMDEsign） platform, which contains aerodynamics, stealth and structure weight analysis and optimization module.Then the proposed method is used for the multi-point integrated aerodynamic optimization of a wide-body passenger aircraft, in which the redundant objectives identified by PCA are transformed to optimization constraints, and several design methods are compared. The design results illustrate that the strategy used in this paper is sufficient and multi-point design requirements of the passenger aircraft are reached. The visualization level of non-dominant Pareto set is improved by effectively reducing the dimension without losing the primary feature of the problem.展开更多
The Stopped-Rotor(SR)UAV combines the advantages of vertical take-off and landing of helicopter and high-speed cruise of fixed-wing aircraft.At the same time,it also has a unique aerodynamic layout,which leads to grea...The Stopped-Rotor(SR)UAV combines the advantages of vertical take-off and landing of helicopter and high-speed cruise of fixed-wing aircraft.At the same time,it also has a unique aerodynamic layout,which leads to great differences in the control and aerodynamic characteristics of various flight modes,and brings great challenges to the flight dynamics modelling and control in full-mode flight.In this paper,the flight dynamics modelling and control method of SR UAV in full-mode flight is studied.First,based on the typical flight profile of SR UAV when performing missions,using the theory and method of fuzzy mathematics,the T-S flight dynamics model of SR UAV in full-mode flight is established by synthesizing the flight dynamics model of each flight mode.Then,an explicit model tracking and parameter adjusting control system based on fuzzy theory is designed to enhance the stability of the inner loop of SR UAV in full-mode flight,which effectively reduces the coupling between axes and improves the control quality of the system.Finally,the outer loop control system is designed by using classical control method,and the control law of SR UAV in full-mode automatic flight is obtained.The simulation results show that the proposed control system design method is feasible and effective,which lays a solid foundation for the subsequent engineering implementation of the SR UAV.展开更多
文摘It is a major challenge for the airframe-inlet design of modern combat aircrafts,as the flow and electromagnetic wave propagation in the inlet of stealth aircraft are very complex.In this study,an aerodynamic/stealth optimization design method for an S-duct inlet is proposed.The upwind scheme is introduced to the aerodynamic adjoint equation to resolve the shock wave and flow separation.The multilevel fast multipole algorithm(MLFMA)is utilized for the stealth adjoint equation.A dorsal S-duct inlet of flying wing layout is optimized to improve the aerodynamic and stealth characteristics.Both the aerodynamic and stealth characteristics of the inlet are effectively improved.Finally,the optimization results are analyzed,and it shows that the main contradiction between aerodynamic characteristics and stealth characteristics is the centerline and crosssectional area.The S-duct is smoothed,and the cross-sectional area is increased to improve the aerodynamic characteristics,while it is completely opposite for the stealth design.The radar cross section(RCS)is reduced by phase cancelation for low frequency conditions.The method is suitable for the aerodynamic/stealth design of the aircraft airframe-inlet system.
文摘Variable-fidelity optimization (VFO) has emerged as an attractive method of performing, both, high-speed and high-fidelity optimization. VFO uses computationally inexpensive low-fidelity models, complemented by a surrogate to account for the difference between the high-and low-fidelity models, to obtain the optimum of the function efficiently and accurately. To be effective, however, it is of prime importance that the low fidelity model be selected prudently. This paper outlines the requirements for selecting the low fidelity model and shows pitfalls in case the wrong model is chosen. It then presents an efficient VFO framework and demonstrates it by performing transonic airfoil drag optimization at constant lift, subject to thickness constraints, using several low fidelity solvers. The method is found to be efficient and capable of finding the optimum that closely agrees with the results of high-fidelity optimization alone.
基金Project supported by the National Natural Science Foundation of China(No.11372254)
文摘A new approach for selecting proper discretization schemes and grid size is presented. This method is based on the convection-diffusion equation and can provide insight for the Navier-Stokes equation. The approach mainly addresses two aspects, i.e., the practical accuracy of diffusion term discretization and the behavior of high wavenum- ber disturbances. Two criteria are included in this approach. First, numerical diffusion should not affect the theoretical diffusion accuracy near the length scales of interest. This is achieved by requiring numerical diffusion to be smaller than the diffusion discretization error. Second, high wavenumber modes that are.much smaller than the length scales of interest should not be amplified. These two criteria provide a range of suitable scheme combinations for convective flux and diffusive flux and an ideal interval for grid spacing. The effects of time discretization on these criteria are briefly discussed.
基金the National Key Research and Development Program of China(No.2020YFB1709500)Natural Science Basic Research Program of Shaanxi province(No.2021JQ-076)Fundamental Research Funds for the Central Universities.
文摘Mitigating the sonic boom to an acceptable stage is crucial for the next generation of supersonic transports.The primary way to suppress sonic booms is to develop a low sonic boom aerodynamic shape design.This paper proposes an inverse design approach to optimize the near-field signature of an aircraft,making it close to the shaped ideal ground signature after propagation in the atmosphere.By introducing the Deep Neural Network(DNN)model for the first time,a predicted input of Augmented Burgers equation is inversely achieved.By the K-fold cross-validation method,the pre-dicted ground signature closest to the target ground signature is obtained.Then,the corresponding equivalent area distribution is calculated using the classical Whitham’s F-function theory from the optimal near-field signature.The inversion method is vali-dated using the classic example of the C608 vehicle provided by the Third Sonic Boom Prediction Workshop(SBPW-3).The results show that the design ground signature is consistent with the target signature.The equivalent area distribution of the design result is smoother than the baseline distribution,and it shrinks significantly in the rear section.Finally,the robustness of this method is verified through the inverse design of sonic boom for the non-physical ground signature target.
文摘Mitigation of sonic boom to an acceptable stage is a key point for the next generation of supersonic transports. Meanwhile, designing a supersonic aircraft with an ideal ground signature is always the focus of research on sonic boom reduction. This paper presents an inverse design approach to optimize the near-field signature of an aircraft, making it close to the shaped ideal ground signature after the propagation in the atmosphere. Using the Proper Orthogonal Decomposition(POD) method, a guessed input of augmented Burgers equation is inversely achieved. By multiple POD iterations, the guessed ground signatures successively approach the target ground signature until the convergence criteria is reached. Finally, the corresponding equivalent area distribution is calculated from the optimal near-field signature through the classical Whitham F-function theory. To validate this method, an optimization example of Lockheed Martin 1021 is demonstrated. The modified configuration has a fully shaped ground signature and achieves a drop of perceived loudness by 7.94 PLdB. This improvement is achieved via shaping the original near-field signature into wiggles and damping it by atmospheric attenuation. At last, a nonphysical ground signature is set as the target to test the robustness of this inverse design method and shows that this method is robust enough for various inputs.
文摘Common,unsteady aerodynamic modeling methods usually use wind tunnel test data from forced vibration tests to predict stable hysteresis loop.However,these methods ignore the initial unstable process of entering the hysteresis loop that exists in the actual maneuvering process of the aircraft.Here,an excitation input suitable for nonlinear system identification is introduced to model unsteady aerodynamic forces with any motion in the amplitude and frequency ranges based on the Least Squares Support Vector Machines(LS-SVMs).In the selection of the input form,avoiding the use of reduced frequency as a parameter makes the model more universal.After model training is completed,the method is applied to predict the lift coefficient,drag coefficient and pitching moment coefficient of the RAE2822 airfoil,in sine and sweep motions under the conditions of plunging and pitching at Mach number 0.8.The predicted results of the initial unstable process and the final stable process are in close agreement with the Computational Fluid Dynamics(CFD)data,demonstrating the feasibility of the model for nonlinear unsteady aerodynamics modeling and the effectiveness of the input design approach.
基金supported by the National Natural Science Foundation of China (No.11372254)
文摘The achievement of laminar flow in the boundary layer at high-speed cruise conditions may further, in addition to shock-wave control, reduce the drag and extend the range of military fighter aircraft. To this end, a further investigation on transitional boundary-layer flow of fighter wings is needed due to different configurations from the wings used on conventional transport aircraft. In this paper, wind tunnel experiments and numerical simulations were conducted on three-dimensional transition of thin diamond-shaped wings used on advanced fighter aircraft at tran/supersonic design points. A newly proposed correlation of crossflow transition which includes the effect of surface roughness was introduced into the c-Rehttransition model. Predicted results were in good agreement with flow visualizations. Results showed that the strength of the crossflow component grew rapidly around the leading edge because of the severe flow acceleration, just as the same as wings with a large aspect ratio. However, there seemed no regular pattern of instabilitydominance variation in span-wise for a diamond configuration. The dominance of different instability mechanisms strongly depended on the local pressure distribution. Hereby, the research recommended a ‘‘roof-like" shape of pressure distribution to suppress both crossflow and Tollmien-Schlichting(T-S) instabilities. Besides, a sharp suction peak with a serious pressure rise should be cut off to avoid stronger instabilities. Further discussions also revealed an independence of the unit Reynolds number when transition was triggered by T-S instabilities. Aerodynamic force comparisons indicated that further benefit on drag reduction could be expected by including the three-dimensional transition effect into a wing design process.
基金supported by the Space Science and Technology Fund Project of China(No.2020-HT-XG-14)。
文摘This paper develops a low-diffusion robust flux splitting scheme termed TVAP to achieve the simulation of wide-ranging Mach number flows.Based on Toro-Vazquez splitting approach,the new scheme splits inviscid flux into convective system and pressure system.This method introduces Mach number splitting function and numerical sound speed to evaluate advection system.Meanwhile,pressure diffusion term,pressure momentum flux,interface pressure and interface velocity are constructed to measure pressure system.Then,typical test problems are utilized to systematically assess the robustness and accuracy of the resulting scheme.Matrix stability analysis and a series of numerical cases,such as double shear-layer problem and hypersonic viscous flow over blunt cone,demonstrate that TVAP scheme achieves excellent low diffusion,shock stability,contact discontinuity and low-speed resolution,and is potentially a good candidate for wide-ranging Mach number flows.
文摘The aerodynamic layout of the Canard Rotor/Wing(CRW) aircraft in helicopter flight mode differs significantly from that of conventional helicopters. In order to study the flight dynamics characteristics of CRW aircraft in helicopter mode, first, the aerodynamic model of the main rotor system is established based on the blade element theory and wind tunnel test results. The aerodynamic forces and moments of the canard wing, horizontal tail, vertical tail and fuselage are obtained via theoretical analysis and empirical formula. The flight dynamics model of the CRW aircraft in helicopter mode is developed and validated by flight test data. Next, a method of model trimming using an optimization algorithm is proposed. The flight dynamics characteristics of the CRW are investigated by the method of linearized small perturbations via Simulink. The trim results are consistent with the conventional helicopter characteristics, and the results show that with increasing forward flight speed, the canard wing and horizontal tail can provide considerable lift,which reflects the unique characteristics of the CRW aircraft. Finally, mode analysis is implemented for the linearized CRW in helicopter mode. The results demonstrate that the stability of majority modes increases with increasing flight speed. However, one mode that diverges monotonously,and the reason is that the CRW helicopter mode has a large vertical tail compared to the conventional helicopter. The results of the dynamic analysis provide optimization guidance and reference for the overall design of the CRW aircraft in helicopter mode, and the model developed can be used for control system design.
基金The authors would like to acknowledge the financial support received from the project“Drag Reduction via Turbulent Boundary Layer Flow Control(DRAGY)”.The DRAGY project(April 2016-March 2019)is a China-EU Aeronautical Cooperation project,which is co-funded by Ministry of Industry and Information Technology(MIIT),China,and Directorate-General for Research and Innovation(DG RTD),European Commission.
文摘An array of distributed round synthetic jets was used to control a fully developed turbulent boundary layer.The study focused on the related skin friction drag reduction and mechanisms involved.The control effects were analyzed by measuring the streamwise velocities using a hot-wire anemometer downstream of the array.A reduction in the skin friction was observed both in the regions downstream of the orifices and in the regions between two adjacent orifices.A statistical analysis with the variable-interval time-averaging(VITA)technique demonstrated a weakened bursting intensity with synthetic jet in the near-wall region.The streamwise vortices were lifted by the upwash effect caused by synthetic jet and induced less low-speed streaks.The control mechanism acted in a way to suppress the dynamic interaction between the streamwise vortices and low-speed streaks and to attenuate the turbulence production in the near-wall region.The forcing frequency was found to be a more relevant parameter when synthetic jet was applied in turbulent boundary layer flow control.A higher forcing frequency induced a higher reduction in the skin friction.The power spectral density and autocorrelation of the fluctuating velocities showed that the synthetic jets gradually decayed in the streamwise direction,having an effect as far as 34.5 times the displacement thickness that was on the trailing edge of the distributed synthetic jets array.
基金supported by the National Natural Science Foundation of China (No.11402288)
文摘Based on improved multi-objective particle swarm optimization（MOPSO） algorithm with principal component analysis（PCA） methodology, an efficient high-dimension multiobjective optimization method is proposed, which, as the purpose of this paper, aims to improve the convergence of Pareto front in multi-objective optimization design. The mathematical efficiency,the physical reasonableness and the reliability in dealing with redundant objectives of PCA are verified by typical DTLZ5 test function and multi-objective correlation analysis of supercritical airfoil,and the proposed method is integrated into aircraft multi-disciplinary design（AMDEsign） platform, which contains aerodynamics, stealth and structure weight analysis and optimization module.Then the proposed method is used for the multi-point integrated aerodynamic optimization of a wide-body passenger aircraft, in which the redundant objectives identified by PCA are transformed to optimization constraints, and several design methods are compared. The design results illustrate that the strategy used in this paper is sufficient and multi-point design requirements of the passenger aircraft are reached. The visualization level of non-dominant Pareto set is improved by effectively reducing the dimension without losing the primary feature of the problem.
基金the Natural Science Foundation of China (No. 12102345)the Natural Science Basic Research Program of Shaanxi Province (Nos. 2021JQ-086 and 2021JQ076)Taicang Scientific Research Institute Innovation Leading Special Plan (No. tc2019dyds11)
文摘The Stopped-Rotor(SR)UAV combines the advantages of vertical take-off and landing of helicopter and high-speed cruise of fixed-wing aircraft.At the same time,it also has a unique aerodynamic layout,which leads to great differences in the control and aerodynamic characteristics of various flight modes,and brings great challenges to the flight dynamics modelling and control in full-mode flight.In this paper,the flight dynamics modelling and control method of SR UAV in full-mode flight is studied.First,based on the typical flight profile of SR UAV when performing missions,using the theory and method of fuzzy mathematics,the T-S flight dynamics model of SR UAV in full-mode flight is established by synthesizing the flight dynamics model of each flight mode.Then,an explicit model tracking and parameter adjusting control system based on fuzzy theory is designed to enhance the stability of the inner loop of SR UAV in full-mode flight,which effectively reduces the coupling between axes and improves the control quality of the system.Finally,the outer loop control system is designed by using classical control method,and the control law of SR UAV in full-mode automatic flight is obtained.The simulation results show that the proposed control system design method is feasible and effective,which lays a solid foundation for the subsequent engineering implementation of the SR UAV.
