This study is dedicated to numerically investigate the dynamic behavior of a missile released from a helicopter under the influence of downwash from the rotating rotors using AGM-114 Hellfire and UH-60 as the testcase...This study is dedicated to numerically investigate the dynamic behavior of a missile released from a helicopter under the influence of downwash from the rotating rotors using AGM-114 Hellfire and UH-60 as the testcase.Simulations are conducted using unsteady Reynolds-averaged Navier-Stokes(URANS)with shear stress transport(SST)k-ωturbulence model,incorporating six-degree-of-freedom(6-DOF)motion and overset grid.Two releasing scenarios,viz.,hover and forward flight,are analyzed under varying missile launch thrust and helicopter forward flight speed.Results reveal that the rotor downwash significantly affects the stability of the missile,particularly during hovering case,where low thrust prolongs wake interaction.In forward flight,the increased airspeed can in principle reduce wake influence but introduces asymmetrical aerodynamic effects on the trajectory of the missile.The findings offer guidance for missile release and launch planning in rotorcraft operations.展开更多
Accurate measurement of helicopter rotor motion parameters(flap,lead-lag,torsion,and azimuth angles)is essential for rotor blade design,helicopter dynamics modeling,and flight safety and health monitoring.However,the ...Accurate measurement of helicopter rotor motion parameters(flap,lead-lag,torsion,and azimuth angles)is essential for rotor blade design,helicopter dynamics modeling,and flight safety and health monitoring.However,the existing methods face challenges in testing equipment installation,calibration,and data transmission,resulting in limited reports on real-time in-flight measurements of blade motion parameters.This paper proposes a non-contact optoelectronic method based on two-dimensional position-sensitive detectors for in-flight measurement and a ground calibration system to obtain real-time rotor motion parameters during helicopter flight.The proposed method establishes the time evolution relationship of rotor motion parameters and verifies the performance of the in-flight measurement system regarding measurement resolution and accuracy through the construction of a blade motion posture experimental platform.The proposed method has been applied to the flight measurement of a medium-sized single-rotor helicopter,and the obtained results have been compared with theoretical analysis outcomes.Furthermore,this paper examines the characteristics of blade motion parameters during flight and discusses the challenges and potential solutions for measuring rotor motion parameters during helicopter flight using the proposed method.展开更多
To achieve high-precision trajectory following during helicopter maneuver tasks and reduce the disruptive influences of unknown variabilities,this study introduces a cascaded-loop helicopter trajectory tracking contro...To achieve high-precision trajectory following during helicopter maneuver tasks and reduce the disruptive influences of unknown variabilities,this study introduces a cascaded-loop helicopter trajectory tracking controller,whose parameters are set using an Ant Colony OptimizationSlime Mould Algorithm(ACO-SMA).Initially,a nonlinear flight dynamics model of the helicopter is constructed.Observer gain functions and nonlinear feedback from a vibrational suppression function to improve the tracking performance of the controller,addressing issues in disturbance estimation and compensation of the Active Disturbance Rejection Control(ADRC).Simultaneously,a cascaded loop system,comprising an internal attitude loop and an external position loop,is created,and the ant colony-slime mold hybrid algorithm optimizes the system parameters of the trajectory tracking controller.Finally,helicopter trajectory tracking simulation experiments are conducted,including spiral ascending and“8”shape climbing maneuvers.The findings indicate that the ADRC employed for helicopter trajectory tracking exhibits outstanding performance in rejecting disturbances caused by gusts and accurately tracking trajectories.The trajectory tracking controller,whose parameters are optimized by the ACO-SMA,shows higher tracking precision compared to the conventional PID and ADRC,thereby substantially improving the precision of maneuver tasks.展开更多
To reduce the vibration of the Coaxial Helicopter Main Transmission System(CHMTS)considering both level and vertical flight conditions,a vibration evaluation and optimization model for the CHMTS was built.The vibratio...To reduce the vibration of the Coaxial Helicopter Main Transmission System(CHMTS)considering both level and vertical flight conditions,a vibration evaluation and optimization model for the CHMTS was built.The vibration simulation model of the CHMTS was set up by gear dynamics theory and loaded contact analysis.For better evaluation of the system vibration,a vibration evaluation method for the CHMTS was established by the G1 method-variation coefficient method.A hybrid Gravitational Search Algorithm-Simulated Annealing(GSA-SA)algorithm was combined to balance convergence speed and searching accuracy.The principle test was conducted to prove the accuracy of theoretical method,in which the maximum relative error is16.26%.The optional results show that the vibration of the optimized transmission system decreases significantly,in which the maximum reduction of key vibration indicators reaches more than 20%.The theoretical results have been compared to the experiment to verify the effectiveness of the vibration optimization method.The proposed method could be extended to other fields.展开更多
This paper presents a comprehensive test and systematic evaluation analysis of cabin noise in the Robinson R44 RAVEN II helicopter.Initially,microphones were placed within the cabin to conduct systematic assessments o...This paper presents a comprehensive test and systematic evaluation analysis of cabin noise in the Robinson R44 RAVEN II helicopter.Initially,microphones were placed within the cabin to conduct systematic assessments of noise levels under various flight conditions,including takeoff,climbing,level flight,landing,hovering,etc.Subsequently,time–frequency analysis was conducted on the test data utilizing traditional A-weighted sound pressure levels,which was followed by quantitative comparisons across different flight conditions.Then,detailed evaluation and discussion were conducted,taking into account the subjective perceptions and communication challenges of cabin crew members.This assessment incorporated the use of aviation noise indicators,speech interference levels,and metrics related to sound quality.Finally,potential noise reduction measures and their effects were preliminarily discussed.The results indicate that helicopter cabin noise exhibited variations across different flight states or positions within the same state,ranging from 87.6 d B(A)to 92.6 dB(A).Discrepancies between A-weighted sound pressure level and psychoacoustic parameters were observed,particularly during hovering states,which indicate that there is a necessity for the combination of multiple evaluation indicators.Notably,damping measure can serve as a pivotal factor in mitigating cabin noise.展开更多
An advanced Actuator Surface Method(ASM)coupled with Computational Fluid Dynamics(CFD)is developed and applied to the complex unsteady aerodynamic simulation of helicopter.By introducing an improved three-dimensional ...An advanced Actuator Surface Method(ASM)coupled with Computational Fluid Dynamics(CFD)is developed and applied to the complex unsteady aerodynamic simulation of helicopter.By introducing an improved three-dimensional anisotropic Gaussian kernel,this method effectively addresses the severe aerodynamic load fluctuations commonly associated with traditional Virtual Blade Method(VBM)due to turbulent flow around blade elements.To manage the issues of regional shape and grid cell quantity variations caused by virtual blade sweeping,a universal hybrid grid generation strategy is established without body-fitted and disk interpolation grids,which enhances the computational stability at both blade elements and blade edges.Aerodynamic numerical simulations of helicopter are performed using this method,focusing on rotor/fuselage interaction dominated by rotor wake motion and fuselage blockage effects,Blade-Vortex Interaction(BVI)induced by tip vortices,and maneuvering flights involving collective pitch ramp increases.The results indicate that the advanced ASM demonstrates reliability and robustness in the simulation of complex unsteady flow fields around helicopter.Under similar computational accuracy,the advanced ASM improves computational efficiency by nearly 40 times compared to the oversetgrid-based full Blade-Resolved(B-R)method,and by 6 times compared to the VBM.It shows significant advantages when applied to complex full-aircraft interaction and maneuvering flight conditions that require substantial computational resources.展开更多
The ballistic impact identification method for the helicopter Tail Drive Shaft System(TDSS)isn't yet comprehensive,which affects helicopter flight safety.This paper proposes a ballistic impact identification metho...The ballistic impact identification method for the helicopter Tail Drive Shaft System(TDSS)isn't yet comprehensive,which affects helicopter flight safety.This paper proposes a ballistic impact identification method for the TDSS based on vibration response analysis.Based on the Johnson-Cook constitutive model and failure criteria,the ballistic impact finite element simulation model is established,which is verified by the ballistic impact experiment of the Tail Drive Shaft(TDS).Considering the ballistic impact excitation force,the dynamic model of the TDSS with ballistic impact is established,which is verified by finite element commercial software.If a bullet hits the TDS,the bending vibration displacement increases sharply at a certain moment and then significantly increases but remains stable.Meanwhile,the critical speed component appears in the frequency-domain response of bending vibration,and then the speed component significantly increases but remains stable.What's more,the axis trajectory exhibits a sudden,large-scale,and irregular whirling motion at a certain moment,followed by a significant increase but remains stable.Furthermore,if the axial vibration response is small,the bullet core shooting should be considered vertically or at a small incident angle,otherwise,it should be considered at a large incident angle.展开更多
The muhi-body analysis of the aeroelastic stability of the tiltrotor aircraft is presented. Muhi-body dynamic differential equations are combined with the equations of the unsteady dynamic inflow model to establish th...The muhi-body analysis of the aeroelastic stability of the tiltrotor aircraft is presented. Muhi-body dynamic differential equations are combined with the equations of the unsteady dynamic inflow model to establish the complete unsteadily aeroelastic coupling analytical model of the tiltrotor. The stability of the tiltrotor in the helicopter mode is analyzed aiming at a semi span soft-inplane tihrotor model with an elastic wing. Parametric effects of the lag stiffness of blades and the flight speed are analyzed. Numerical simulations demonstrate that the multibody analytical model can analyze the aeroelastic stability of the tiltrotor aircraft in the helicopter mode.展开更多
An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from ...An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°展开更多
The particularity and practicality of harmony operations of close-coupling multiple helicopters indicate that the researches on it are urgent and necessary, Using the model that describes two hovering helicopters carr...The particularity and practicality of harmony operations of close-coupling multiple helicopters indicate that the researches on it are urgent and necessary, Using the model that describes two hovering helicopters carrying one heavy load, an inertia coordinate system and body coordinate systems of each sub-system are established. A nonlinear force model is established too. The equilibrium computation results can be regarded as the reference control inputs of the flight control system under hovering or low-speed flight condition. After the establishment of a translation kinematics model and a posture kinematics model, a coupling dynamics model of the multiple helicopter system is set up. The results can also be regarded as the base to analyze stabilization and design a controller for a close-coupling multiple helicopters harmony operation system.展开更多
This article presents a new mathematical model for helicopter comprehensive analysis with the features of flexibility and mathematical simplicity. The model synthesizes the rigid fuselage motion model with 6 degrees o...This article presents a new mathematical model for helicopter comprehensive analysis with the features of flexibility and mathematical simplicity. The model synthesizes the rigid fuselage motion model with 6 degrees of freedom,coupled flap-lag-torsion elastic rotor blade motion model,unsteady aerodynamics model with dynamic stall and high order generalized dynamic wake model. A new blade structural operator with implicit form is formulated,and the components of the blade structure model are independent of each other so that it is convenient to change or handle any component of blade structure without changing the others. What is more,the entire model is developed in a strict state-space form to simplify the comprehensive analysis. Finally,the UH-60 helicopter is taken as an example to predict the blade natural characteristics,the trim characteristics including controls and fuselage attitudes as well as the airloads at blade section under the flight conditions of high speed with moderate thrust and high thrust with moderate speed. The results are compared with UH-60 flight test data and those predicted by two well-known comprehensive codes. The validity of the model presented in this article is verified.展开更多
A direct self-repairing control approach is proposed for helicopter via quantum control techniques and adaptive compensator when some complex faults occur. For a linear varying-parameter helicopter control system, the...A direct self-repairing control approach is proposed for helicopter via quantum control techniques and adaptive compensator when some complex faults occur. For a linear varying-parameter helicopter control system, the model reference adaptive control law is designed and an adaptive compensator is used for improving its self-re- pairing capability. To enhance anti-interference capability of helicopter, quantum control feedforward is added be- tween fault and disturbance. Simulation results illustrate the effectiveness and feasibility of the approach.展开更多
Gust response alleviation is very important for helicopters which have strong coupling and vibration. Gust disturbance not only influences the ride quality and the precision of the weapon delivery, but also affects to...Gust response alleviation is very important for helicopters which have strong coupling and vibration. Gust disturbance not only influences the ride quality and the precision of the weapon delivery, but also affects to the structural fatigue load and the strength. The method of an optimal control law to suppress the gust disturbance for helicopters is presented. The optimization requires the minimization of the vertical overload at the pilot′s seat, the attitude variation and the control energy consumption under the gust disturbance. Based on the original control system, the new system can be easily realized by adding a vertical speed feedback passage. In order to develop the real-time operational flight control system, the optimized control law is written in C language. The hybrid simulations prove that the performance of gust response alleviation and the efficiency of digitalization are satisfactory.展开更多
Aim To extract harmonic frequencies of helicopter acoustic signal as features for hel icopter identification. Methods Estimation of signal parameters via rotational invariance techniques(ESPRIT) was selected to ext...Aim To extract harmonic frequencies of helicopter acoustic signal as features for hel icopter identification. Methods Estimation of signal parameters via rotational invariance techniques(ESPRIT) was selected to extract harmonic frequencies from really measured helicopter acoustic signal and an algorithm based on the SVD TLS was used. Results ESPRIT correctly extracted harmonic frequencies of helicopter using the data of limited length under the variousflight conditions. Conclusion ESPRIT is an effective method of extracting harmonic frequencies and using harmonic frequencies of helicopter acoustic signal to recognize helicopter is feasible.展开更多
Helicopter mathematical model mainly depends on design helicopter control system, flight simulator, and real time control simulation system. But it is difficult to establish a helicopter flight dynamics mathematical ...Helicopter mathematical model mainly depends on design helicopter control system, flight simulator, and real time control simulation system. But it is difficult to establish a helicopter flight dynamics mathematical model that has features such as rapidness, reliability and precision, because there is no unique and precise expression to some sophisticated phenomenon of helicopter. In this paper a fuzzy helicopter flight model is constructed based on the flight experimental data. The fuzzy model, which is identified by fuzzy inference, has characteristics of computed rapidness and high precision. In order to guarantee the precision of the identified fuzzy model, a new method is adopted to handle the conflict fuzzy rules. Additionally, using fuzzy clustering technology can effectively reduce the number of rules of fuzzy model, namely, the order of the fuzzy model. The simulation results indicate that the method of this paper is effective and feasible.展开更多
Adaptive flight control technology, feedback linearization, model inversion theory are reviewed and the error dynamic characteristics are analyzed, and an adaptive on-line neural network attitude control system is pre...Adaptive flight control technology, feedback linearization, model inversion theory are reviewed and the error dynamic characteristics are analyzed, and an adaptive on-line neural network attitude control system is presented. The model inversion is under the hover condition. And the adaptive control law based on the neural network is designed to guarantee the boundedness of tracking error and control signals. Simulation results demonstrate that the nonlinear neural network augmented model inversion can self-adapt to the uncertainty and modeling errors of unmanned helicopters. Results are compared while the parameters of PD controller and robustness items are changed.展开更多
The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled ...The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled rigid/elastic blended element based on the fle xible multibody system theory in this paper. It accounts for the effects of prec one, sweep, and the moderately large elastic deflections on the blade and elasti city of shaft and fuselage of the helicopter. The dynamic coupling between the r igid motion of blades about the flap, lag and pitch hinges of articulated rotor and moderately large elastic deflections are included. There is no restriction o n the rotation amplitudes of flap, lag and pitch in the formulation. The stabili ty of periodic solution is studied using the Floquet theory. The transition matr ix is calculated by the Newmark integration method. The aeromechanical stability of a new helicopter is studied. The results show that it is stable in the given forward flight. But the instability arises with the decrease of the bending and torsion stiffness of the shaft.展开更多
基金supported by the National Natural Science Foundation of China(No.11602200)。
文摘This study is dedicated to numerically investigate the dynamic behavior of a missile released from a helicopter under the influence of downwash from the rotating rotors using AGM-114 Hellfire and UH-60 as the testcase.Simulations are conducted using unsteady Reynolds-averaged Navier-Stokes(URANS)with shear stress transport(SST)k-ωturbulence model,incorporating six-degree-of-freedom(6-DOF)motion and overset grid.Two releasing scenarios,viz.,hover and forward flight,are analyzed under varying missile launch thrust and helicopter forward flight speed.Results reveal that the rotor downwash significantly affects the stability of the missile,particularly during hovering case,where low thrust prolongs wake interaction.In forward flight,the increased airspeed can in principle reduce wake influence but introduces asymmetrical aerodynamic effects on the trajectory of the missile.The findings offer guidance for missile release and launch planning in rotorcraft operations.
基金the funding provided by the National Helicopter Development Project of China。
文摘Accurate measurement of helicopter rotor motion parameters(flap,lead-lag,torsion,and azimuth angles)is essential for rotor blade design,helicopter dynamics modeling,and flight safety and health monitoring.However,the existing methods face challenges in testing equipment installation,calibration,and data transmission,resulting in limited reports on real-time in-flight measurements of blade motion parameters.This paper proposes a non-contact optoelectronic method based on two-dimensional position-sensitive detectors for in-flight measurement and a ground calibration system to obtain real-time rotor motion parameters during helicopter flight.The proposed method establishes the time evolution relationship of rotor motion parameters and verifies the performance of the in-flight measurement system regarding measurement resolution and accuracy through the construction of a blade motion posture experimental platform.The proposed method has been applied to the flight measurement of a medium-sized single-rotor helicopter,and the obtained results have been compared with theoretical analysis outcomes.Furthermore,this paper examines the characteristics of blade motion parameters during flight and discusses the challenges and potential solutions for measuring rotor motion parameters during helicopter flight using the proposed method.
基金support of the National Natural Science Foundation of China(No.12032012)the Priority Academic Program Development of Jiangsu Higher Education Institutions,China。
文摘To achieve high-precision trajectory following during helicopter maneuver tasks and reduce the disruptive influences of unknown variabilities,this study introduces a cascaded-loop helicopter trajectory tracking controller,whose parameters are set using an Ant Colony OptimizationSlime Mould Algorithm(ACO-SMA).Initially,a nonlinear flight dynamics model of the helicopter is constructed.Observer gain functions and nonlinear feedback from a vibrational suppression function to improve the tracking performance of the controller,addressing issues in disturbance estimation and compensation of the Active Disturbance Rejection Control(ADRC).Simultaneously,a cascaded loop system,comprising an internal attitude loop and an external position loop,is created,and the ant colony-slime mold hybrid algorithm optimizes the system parameters of the trajectory tracking controller.Finally,helicopter trajectory tracking simulation experiments are conducted,including spiral ascending and“8”shape climbing maneuvers.The findings indicate that the ADRC employed for helicopter trajectory tracking exhibits outstanding performance in rejecting disturbances caused by gusts and accurately tracking trajectories.The trajectory tracking controller,whose parameters are optimized by the ACO-SMA,shows higher tracking precision compared to the conventional PID and ADRC,thereby substantially improving the precision of maneuver tasks.
基金funded by the National Natural Science Foundation of China(No.52105060)the Special Transmission Project,China(No.KY-1044-2023-0458)。
文摘To reduce the vibration of the Coaxial Helicopter Main Transmission System(CHMTS)considering both level and vertical flight conditions,a vibration evaluation and optimization model for the CHMTS was built.The vibration simulation model of the CHMTS was set up by gear dynamics theory and loaded contact analysis.For better evaluation of the system vibration,a vibration evaluation method for the CHMTS was established by the G1 method-variation coefficient method.A hybrid Gravitational Search Algorithm-Simulated Annealing(GSA-SA)algorithm was combined to balance convergence speed and searching accuracy.The principle test was conducted to prove the accuracy of theoretical method,in which the maximum relative error is16.26%.The optional results show that the vibration of the optimized transmission system decreases significantly,in which the maximum reduction of key vibration indicators reaches more than 20%.The theoretical results have been compared to the experiment to verify the effectiveness of the vibration optimization method.The proposed method could be extended to other fields.
基金the Natural Science Foundation of Sichuan Province,China(Nos.2023NSFSC0902,2024ZYD0083)the Open Project of Key Laboratory of Aerodynamic Noise Control of China Aerodynamics Research and Development Center(No.ANCL20220202)+1 种基金the Open Project of Key Laboratory of Flight Techniques and Flight Safety,CAAC,China(No.FZ2022KF01)the Fundamental Research Funds for the Central Universities,China(No.24CAFUC01007)。
文摘This paper presents a comprehensive test and systematic evaluation analysis of cabin noise in the Robinson R44 RAVEN II helicopter.Initially,microphones were placed within the cabin to conduct systematic assessments of noise levels under various flight conditions,including takeoff,climbing,level flight,landing,hovering,etc.Subsequently,time–frequency analysis was conducted on the test data utilizing traditional A-weighted sound pressure levels,which was followed by quantitative comparisons across different flight conditions.Then,detailed evaluation and discussion were conducted,taking into account the subjective perceptions and communication challenges of cabin crew members.This assessment incorporated the use of aviation noise indicators,speech interference levels,and metrics related to sound quality.Finally,potential noise reduction measures and their effects were preliminarily discussed.The results indicate that helicopter cabin noise exhibited variations across different flight states or positions within the same state,ranging from 87.6 d B(A)to 92.6 dB(A).Discrepancies between A-weighted sound pressure level and psychoacoustic parameters were observed,particularly during hovering states,which indicate that there is a necessity for the combination of multiple evaluation indicators.Notably,damping measure can serve as a pivotal factor in mitigating cabin noise.
基金co-supported by the Foundation of the State Key Laboratory of Aerodynamics(No.RAL202203)the National Key Laboratory Foundation(No.6142202202)the China Postdoctoral Science Foundation(No.2024M754133)。
文摘An advanced Actuator Surface Method(ASM)coupled with Computational Fluid Dynamics(CFD)is developed and applied to the complex unsteady aerodynamic simulation of helicopter.By introducing an improved three-dimensional anisotropic Gaussian kernel,this method effectively addresses the severe aerodynamic load fluctuations commonly associated with traditional Virtual Blade Method(VBM)due to turbulent flow around blade elements.To manage the issues of regional shape and grid cell quantity variations caused by virtual blade sweeping,a universal hybrid grid generation strategy is established without body-fitted and disk interpolation grids,which enhances the computational stability at both blade elements and blade edges.Aerodynamic numerical simulations of helicopter are performed using this method,focusing on rotor/fuselage interaction dominated by rotor wake motion and fuselage blockage effects,Blade-Vortex Interaction(BVI)induced by tip vortices,and maneuvering flights involving collective pitch ramp increases.The results indicate that the advanced ASM demonstrates reliability and robustness in the simulation of complex unsteady flow fields around helicopter.Under similar computational accuracy,the advanced ASM improves computational efficiency by nearly 40 times compared to the oversetgrid-based full Blade-Resolved(B-R)method,and by 6 times compared to the VBM.It shows significant advantages when applied to complex full-aircraft interaction and maneuvering flight conditions that require substantial computational resources.
基金co-supported by the National Natural Science Foundation of China(No.52275061)the Postgraduate Research&Practice Innovation Program of Jiangsu Province,China(No.KYCX24_0562)。
文摘The ballistic impact identification method for the helicopter Tail Drive Shaft System(TDSS)isn't yet comprehensive,which affects helicopter flight safety.This paper proposes a ballistic impact identification method for the TDSS based on vibration response analysis.Based on the Johnson-Cook constitutive model and failure criteria,the ballistic impact finite element simulation model is established,which is verified by the ballistic impact experiment of the Tail Drive Shaft(TDS).Considering the ballistic impact excitation force,the dynamic model of the TDSS with ballistic impact is established,which is verified by finite element commercial software.If a bullet hits the TDS,the bending vibration displacement increases sharply at a certain moment and then significantly increases but remains stable.Meanwhile,the critical speed component appears in the frequency-domain response of bending vibration,and then the speed component significantly increases but remains stable.What's more,the axis trajectory exhibits a sudden,large-scale,and irregular whirling motion at a certain moment,followed by a significant increase but remains stable.Furthermore,if the axial vibration response is small,the bullet core shooting should be considered vertically or at a small incident angle,otherwise,it should be considered at a large incident angle.
文摘The muhi-body analysis of the aeroelastic stability of the tiltrotor aircraft is presented. Muhi-body dynamic differential equations are combined with the equations of the unsteady dynamic inflow model to establish the complete unsteadily aeroelastic coupling analytical model of the tiltrotor. The stability of the tiltrotor in the helicopter mode is analyzed aiming at a semi span soft-inplane tihrotor model with an elastic wing. Parametric effects of the lag stiffness of blades and the flight speed are analyzed. Numerical simulations demonstrate that the multibody analytical model can analyze the aeroelastic stability of the tiltrotor aircraft in the helicopter mode.
文摘An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°
基金National Natural Science Foundation of China(60475039)
文摘The particularity and practicality of harmony operations of close-coupling multiple helicopters indicate that the researches on it are urgent and necessary, Using the model that describes two hovering helicopters carrying one heavy load, an inertia coordinate system and body coordinate systems of each sub-system are established. A nonlinear force model is established too. The equilibrium computation results can be regarded as the reference control inputs of the flight control system under hovering or low-speed flight condition. After the establishment of a translation kinematics model and a posture kinematics model, a coupling dynamics model of the multiple helicopter system is set up. The results can also be regarded as the base to analyze stabilization and design a controller for a close-coupling multiple helicopters harmony operation system.
文摘This article presents a new mathematical model for helicopter comprehensive analysis with the features of flexibility and mathematical simplicity. The model synthesizes the rigid fuselage motion model with 6 degrees of freedom,coupled flap-lag-torsion elastic rotor blade motion model,unsteady aerodynamics model with dynamic stall and high order generalized dynamic wake model. A new blade structural operator with implicit form is formulated,and the components of the blade structure model are independent of each other so that it is convenient to change or handle any component of blade structure without changing the others. What is more,the entire model is developed in a strict state-space form to simplify the comprehensive analysis. Finally,the UH-60 helicopter is taken as an example to predict the blade natural characteristics,the trim characteristics including controls and fuselage attitudes as well as the airloads at blade section under the flight conditions of high speed with moderate thrust and high thrust with moderate speed. The results are compared with UH-60 flight test data and those predicted by two well-known comprehensive codes. The validity of the model presented in this article is verified.
基金Supported by the National Natural Science Foundation of China(61074080)the Innovation Foundation for Aeronautical Science and Technology(08C52001)~~
文摘A direct self-repairing control approach is proposed for helicopter via quantum control techniques and adaptive compensator when some complex faults occur. For a linear varying-parameter helicopter control system, the model reference adaptive control law is designed and an adaptive compensator is used for improving its self-re- pairing capability. To enhance anti-interference capability of helicopter, quantum control feedforward is added be- tween fault and disturbance. Simulation results illustrate the effectiveness and feasibility of the approach.
文摘Gust response alleviation is very important for helicopters which have strong coupling and vibration. Gust disturbance not only influences the ride quality and the precision of the weapon delivery, but also affects to the structural fatigue load and the strength. The method of an optimal control law to suppress the gust disturbance for helicopters is presented. The optimization requires the minimization of the vertical overload at the pilot′s seat, the attitude variation and the control energy consumption under the gust disturbance. Based on the original control system, the new system can be easily realized by adding a vertical speed feedback passage. In order to develop the real-time operational flight control system, the optimized control law is written in C language. The hybrid simulations prove that the performance of gust response alleviation and the efficiency of digitalization are satisfactory.
文摘Aim To extract harmonic frequencies of helicopter acoustic signal as features for hel icopter identification. Methods Estimation of signal parameters via rotational invariance techniques(ESPRIT) was selected to extract harmonic frequencies from really measured helicopter acoustic signal and an algorithm based on the SVD TLS was used. Results ESPRIT correctly extracted harmonic frequencies of helicopter using the data of limited length under the variousflight conditions. Conclusion ESPRIT is an effective method of extracting harmonic frequencies and using harmonic frequencies of helicopter acoustic signal to recognize helicopter is feasible.
文摘Helicopter mathematical model mainly depends on design helicopter control system, flight simulator, and real time control simulation system. But it is difficult to establish a helicopter flight dynamics mathematical model that has features such as rapidness, reliability and precision, because there is no unique and precise expression to some sophisticated phenomenon of helicopter. In this paper a fuzzy helicopter flight model is constructed based on the flight experimental data. The fuzzy model, which is identified by fuzzy inference, has characteristics of computed rapidness and high precision. In order to guarantee the precision of the identified fuzzy model, a new method is adopted to handle the conflict fuzzy rules. Additionally, using fuzzy clustering technology can effectively reduce the number of rules of fuzzy model, namely, the order of the fuzzy model. The simulation results indicate that the method of this paper is effective and feasible.
文摘Adaptive flight control technology, feedback linearization, model inversion theory are reviewed and the error dynamic characteristics are analyzed, and an adaptive on-line neural network attitude control system is presented. The model inversion is under the hover condition. And the adaptive control law based on the neural network is designed to guarantee the boundedness of tracking error and control signals. Simulation results demonstrate that the nonlinear neural network augmented model inversion can self-adapt to the uncertainty and modeling errors of unmanned helicopters. Results are compared while the parameters of PD controller and robustness items are changed.
文摘The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled rigid/elastic blended element based on the fle xible multibody system theory in this paper. It accounts for the effects of prec one, sweep, and the moderately large elastic deflections on the blade and elasti city of shaft and fuselage of the helicopter. The dynamic coupling between the r igid motion of blades about the flap, lag and pitch hinges of articulated rotor and moderately large elastic deflections are included. There is no restriction o n the rotation amplitudes of flap, lag and pitch in the formulation. The stabili ty of periodic solution is studied using the Floquet theory. The transition matr ix is calculated by the Newmark integration method. The aeromechanical stability of a new helicopter is studied. The results show that it is stable in the given forward flight. But the instability arises with the decrease of the bending and torsion stiffness of the shaft.
