Transactions of Nanjing University of Aeronautics & Astronautics
Volume 32,Issue 1,2015 Table of Contents
2015, 32(1):1-8.
DOI: 10.16356/j.1005-1120.2015.01.001
Abstract:
To date unmanned aerial system (UAS) technologies have attracted more and more attention from countries in the world. Unmanned aerial vehicles (UAVs) play an important role in reconnaissance, surveillance, and target tracking within military and civil fields. Here one briefly introduces the development of UAVs, and reviews its various subsystems including autopilot, ground station, mission planning and management subsystem, navigation system and so on. Furthermore, an overview is provided for advanced design methods of UAVs control system, including the linear feedback control, adaptive and nonlinear control, and intelligent control techniques. Finally, the future of UAVs flight control techniques is forecasted.
To date unmanned aerial system (UAS) technologies have attracted more and more attention from countries in the world. Unmanned aerial vehicles (UAVs) play an important role in reconnaissance, surveillance, and target tracking within military and civil fields. Here one briefly introduces the development of UAVs, and reviews its various subsystems including autopilot, ground station, mission planning and management subsystem, navigation system and so on. Furthermore, an overview is provided for advanced design methods of UAVs control system, including the linear feedback control, adaptive and nonlinear control, and intelligent control techniques. Finally, the future of UAVs flight control techniques is forecasted.
2015, 32(1):9-15.
DOI: 10.16356/j.1005-1120.2015.01.009
Abstract:
To satisfy the demand of measuring the velocity of ground moving target through unmanned aerial vehicle (UAV) electro-optical platform, two velocity measurement methods are proposed. Firstly, a velocity measurement method based on target localization is derived, using the position difference between two points with the advantages of easy deployment and realization. Then a mathematical model for measuring target velocity is built and described by 15 variables, i.e. UAV velocity, UAV attitude angular rate, camera direction angular rate and so on. Moreover, the causes of velocity measurement error are analyzed and a formula is derived for calculating the measurement error. Finally, the simulation results show that angular rate error has a strong influence on the velocity measurement accuracy, especially the UAV pitch angular rate error, roll angular rate error and the camera angular altitude rate error, thus indicating the direction for improving velocity measurement precision.
To satisfy the demand of measuring the velocity of ground moving target through unmanned aerial vehicle (UAV) electro-optical platform, two velocity measurement methods are proposed. Firstly, a velocity measurement method based on target localization is derived, using the position difference between two points with the advantages of easy deployment and realization. Then a mathematical model for measuring target velocity is built and described by 15 variables, i.e. UAV velocity, UAV attitude angular rate, camera direction angular rate and so on. Moreover, the causes of velocity measurement error are analyzed and a formula is derived for calculating the measurement error. Finally, the simulation results show that angular rate error has a strong influence on the velocity measurement accuracy, especially the UAV pitch angular rate error, roll angular rate error and the camera angular altitude rate error, thus indicating the direction for improving velocity measurement precision.
2015, 32(1):16-21.
DOI: 10.16356/j.1005-1120.2015.01.016
Abstract:
A novel fuzzy support vector machine based on unbalanced samples (FSVM-US) is proposed to solve the high false positive rate problem since the gyroscope output is susceptible to unmanned aerial vehicle (UAV) airborne electromagnetic environment and the gyroscope abnormal signal sample is rather rare. Firstly, the standard deviation of samples projection to normal vector for SVM classifier hyper plane is analyzed. The imbalance feature expression reflecting the hyper plane shift for the number imbalance between samples and the dispersion imbalance within samples is derived. At the same time, the denoising factor is designed as the exponential decay function based on the Euclidean distance between each sample and the class center. Secondly, the imbalance feature expression and denoising factor are configured into the membership function. Each sample has its own weight denoted the importance to the classifier. Finally, the classification simulation experiments on the gyroscope fault diagnosis system are conducted and FSVM-US is compared with the standard SVM, FSVM, and the four typical class imbalance learning (CIL) methods. The results show that FSVM-US classifier accuracy is 12% higher than that of the standard SVM. Generally, FSVM-US is superior to the four CIL methods in total performance. Moreover, the FSVM-US noise tolerance is also 17% higher than that of the standard SVM.
A novel fuzzy support vector machine based on unbalanced samples (FSVM-US) is proposed to solve the high false positive rate problem since the gyroscope output is susceptible to unmanned aerial vehicle (UAV) airborne electromagnetic environment and the gyroscope abnormal signal sample is rather rare. Firstly, the standard deviation of samples projection to normal vector for SVM classifier hyper plane is analyzed. The imbalance feature expression reflecting the hyper plane shift for the number imbalance between samples and the dispersion imbalance within samples is derived. At the same time, the denoising factor is designed as the exponential decay function based on the Euclidean distance between each sample and the class center. Secondly, the imbalance feature expression and denoising factor are configured into the membership function. Each sample has its own weight denoted the importance to the classifier. Finally, the classification simulation experiments on the gyroscope fault diagnosis system are conducted and FSVM-US is compared with the standard SVM, FSVM, and the four typical class imbalance learning (CIL) methods. The results show that FSVM-US classifier accuracy is 12% higher than that of the standard SVM. Generally, FSVM-US is superior to the four CIL methods in total performance. Moreover, the FSVM-US noise tolerance is also 17% higher than that of the standard SVM.
2015, 32(1):22-28.
DOI: 10.16356/j.1005-1120.2015.01.022
Abstract:
Heavy fuel aviation piston engines (HF-APEs) refer to the engine using fuels with high flash point, such as kerosene or light diesel. Here technique specifications of some classical foreign HF-APE (Hirth3503, Zanzottera 498) are introduced. Recent progress and trend of fuel injection, fuel ignition, working cycle, intake charging, thermal management and electronic control of HF-APE are compared and summarized. Emphases are put on the technological difficulties, solutions and development tendency in the design, retrofitting and manufacturing of HF-APE aiming to provide references for the research of related area and the development of prototype HF-APE in China.
Heavy fuel aviation piston engines (HF-APEs) refer to the engine using fuels with high flash point, such as kerosene or light diesel. Here technique specifications of some classical foreign HF-APE (Hirth3503, Zanzottera 498) are introduced. Recent progress and trend of fuel injection, fuel ignition, working cycle, intake charging, thermal management and electronic control of HF-APE are compared and summarized. Emphases are put on the technological difficulties, solutions and development tendency in the design, retrofitting and manufacturing of HF-APE aiming to provide references for the research of related area and the development of prototype HF-APE in China.
2015, 32(1):29-34.
DOI: 10.16356/j.1005-1120.2015.01.029
Abstract:
Quadrotor unmanned helicopter is a new popular research platform for unmanned aerial vehicle (UAV), thanks to its simple construction, vertical take-off and landing (VTOL) capability. Here a nonlinear intelligent flight control system is developed for quadrotor unmanned helicopter, including trajectory control loop composed of co-controller and state estimator, and attitude control loop composed of brain emotional learning (BEL) intelligent controller. BEL intelligent controller based on mammalian middle brain is characterized as self-leaning capability, model-free and robustness. Simulation results of a small quadrotor unmanned helicopter show that the BEL intelligent controller-based flight control system has faster dynamical responses with higher precision than the traditional controller-based system.
Quadrotor unmanned helicopter is a new popular research platform for unmanned aerial vehicle (UAV), thanks to its simple construction, vertical take-off and landing (VTOL) capability. Here a nonlinear intelligent flight control system is developed for quadrotor unmanned helicopter, including trajectory control loop composed of co-controller and state estimator, and attitude control loop composed of brain emotional learning (BEL) intelligent controller. BEL intelligent controller based on mammalian middle brain is characterized as self-leaning capability, model-free and robustness. Simulation results of a small quadrotor unmanned helicopter show that the BEL intelligent controller-based flight control system has faster dynamical responses with higher precision than the traditional controller-based system.
2015, 32(1):35-41.
DOI: 10.16356/j.1005-1120.2015.01.035
Abstract:
The guidance and control for UAV aerial refueling docking based on dynamic inversion with L1 adaptive augmentation is studied. In order to improve the tracking performance of UAV aerial refueling docking, a guidance algorithm is developed to satisfy the tracking requirement of position and velocity, and it generates the UAV flight control loop commands. In flight control loop, based on the 6-DOF nonlinear model, the angular rate loop and the attitude loop are separated based on time-scale principle and the control law is designed using dynamic inversion. The throttle control is also derived from dynamic inversion method. Moreover, an L1 adaptive augmentation is developed to compensate for the undesirable effects of modeling uncertainty and disturbance. Nonlinear digital simulations are carried out. The results show that the guidance and control system has good tracking performance and robustness at achieving accurate aerial refueling docking.
The guidance and control for UAV aerial refueling docking based on dynamic inversion with L1 adaptive augmentation is studied. In order to improve the tracking performance of UAV aerial refueling docking, a guidance algorithm is developed to satisfy the tracking requirement of position and velocity, and it generates the UAV flight control loop commands. In flight control loop, based on the 6-DOF nonlinear model, the angular rate loop and the attitude loop are separated based on time-scale principle and the control law is designed using dynamic inversion. The throttle control is also derived from dynamic inversion method. Moreover, an L1 adaptive augmentation is developed to compensate for the undesirable effects of modeling uncertainty and disturbance. Nonlinear digital simulations are carried out. The results show that the guidance and control system has good tracking performance and robustness at achieving accurate aerial refueling docking.
2015, 32(1):42-47.
DOI: 10.16356/j.1005-1120.2015.01.042
Abstract:
Aiming at the requirements of accurate target positioning and autonomous capability for adapting to the environmental changes of unmanned aerial vehicle (UAV), a new method for wind estimation and airspeed calibration is proposed. The method is implemented to obtain both wind speed and wind direction based on the information from a GPS receiver, an air data computer and a magnetic compass, combining with the velocity vector triangle relationships among ground speed, wind speed and air speed. Considering the installation error of Pitot tube, cubature Kalman filter (CKF) is applied to determine proportionality calibration coefficient of true airspeed, thus improving the accuracy of wind field information further. The entire autonomous flight simulation is performed in a constant 2-D wind using a digital simulation platform for UAV. Simulation results show that the wind speed and wind direction can be accurately estimated both in straight line and in turning segment during the path tracking by using the proposed method. The measurement accuracies of the wind speed and wind direction are 0.62 m/s and 2.57°, respectively.
Aiming at the requirements of accurate target positioning and autonomous capability for adapting to the environmental changes of unmanned aerial vehicle (UAV), a new method for wind estimation and airspeed calibration is proposed. The method is implemented to obtain both wind speed and wind direction based on the information from a GPS receiver, an air data computer and a magnetic compass, combining with the velocity vector triangle relationships among ground speed, wind speed and air speed. Considering the installation error of Pitot tube, cubature Kalman filter (CKF) is applied to determine proportionality calibration coefficient of true airspeed, thus improving the accuracy of wind field information further. The entire autonomous flight simulation is performed in a constant 2-D wind using a digital simulation platform for UAV. Simulation results show that the wind speed and wind direction can be accurately estimated both in straight line and in turning segment during the path tracking by using the proposed method. The measurement accuracies of the wind speed and wind direction are 0.62 m/s and 2.57°, respectively.
2015, 32(1):48-60.
DOI: 10.16356/j.1005-1120.2015.01.048
Abstract:
To track human across non-overlapping cameras in depression angles for applications such as multi-airplane visual human tracking and urban multi-camera surveillance, an adaptive human tracking method is proposed, focusing on both feature representation and human tracking mechanism. Feature representation describes individual by using both improved local appearance descriptors and statistical geometric parameters. The improved feature descriptors can be extracted quickly and make the human feature more discriminative. Adaptive human tracking mechanism is based on feature representation and it arranges the human image blobs in field of view into matrix. Primary appearance models are created to include the maximum inter-camera appearance information captured from different visual angles. The persons appeared in camera are first filtered by statistical geometric parameters. Then the one among the filtered persons who has the maximum matching scale with the primary models is determined to be the target person. Subsequently, the image blobs of the target person are used to update and generate new primary appearance models for the next camera, thus being robust to visual angle changes. Experimental results prove the excellence of the feature representation and show the good generalization capability of tracking mechanism as well as its robustness to condition variables.
To track human across non-overlapping cameras in depression angles for applications such as multi-airplane visual human tracking and urban multi-camera surveillance, an adaptive human tracking method is proposed, focusing on both feature representation and human tracking mechanism. Feature representation describes individual by using both improved local appearance descriptors and statistical geometric parameters. The improved feature descriptors can be extracted quickly and make the human feature more discriminative. Adaptive human tracking mechanism is based on feature representation and it arranges the human image blobs in field of view into matrix. Primary appearance models are created to include the maximum inter-camera appearance information captured from different visual angles. The persons appeared in camera are first filtered by statistical geometric parameters. Then the one among the filtered persons who has the maximum matching scale with the primary models is determined to be the target person. Subsequently, the image blobs of the target person are used to update and generate new primary appearance models for the next camera, thus being robust to visual angle changes. Experimental results prove the excellence of the feature representation and show the good generalization capability of tracking mechanism as well as its robustness to condition variables.
2015, 32(1):61-69.
DOI: 10.16356/j.1005-1120.2015.01.061
Abstract:
Transonic rudder buzz responses based on the computational fluid dynamics or computational structural dynamics (CFD/CSD) loosely method are analyzed for a tailless flying wing unmanned aerial vehicle (UAV). The Reynolds-averaged Navier-Stokes (RANS) equations and finite element methods based on the detailed aerodynamic and structural model are established, in which the aerodynamic dynamic meshes adopt the unstructured dynamic meshes based on the combination of spring-based smoothing and local remeshing methods, and the lower upper symmetric-Gauss-Seidel (LU-SGS) iteration and Harten-Lax-van Leer-Einfeldt-Wada (HLLEW) space discrete methods based on the shear stress transport (SST) turbulence model are used to calculate the aerodynamic force. The constraints of the rudder motions are fixed at the end of structural model of the flying wing UAV, and the structural geometric nonlinearities are also considered in the flying wing UAV with a high aspect ratio. The interfaces between structural and aerodynamic models are built with an exact match surface where load transferring is performed based on 3D interpolation. The flying wing UAV transonic buzz responses based on the aerodynamic structural coupling method are studied, and the rudder buzz responses and aileron, elevator and flap vibration responses caused by rudder motion are also investigated. The effects of attack, height, rotating angular frequency and Mach number under transonic conditions on the flying wing UAV rudder buzz responses are discussed. The results can be regarded as a reference for the flying wing UAV engineering vibration analysis.
Transonic rudder buzz responses based on the computational fluid dynamics or computational structural dynamics (CFD/CSD) loosely method are analyzed for a tailless flying wing unmanned aerial vehicle (UAV). The Reynolds-averaged Navier-Stokes (RANS) equations and finite element methods based on the detailed aerodynamic and structural model are established, in which the aerodynamic dynamic meshes adopt the unstructured dynamic meshes based on the combination of spring-based smoothing and local remeshing methods, and the lower upper symmetric-Gauss-Seidel (LU-SGS) iteration and Harten-Lax-van Leer-Einfeldt-Wada (HLLEW) space discrete methods based on the shear stress transport (SST) turbulence model are used to calculate the aerodynamic force. The constraints of the rudder motions are fixed at the end of structural model of the flying wing UAV, and the structural geometric nonlinearities are also considered in the flying wing UAV with a high aspect ratio. The interfaces between structural and aerodynamic models are built with an exact match surface where load transferring is performed based on 3D interpolation. The flying wing UAV transonic buzz responses based on the aerodynamic structural coupling method are studied, and the rudder buzz responses and aileron, elevator and flap vibration responses caused by rudder motion are also investigated. The effects of attack, height, rotating angular frequency and Mach number under transonic conditions on the flying wing UAV rudder buzz responses are discussed. The results can be regarded as a reference for the flying wing UAV engineering vibration analysis.
2015, 32(1):70-80.
DOI: 10.16356/j.1005-1120.2015.01.070
Abstract:
Since the subsystems of aerodynamics, propulsion, structure and so on in hypersonic vehicles involve characteristics of nonlinearity, strong coupling and uncertainty, and typical hypersonic vehicles adopt slender-body and wave-rider layout with widely-used lightweight materials, the accuracy of the modeling with a conventional rigid-body assumption is challenged. Therefore, a nonlinear mathematical longitudinal model of a hypersonic vehicle is established with its geometry provided to estimate aerodynamic force and thrust using hypersonic aerodynamics and quasi-one-dimensional flow with heat added and capture vehicle aeroelasticity using a single free-free Bernoulli-Euler beam model. Then the static and dynamic properties of the rigid and rigid-aeroelasticity coupling model are compared via theoretical analysis and numerical simulation under the given flight condition. Finally, a LQR controller for rigid model is designed and the comparable results are obtained to explain the aerolasticity influence on the control effect. The simulation results show that the aeroelasticity mode of slender-body hypersonic vehicles affects short period mode significantly and it cannot be simply neglected.
Since the subsystems of aerodynamics, propulsion, structure and so on in hypersonic vehicles involve characteristics of nonlinearity, strong coupling and uncertainty, and typical hypersonic vehicles adopt slender-body and wave-rider layout with widely-used lightweight materials, the accuracy of the modeling with a conventional rigid-body assumption is challenged. Therefore, a nonlinear mathematical longitudinal model of a hypersonic vehicle is established with its geometry provided to estimate aerodynamic force and thrust using hypersonic aerodynamics and quasi-one-dimensional flow with heat added and capture vehicle aeroelasticity using a single free-free Bernoulli-Euler beam model. Then the static and dynamic properties of the rigid and rigid-aeroelasticity coupling model are compared via theoretical analysis and numerical simulation under the given flight condition. Finally, a LQR controller for rigid model is designed and the comparable results are obtained to explain the aerolasticity influence on the control effect. The simulation results show that the aeroelasticity mode of slender-body hypersonic vehicles affects short period mode significantly and it cannot be simply neglected.
2015, 32(1):81-88.
DOI: 10.16356/j.1005-1120.2015.01.081
Abstract:
With the liquid propellant making up 60%—70% of the takeoff weight of the hypersonic vehicle, the dynamic load caused by great propellant sloshing interacts with the flexible structure of the aircraft. Therefore, the dynamic model displays characteristics of strong coupling with structure/control and nonlinearity. Based on the sloshing mass dynamic simplified as a spring-mass-damping model, a rigid-flexible-sloshing model is constructed. Moreover, the effect on the dynamic performance of the coupled model is analyzed with changing frequency and damping. The results show that propellant sloshing dynamics significantly affects the rigid body motion modes, especially flexible mode and short mode. The right half plane pole (RHP) moves far from the imaginary axis with the consumption of the propellant. The flexible mode attenuates with the increase of the sloshing damping, and the coupling becomes stronger when sloshing frequency is close to the short mode frequency or the flexible frequency of the beam.
With the liquid propellant making up 60%—70% of the takeoff weight of the hypersonic vehicle, the dynamic load caused by great propellant sloshing interacts with the flexible structure of the aircraft. Therefore, the dynamic model displays characteristics of strong coupling with structure/control and nonlinearity. Based on the sloshing mass dynamic simplified as a spring-mass-damping model, a rigid-flexible-sloshing model is constructed. Moreover, the effect on the dynamic performance of the coupled model is analyzed with changing frequency and damping. The results show that propellant sloshing dynamics significantly affects the rigid body motion modes, especially flexible mode and short mode. The right half plane pole (RHP) moves far from the imaginary axis with the consumption of the propellant. The flexible mode attenuates with the increase of the sloshing damping, and the coupling becomes stronger when sloshing frequency is close to the short mode frequency or the flexible frequency of the beam.
2015, 32(1):89-96.
DOI: 10.16356/j.1005-1120.2015.01.089
Abstract:
The unsteady aerodynamic loads generated by the thin-shell object separating from aircraft affect flying safety. To investigate the loads, a method combining numerical simulation and experiment is proposed. Firstly, the motional tendency of the thin-shell object separating from aircraft is calculated, and then the high-speed air blowing test on ground is designed. Thereafter, the external store is employed to avoid colliding with the thin-shell object in air. Finally, the hanging and flight test is conducted by a high-speed unmanned aerial vehicle (UAV), and the feasibility of the thin-shell object separating from aircraft at high speed is proved. Consequently, the separating problem of a thin-shell object with an unconventional aerodynamic configuration is solved, and the collision with aircraft is prevented.
The unsteady aerodynamic loads generated by the thin-shell object separating from aircraft affect flying safety. To investigate the loads, a method combining numerical simulation and experiment is proposed. Firstly, the motional tendency of the thin-shell object separating from aircraft is calculated, and then the high-speed air blowing test on ground is designed. Thereafter, the external store is employed to avoid colliding with the thin-shell object in air. Finally, the hanging and flight test is conducted by a high-speed unmanned aerial vehicle (UAV), and the feasibility of the thin-shell object separating from aircraft at high speed is proved. Consequently, the separating problem of a thin-shell object with an unconventional aerodynamic configuration is solved, and the collision with aircraft is prevented.
2015, 32(1):97-109.
DOI: 10.16356/j.1005-1120.2015.01.097
Abstract:
Tradeoff analysis of the factors, including external environment and unmanned aerial vehicle (UAV) aerodynamic attributes, which affect longitudinal carrier landing performance, is important for small UAV. First, small UAV longitudinal carrier landing system is established, as well as the nonlinear dynamics and kinematics model, and then the longitudinal flight control system using backstepping technology with minimum information about the aerodynamic is designed. To assess the landing performance, a variety of influencing factors are considered, resulting in the constraints of aerodynamic attributes of carrier UAV. The simulation results show that the severe sea condition has the greatest influence on landing dispersion, while air wake is the primary factor on impact velocity. Among the longitudinal aerodynamic parameters, the lift curve slope is the most important factor affecting the landing performance, and increasing lift curve slope can improve the landing performance significantly. A better system performance will be achieved when the lift curve slope is larger than 2 per radian.
Tradeoff analysis of the factors, including external environment and unmanned aerial vehicle (UAV) aerodynamic attributes, which affect longitudinal carrier landing performance, is important for small UAV. First, small UAV longitudinal carrier landing system is established, as well as the nonlinear dynamics and kinematics model, and then the longitudinal flight control system using backstepping technology with minimum information about the aerodynamic is designed. To assess the landing performance, a variety of influencing factors are considered, resulting in the constraints of aerodynamic attributes of carrier UAV. The simulation results show that the severe sea condition has the greatest influence on landing dispersion, while air wake is the primary factor on impact velocity. Among the longitudinal aerodynamic parameters, the lift curve slope is the most important factor affecting the landing performance, and increasing lift curve slope can improve the landing performance significantly. A better system performance will be achieved when the lift curve slope is larger than 2 per radian.
2015, 32(1):110-114.
DOI: 10.16356/j.1005-1120.2015.01.110
Abstract:
To solve the flight control problem for unmanned hypersonic vehicles, a novel intelligent optimized control method is proposed. A flight control system based on integral separated proportional-integral-derivative (PID) control is designed for hypersonic vehicle, and an improved shuffled frog leaping algorithm is presented to optimize the control parameters. A nonlinear model of hypersonic vehicle is established to examine the dynamic characteristics achieved by the flight control system. Simulation results demonstrate that the proposed optimized controller can effectively achieve better flight control performance than the traditional controller.
To solve the flight control problem for unmanned hypersonic vehicles, a novel intelligent optimized control method is proposed. A flight control system based on integral separated proportional-integral-derivative (PID) control is designed for hypersonic vehicle, and an improved shuffled frog leaping algorithm is presented to optimize the control parameters. A nonlinear model of hypersonic vehicle is established to examine the dynamic characteristics achieved by the flight control system. Simulation results demonstrate that the proposed optimized controller can effectively achieve better flight control performance than the traditional controller.
2015, 32(1):115-120.
DOI: 10.16356/j.1005-1120.2015.01.115
Abstract:
The Cramer-Rao bound (CRB) for two-dimensional (2-D) direction of arrival （DOA） estimation in multiple-input multiple-output (MIMO) radar with uniform circular array (UCA) is studied. Compared with the uniform linear array (ULA), UCA can obtain the similar performance with fewer antennas and can achieve DOA estimation in the range of 360°. This paper investigates the signal model of the MIMO radar with UCA and 2-D DOA estimation with the multiple signal classification (MUSIC) method. The CRB expressions are derived for DOA estimation and the relationship between the CRB and several parameters of the MIMO radar system is discussed. The simulation results show that more antennas and larger radius of the UCA leads to lower CRB and more accurate DOA estimation performance for the monostatic MIMO radar. Also the interference during the 2-D DOA estimation will be well restrained when the number of the transmitting antennas is different from that of the receiving antennas.
The Cramer-Rao bound (CRB) for two-dimensional (2-D) direction of arrival （DOA） estimation in multiple-input multiple-output (MIMO) radar with uniform circular array (UCA) is studied. Compared with the uniform linear array (ULA), UCA can obtain the similar performance with fewer antennas and can achieve DOA estimation in the range of 360°. This paper investigates the signal model of the MIMO radar with UCA and 2-D DOA estimation with the multiple signal classification (MUSIC) method. The CRB expressions are derived for DOA estimation and the relationship between the CRB and several parameters of the MIMO radar system is discussed. The simulation results show that more antennas and larger radius of the UCA leads to lower CRB and more accurate DOA estimation performance for the monostatic MIMO radar. Also the interference during the 2-D DOA estimation will be well restrained when the number of the transmitting antennas is different from that of the receiving antennas.
2015, 32(1):121-127.
DOI: 10.16356/j.1005-1120.2015.01.121
Abstract:
Aircraft laminated composite components often suffer a variety of high velocity impacts with large quantity of energy, which usually affects aircraft behavior and would incur component damages, even disastrous consequences. Therefore, one investigates the impact resistance of a new type of composite material, Ti/CFRP/Ti sandwich structure, and launches impact tests by using an air gun test system. Then one acquires the critical breakthrough rate of the structure and analyzes the damages. The results show that the main failure mode of the front titanium sheet is shear plugging and brittle fracture of adhesive layer with fiber breakage, while the back titanium sheet is severely ripped. The rear damage is worse than the front one. Compared with traditional CFRP laminates, the critical breakthrough rate of Ti/CFRP/Ti sandwich structure is improved by 69.9% when suffered the impact of a bearing ball with 2 mm radius.
Aircraft laminated composite components often suffer a variety of high velocity impacts with large quantity of energy, which usually affects aircraft behavior and would incur component damages, even disastrous consequences. Therefore, one investigates the impact resistance of a new type of composite material, Ti/CFRP/Ti sandwich structure, and launches impact tests by using an air gun test system. Then one acquires the critical breakthrough rate of the structure and analyzes the damages. The results show that the main failure mode of the front titanium sheet is shear plugging and brittle fracture of adhesive layer with fiber breakage, while the back titanium sheet is severely ripped. The rear damage is worse than the front one. Compared with traditional CFRP laminates, the critical breakthrough rate of Ti/CFRP/Ti sandwich structure is improved by 69.9% when suffered the impact of a bearing ball with 2 mm radius.
2015, 32(1):128-132.
DOI: 10.16356/j.1005-1120.2015.01.128
Abstract:
The whole airspace phased array telemetry, track and command (TT&C) system is regarded as the development tendency of next generation TT&C system, and the distribution of the antenna units and the beamforming technology have sparked wide interest in this field. A method for antenna distribution is proposed based on the linear subarrays technology. A symmetrical truncated cone conformal array is composed of the linear subarrays placed on the generatrix. The impact of truncated cone bottom radius and elevation angle on beamforming are studied and simulated. Simulation results verify the system design.
The whole airspace phased array telemetry, track and command (TT&C) system is regarded as the development tendency of next generation TT&C system, and the distribution of the antenna units and the beamforming technology have sparked wide interest in this field. A method for antenna distribution is proposed based on the linear subarrays technology. A symmetrical truncated cone conformal array is composed of the linear subarrays placed on the generatrix. The impact of truncated cone bottom radius and elevation angle on beamforming are studied and simulated. Simulation results verify the system design.
2015, 32(1):133-136.
DOI: 10.16356/j.1005-1120.2015.01.133
Abstract:
A rapid method of the trim drag prediction for the blended-wing-body unmanned aerial vehicle (UAV) configuration is proposed. The method consists of four steps. The first step is to parameterizedly model the blended-wing-body UAV configuration; the second is to analyze the aerodynamics of the geometric model; the third is to create aerodynamic surrogate model; and the final step is to predict the trim drag using the surrogate model. Hence, a tool for trim drag prediction is developed by integration of the four steps. The impacts of the allocation of control surfaces, position of gravity center and planform parameters on the trim drag are investigated by using the tool. Results show that using the control surface in outer wing for trim has an advantage of lower trim drag, and the position of gravity center has a primary impact on the trim drag. Moreover, the planform has secondary impacts on the trim drag.
A rapid method of the trim drag prediction for the blended-wing-body unmanned aerial vehicle (UAV) configuration is proposed. The method consists of four steps. The first step is to parameterizedly model the blended-wing-body UAV configuration; the second is to analyze the aerodynamics of the geometric model; the third is to create aerodynamic surrogate model; and the final step is to predict the trim drag using the surrogate model. Hence, a tool for trim drag prediction is developed by integration of the four steps. The impacts of the allocation of control surfaces, position of gravity center and planform parameters on the trim drag are investigated by using the tool. Results show that using the control surface in outer wing for trim has an advantage of lower trim drag, and the position of gravity center has a primary impact on the trim drag. Moreover, the planform has secondary impacts on the trim drag.
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