Abstract：Considering the problem of stability analysis of the satellite attitude system controlled by a phaseplane controller, a simplified phaseplane control method is proposed, and then the stability of the rigid satellite attitude system under the proposed controller is investigated in detail. By estimating the closedloop system’s trajectories with the phaseplane analysis method, the existence of a stable region of the closedloop system is proved, which satisfies certain properties. The formula to determine the stable region is derived in terms of control parameters. Under certain conditions, it is proved that the trajectories can reach the stable region in finite time from any initial value. Therefore, the sufficient condition is obtained to guarantee the globally uniformly ultimate boundedness of the closedloop system. In addition, the formula to estimate the reach time is provided. Simulations are provided to illustrate the effectiveness of proposed formulas and methods.  

Abstract：The problems of spinning angle velocity determination for satellites with attitude anomaly are investigated. A spinning angle velocity determination method based on sun sensors measurement is proposed, and the influence of individual error source on the angle rate determination error is obtained by deliberate analysis. According to the error analysis conclusion, a strategy is presented to maintain the determination precision for a smaller spinning angle rate by using different measurement data with larger time interval in spinning angle velocity calculation. The effectiveness of the proposed method is demonstrated by numerical simulations, and by applying the method with telemetry data of sun sensors the actual spinning angle velocity is also obtained for a satellite inorbit out of control．  

Abstract：A new angular momentum management method for variable speed control moment gyroscope (VSCMGs) is proposed in this paper. The control torque (CMGs) mode and the flywheel mode from the VSCMGs is seperated, and the singular value decomposition on the CMGs mode torque output is performed. Only when the CMGs mode is close to singular， the command torque in the singular direction be distributed to RWs mode, so as to make full use of VSCMGs execution ability while ensuring VSCMGs output accuracy. At the same time, zero motion is also introduced into the angular momentum management algorithm to achieve rotor wheel speed balance and frame structure avoidance singularity, and the nonlinearity of the deadzone of the frame angular velocity and the compensation of neglected items are considered. Simulation results show that the designed control law can achieve highprecision hightorque output when maneuvering largeangle and highspeed spacecraft.  

Abstract：Considering large projectile envelope and complex flight environment, the model exists high uncertainty, external wind and internal disturbance, which increase difficulties of rolling attitude control. Furthermore, axisymmetric miniature missile with small aspect ratio structural characteristics lead to weakness in quality of rolling control based on classical control theory, especially in robustness and antidisturbance ability. Thus, the reducedorder ESO proposed estimates compound disturbance and then realtimely compensate for inputs. Then, the LQR stability design method with the integral term is adopted for controller designing on the basis of simplified nominal model. Last, the essay proves the stability of closedloop system. And the simulation results show that this combined control system designed has more excellent characteristics than PID controller.  

Abstract：A robust entry guidance law with finite time convergence is designed for dragtracking in this paper. The bank angle is regarded as the control variable. First, a robust sliding mode control (SMC) method, which can guarantee the finite time convergence, is proposed for a class of nonlinear systems with uncertainties and disturbances. Then, after the drag dynamics is given, the finite time guidance law is designed based on the proposed SMC method. A 1 000run Monte Carlo simulation shows the effectiveness of the presented approach.  

Abstract：Based on the Hill equation， a design method of doubleimpulsive flyaround trajectory is given on the basis of rectangular keep out zone. With dynamics model transformed, the relation of the variables of the model is deduced. On this basis, the theorems are proved, and then safety restrictions are given. Furthermore, we can get runofmill design methods of safe flyaround trajectories between Vbar direction and arbitrary directions in orbital plane. Lastly, this method is showed to be feasible by simulations about two kinds of typical flyaround trajectories, as well as simulation about flyaround trajectory between Vbar direction and arbitrary direction.  

Abstract：To realize the high accuracy rendezvous and docking, the 6degree orbit and attitude control is needed. For the same control command sequence, different thruster configurations and different control allocation algorithms lead to different fuel consumptions. As the control command sequence can be divided by time and space region, the regional average fuel consumption is proposed. The computing method of the regional average fuel consumption is given for both the 3degree and the 6degree control tasks. The optimal design method of the thruster configuration is suggested, which has important significance for thruster configuration design and control allocation.  

Abstract：The problems of the quadruped robot joint control and turning gait planning are researched in this paper. Firstly, the singleleg dynamic model of quadruped robot is established. Its multiinput multioutput (MIMO), timevarying, coupled and nonlinear characteristics are analyzed. A MIMO characteristic model with input decoupling is derived. A controller, which combines MIMO goldensection adaptive control law and differential control law, is designed for the model. Secondly, a turning gait is proposed via modifying the step lengths of the left and right legs respectively. Furthermore, the relationship between the step length correction and the turning radius is analyzed. Finally, the virtual prototype technology is used to simulate the straight and turning gaits of the robot under zero load and 50kg load respectively. The simulation results verify the effectiveness of the proposed method.  

Abstract：In view of the shortcomings of the existing groundbased deep space communication system, such as insufficient communication distance and insufficient data rate in deep space, a spacebased active communication and space radio telescope function system based on distributed cooperative control method is proposed. The system is deployed in GEO, consisting of a feed source satellite, a central satellite and 4294 unit satellites, through a precise formation and onorbit assembly. The system integrates the functions of spacebased active communication and radio astronomy telescope. The system can provide a deep space station for deep space detector, due to deploying in GEO, no atmosphere and cloud, which can greatly improve the communication distance and data rate of deep space exploration. What’s more, resolution and sensitivity of the radio telescope can upgrade 12 orders.  

Abstract：Variation of temperature has a strong effect on Hemispherical Resonator Gyro（HRG）output. In this paper, we propose a selfcompensation scheme based on the drivemode resonant frequency. The analysis of relationship between resonant frequency and temperature is used to illustrate the feasibility of frequency used as temperature. The bias temperature drift mode is established and the drivemode resonant frequency is used to compensate the bias drift real timely. The resolution of the resonant frequency is 0.03 Hz which means a resolution of 0.075 ℃ for temperature. The selfcompensation yielded a total bias error of 2.8（°）/h over temperature range of -10 ℃~60 ℃, corresponding to a 30（°）/h bias error before compensation. These results show that the scheme is valid.  

Abstract：Thermal Propellant Gauging System (PGS) method is a kind of new method to measure the amount of remaining propellant It is convenient to carry out and has high measurement precision, which is suitable for the measurement at EndOfLife (EOL). A large number of ground tests on a type of tank are needed to verify the feasibility and accuracy of PGS and give suggestion for its application onorbit. The data are collected through the tests of different filling quantity of test liquid under different heating conditions. According to the data analysis , the measurement precision is obtained . The good performance of PGS is showed. And the suggestion of application is presented. The profile is of reference value in the application of PGS onorbit．