Visualization simulation of satellite orbiting the earth can display the satellite’s orbit and attitude motion intuitionisticly and really. The STK is an important tool for satellite simulation and analysis. It can display 2D and 3D graphics of the satellite. This paper studies the core technique of STK/Connect, makes use of the windows sockets and multi-threads programming technology, and develops the drive program for STK realtime visualization simulation. This drive program receives the data from the telemetry simulator computer according to the UDP protocal, then drives the STK runing through communicatting with the STK’s connect module. Results show that the drive program can realize the realtime simulation for satellites.

The SRAM-based FPGA is one of the on-board electrical devices susceptible to radiation. In addition, effects and errors induced by single event effect on SRAM-based FPGA are distinct from others because of its special structure and operation. This paper analyzes in detail diverse fault modes for SRAM-based FPGA in space applications, and investigates broadly the corresponding fault-tolerant methods by taking the mainstream device in this type of FPGA as objects. Results show that we are able to effectively decrease the possibilities of the radiation-induced faults of SRAM-based FPGA by taking appropriate fault tolerant methods.

There is error of the traditional accelerometer self-calibration method in the presence of an accelerometer bias and it is not easy to validate correctness of calibrated parameters. Based on the contraction mapping theorem, an iterative calibration method for the accelerometer and a simple effective method validating calibrated parameters are brought forward. All these are proved in this paper.

A new approach for nonlinear filtering systems is presented in this paper to meet the requirements both on accuracy and computing time. This method, called as approximate second-order extended Kalman filter (AS-EKF), is based on the frame of recursive linear minimum variance estimation. Other than extended Kalman filter (EKF) linearizing all nonlinear models, the new approach estimates the expectation value to the second-order of accuracy. We show that this technique is more accurate than EKF, and it also costs the less computing time than unscented Kalman filter (UKF).It can be used in the nonlinear filtering concerned with both accuracy and computing time.

A method for designing attitude stabilization controllers is proposed in this paper for flexible spacecraft without angular velocity measurements. In order to solve the attitude control problem when the angular velocity is not available due to gyro failure or gyroless configuration, a velocity-free attitude regulation controller is proposed in this paper based on quaternion with the global asymptotic stability of the closed loop system proved by combining the Lyapunov direct method with the LaSalle’s theorem. An improved controller is also designed in this paper. Numerical simulations demonstrate the effectiveness of the proposed method.

For the guidance law design problem of lunar spacecraft reentry at the second cosmic velocity, a predictor-corrector guidance method is adopted. The flight characteristic of spacecraft at the balanced attack angel is analyzed, and then the three degrees of freedom motion equation for reentry is established. Furthermore, the theory of predictor-corrector and its longitudinal and lateral guidance laws are introduced in detail. Reentry simulation in such two conditions, one in standard initial condition and other in error initial condition shows that the predictor-corrector can not only acquire higher accuracy, but has strong robustness to the initial errors at all constraints to be satistied, even at the uncertain conditions.

The structures of various motion simulators for spacecraft’s rendezvous simulation test are discussed with the emphasis on simulators for RVD proximity operation phase in this paper. A new 9-DOF simulator concept of 3+6 type with a cross beam on ground is presented. And then a new 9-DOF simulator concept of 4+5 type is presented. Finally a new concept of simple pose simulator is proposed and this is the only simulator with wider range.

A standard autonomous astronomical navigation method is based on the information of the earth sensor and the star sensor. The extended Kalman filter (EKF) is implemented to estimate the position vector of the spacecraft according to the spacecraft dynamics model and the measurements from these sensors. In order to decrease the estimation error of the filter during orbit maneuvers of the spacecraft, an adaptive robust extended Kalman filter (AREKF) is designed for spacecraft autonomous navigation. The simulation results show that the AREKF can effectively depress the unfavorable effect of thrust uncertainties, and the navigation performance is improved effectively without additional sensors. The estimate of the AREKF is more accurate than ones of the EKF and the adaptive extended Kalman filter (AEKF).

The attitude control problem of the assembled spacecraft in rendezvous and docking tasks is studied in this paper. Control capability of thrusters whose outputs are bounded is analyzed firstly. Then the controller is designed by using intelligent adaptive control method based on characteristic model. The same controller is applied to the one-shaped and L-shaped assembled spacecraft, and the simulation results are given respectively. In contrast with the robust controller, this method is feasible and of much advantages.

Research involved in this paper focuses on attitude dynamics modeling and control of a spin-stabilized micro-satellite for near-earth space exploration. There are some special requirements for attitude control due to the scientific mission. During modeling, the effect of double-side booms-extended disturbance on the satellite attitude has been considered in detail. Using the nutation characteristics of the spin-stabilized satellite, a controller of attitude control combined with nutation control has been designed. The firing time is determined based on the phase of satellite transverse angular velocity and the orientation of jet torque in the inertial space. Following control strategy has been applied: firstly coarse nutation control and precession control, then fine nutation control. When the satellite nutation angle is large due to the long term action of disturbance torques and orbit maneuver is needed, this controller can be used to adjust the orientation of spin axis conveniently.

A composite attitude controller for the base and the manipulator is designed for a free-flying space robot via the backstepping method. The controller ,which feeds back the position and orientation of the manipulator tip in workspace, joint angle and speed in joint space, attitude and velocity of the base, can fulfill the task in workspace directly, avoiding the motion planning from workspace to joint space and complicated differentiating of Jacobian matrix. Meanwhile, attitude control of the base can extend the capacity of space robot and improved the character of dynamical singularity. Finally, the numerical simulation of a free-flying space robot validates the controller.

This paper first establishes a satellite orbit dynamics model perturbed by the Earth’s non-spherical gravity, and then develops a numerical orbit propagator programmed by C++ language, with the Runge-Kutta-Fehlberg 7(8) integrator. In this modeling process, the precession, nutation and some other astronomical quantities are considered through two standards released by U.S. Naval Observatory in 1981 and 2005, respectively. By comparing the simulation results with that of the professional software STK, it shows that the magnitudes of position vector differences between these two standards and STK are both not more than the decimeter level, and that of velocity vector differences are both not more than the level of millimeter per second. Therefore, the model and orbit propagator are both prove to be valid, and the difference of orbit propagation caused by referencing the old and new standards is slight.

The safety of angular contact ball bearing in CMG under random vibration is investigated. The axial stiffness and radial stiffness of the bearing under both axial force and radial force are derived firstly, then an equivalent linear spring model is established so that the axial force and radial force of the bearing under random vibration can be iteratively obtained using the spectrum analysis of ANSYS. The safety is valuated by comparing radial equivalent static load and axial load with radial base rated static load and limit axial load separately. An analysis of a type of angular contact ball bearing under given input power spectrum density is performed. The result demonstrate that the bearing is safe in any load direction.

Based on varieties of the inertia matrix for a combined spacecraft separated, this paper proposes a adaptive control algorithm. Attitude quaternion is adopted to establish the model of the kind of spacecraft, and based on the Lyapunov stability theory, the algorithm is designed by the backstepping control method. Theoretical analysis proves the algorithm be stable, and the simulation results confirm the effectiveness of the algorithm.

The qualitative model-based fault diagnosis approach can depict the system configuration in qualitative model, and get the predicted behavior qualitatively. The system fault can be diagnosed by determining the differences between the model of an artifact and the artifact itself. This paper investigates the diagnosis mechanism, system modeling, advantages and disadvantages of the qualitative model-based fault diagnosis approach. Its progress, application, and trend developing are also addressed.

Sensors are important components in a satellite navigation system, so both fault detection and identification for them are of great significance for the reliability improvement of system. In this paper, a sensor fault identification method based on fuzzy logic for the satellite navigation system is presented. First of all, the satellite navigation system is described by the T-S fuzzy model, and then full de-coupled parity equations are applied to detect faults, and a Kalman filter is used to identify fault parameters. Simulation results show that, in case of multiple sensor faults in the navigation system this method can effectively detect them and accurately identify their parameters.

The measurement accuracy of Sun Sensor is mainly influenced by earth albedo. In this paper, a model about energy attenuation for sunlight incident upon CCD Sun Sensor in large angle is proposed. Through computing peak value of earth albedo irradiance, the influence is correlated with sunlight’s incident angle in theory. For instance, the irradiance produced by earth albedo is comparable to that of the Sun, when the Sun appears in the edge of field of view of CCD Sun Sensor and earth albedo appears in the worst condition. Finally a method about the improvement of structural parameter of Sun Sensor to decrease the influence of earth albedo on Sun Sensor is presented.

Stellar instrument magnitudes are important for star trackers; however, there are not simple and accurate methods to estimate stellar instrument magnitudes. In this paper, a method using color indexes to compute instrument magnitudes is proposed. On the analysis of errors, better data is got by using the improved method. Because the errors are not greater than 0.1 magnitude, the data satisfy requirements of the star sensors.

The paper surveys the attitude tracking of aircrafts via Fliess functional expansion. First, one step and ¯nite step predictive controls are considered. Then we show that when the system decoupling matrix is of full row rank, using a proper state feedback, the closed-loop system has a Fliess functional expansion with only ¯nite terms. This approach is particularly suitable for the online design of attitude control of spacecraft.

The time-varying parameter estimation is a challenging problem in strapdown inertial navigation system (SINS)/synthetic aperture radar (SAR) integrated navigation systems because SAR measurements are fairly few and have unequal interval. Based on the orthogonality principle, an adaptive recursive least square (RLS) algorithm of SINS/SAR integrated navigation systems is proposed. The adaptive RLS can estimate the time -varying parameters by constructing adaptive fading factors. The adaptive RLS algorithm is deduced and its performance is rigorously analyzed. It is proved that the parameter estimation error of the algorithm is exponentially uniformly bounded. Simulation results demonstrate the feasibility and effectiveness of the proposed approach.

In order to guarantee high-accuracy navigation independent of GPS/GLONASS during long time and long range running, a combined land navigation system using the odometer(OD) and geographic information system(GIS) assisted with the laser-gyro strapdown inertial navigation system(LSINS/OD/GIS) is described in this article. Based on error estimation equation, an automatic zero updating function and odometer parameter calibration and the navigation azimuth error correction by GIS position information synthetically are proposed, using the GIS position information and possible temporary parking states. The results of accuracy verification test are also presented, indicating divergent velocity of main errors has been effectively reduced.

The high frequency jitter resulting from high speed rotor of actuators in its rotating process has a serious effect on the attitude control accuracy and stability. In this paper the whirling characteristics of high speed rotors are analyzed by means of building the dynamics model of high speed rotors in flywheel and control moment gyro (CMG). An experiment is also designed to test related results of the theoretical analysis.

Based on the relative motion dynamic model described by C-W equations, and by considering parameter uncertainties characterized by the norm-bounded method and the control input constraints in practice, a robust orbit design approach is proposed for the spacecraft rendezvous engineering. By introducing a Lyapunov function, the design problem is cast into a convex optimization problem subject to linear matrix inequalities. By solving this optimization problem, the robust controllers satisfying the requirements can be designed.

This paper proposes a fault-tolerant reconfigurable controller for attitude control system of spacecrafts with abrupt changes of centroid. A novel interacting multiple model algorithm, based on the Unscented Kalman filter and a performance index using output residuals, is presented to diagnose the faults. Once a fault is detected and isolated, controller parameters can be reconfigured to compensate the loss of performance and keep the good attitude tracking accuracy of the spacecraft. Numerical simulations validate the effectiveness of the proposed method.

Rendezvous between two spacecraft is a very important spaceflight mission . Based on the C-W equation, equations with both rectangular coordinates and orbital elements for the impulse maneuvers in long-distance guidance are derived. The iterative processes for solving equations are given. As the maneuvers are applied, the variables are not enough to satisfy all the final-time constraints. This paper gives the partial variable aiming method to satisfy some of the final-time constraints. Numerical simulations show that partial variable aiming method can aim concerned variables.

Abstract：Based on the vacuum plume effect test system of some SinoRussian cooperation project built in Beihang University, and the mission requirements of the space station plume effects experimental research, this paper successfully conducts the aerodynamic force experiments of 10N bellshaped thruster plume by impacting on the tablet. The vacuum chamber, the vacuum pumping system, working fluid supply system, data acquisition system, 10N bellshaped thrusters and aerodynamic measurement devices are introduced. To be closer to the engineering practice, the aerodynamic experiments are equipped with real bell thruster, and the associated aerodynamic experimental scheme is given. The pressure distributions of plate surface with three different flatpanel installation locations are presented. This work provides insight into the future study of the plume aerodynamic force.

Nuclear electric propulsion (NEP) is the key propulsion technology for future deep space exploration. The development of nuclear electric propulsion system, nuclear fission reactor, and high power electric propulsion are introduced based on the investigation of nuclear electric propulsion space application. Then the achievements and the key technologies of nuclear electric propulsion system are summarized. Finally several suggestions are proposed for China’s nuclear electric propulsion technology development.

Sun tracking and positioning technique is widely used in areas such as energy, meteorology and space. Related methods is classified systematically and presented comprehensively in this paper. Some newer applications such as porous multiplexing and optical fiber introduced etc, in this field are also analyzed. Finally the “bottleneck” problem in this domain is concluded, and the future trend is looked to.

The paper describes the space operations, gives brief remarks about the design of the current system, and states the future vision. We propose four fundamental abilities of the control system via analysis of typical tasks, including autonomousflying management, noncooperative target reconstruction and recognition, quick orbital transfer and agile pointing, and space complex programming control.

In order to meet the requirement for time performance of system communication, a new design method of lower machine communication software is proposed by using transmitting interrupt and receiving interrupt. When GPS added as another host causes a concurrent communication of CAN bus，the reason of interrupt conflict is analyzed and the algorithm is updated to solve the bus collision.

Space is the new frontier of a country. Space activities are important embodiment of national will and strategic intention, and important guarantee for the expansion of national interests. Space security has become an important part of national security. Space situation awareness refers to the acquisition and cognition of space situation information, including space target monitoring and space environment monitoring. It is the basis for further space control and space confrontation. Firstly, this paper reviews the development of relevant regulations in the field of space situational awareness, introduces the implementation of several typical situation awareness projects in the field of high orbit in the United States, and summarizes four key technologies, including space entry, autonomous operation, rendezvous and docking navigation and control, and multi angle stereo imaging technology. Finally, this paper gives some suggestions on the establishment of situation awareness system, spacebased autonomous sensing system, and the development of space attack and defense confrontation capability.

A gimbal drive assembly (GDA) for intersatellite link antenna (ILA) is required with long life, huge drive torque, high pointing accuracy, and adaptability to complex and atrocious space environments. Harmonic drive with many advantages such as high reduction ratio, high loading capacity, and high transmission accuracy is applied widely in aeronautics and astronautics fields. Design constraints are briefly analyzed, advantages and disadvantages among three different types of dualaxis structures are campared, a type of structure layout for the single axis actuator is proposed, motor, harmonic drive and angle sensor as well as of points thermal interface design are analyzed, critical techniques including compact and lightweight structure, long life lubrication and high pointing accuracy are pointed,and finally test methods for verifing long life and high pointing accuracy are presented in this paper.

Idealized double twobody model is the universal hypothesis in the preliminary trajectory design of deep space. Based on this hypothesis, a return trajectory model with six independent parameters of lunar spacecraft from parking moon orbit including direct return trajectory and indirect return trajectory is designed. Through analysis of the six independent parameters and the time of outlet point, general relationships between constraint and free parameters are given to easily search certain trajectories. Finally, effectiveness of the proposed approach is validated by examples including two types of return trajectory.

The deepspaceexploration return vehicles reenter the earth under bad conditions. To assure the safety and accuracy of the return, both the LQR referencetrajectory and the predictorcorrector guidance approaches are researched in this paper. The adaptability to the reentry initial conditions and the landing accuracy of both guidance methods are analyzed. The numerical simulation is done for the return vehicles with the second cosmic velocity. The results show that the predictorcorrector guidance gets smaller range error than the LQR one with a relatively large initial error.