The CCDbased vision sensor is critical equipment at the proximity phase of rendezvous and docking. The radiation of solar and earth may enter  the field of view of the vision sensor inorbit for rendezvous and docking. The method of stray light modeling is provided, the influence of solar and earth radiation is analyzed, and the design parameters of baffle for vision sensor is determined on the basis of the analysis. Finally ground and flight test results are presented.

Application of an optical navigation sensor (ONS) is a tendency for deep space navigation in the world. The method has many benefits such as good autonomous ability, high accuracy, low cost and so on. Design, development, and validation of the ONS are first discussed in China in this paper. The main technical parameters of the ONS are as follows: the focal length is 953.8mm，the field of view is 0.8°×0.8°,the limiting detection sensitivity is 12 Mv, the measurement accuracy is 0.5″, and the dynamic range of response is 100∶1. The test results in the laboratory and stars observation station indicate all the technical parameters of the ONS’s prototype example are realized.

The star sensor’s low frequency error (LFE) caused by the thermal distortion of the star sensor and its bracket may affect the satellite attitude determination accuracy. In order to cope with this problem, a low frequency error calibration method for star sensor based on the landmark information obtained from the payload is proposed by using the least square (LS) algorithm in this paper. Considering that the satellite attitude determination system serves for the payload, the purpose of the paper is to make the attitude information from the satellite attitude determination system consistent with that from the payload, such that the boresight of the payload can be determined accurately. Thus, the landmark information provided by the payload is adopted as the observation for the calibration for the low frequency error of star sensor. It is illustrated from numerical simulation that the effect of the star sensor’s low frequency error on attitude determination accuracy is effectively decreased by using the proposed method, and the satellite attitude determination accuracy is improved.

The influence of star senor’s characteristics on the accuracy of satellite angular rate estimation is emphatically studied. The relation between errors of star image centre extraction and accuracy of star vector measurement is analyzed. The accuracy of the leastsquare angular rate estimation is deduced, and then factors influence of error of star image centre extraction,the exposure time, the number of star vectors and their mutual angles on estimation accuracy. A simulation is set based on the given APS parameter and actual star distribution.

For a spacecraft with three orthogonally placed reaction wheels and the fourth skewed one, the attitude control problem is considered in this paper when two of three orthogonally placed wheels failed. Under the assumption of zero system angular momentum, a piecewise decoupling control law is designed to stabilize the attitude of the spacecraft by using two residual wheels. By using the Euler angle parameterization method, spacecraft dynamics and kinematics are given and a piecewise control law is proposed based on the decoupling property of the model. After six maneuvers, the proposed scheme ultimately guarantees the attitude stability. Simulation results demonstrate the proposed control approach.

The model of swingscanning infrared earth sensor (SSIRES) at different orbit altitudes is researched in this paper. By reviewing the swingscanning Process of infrared earth sensor(IRES),two models, as disk model and 3D model are introduced by taking orbit altitude as a variable. The transformation from the earth chord length to attitude angles is established approximately in the condition of spacecraft closing to zero position, so that the models are extended to all kinds of orbits through calibration for interested orbit altitudes .

Basic constitute and control principle of HRG is presented in this paper. The design of frequencydomain characteristic of the moment rebalance control loop as the core control loop is described，and its implementation is given. Verification test indicates that the design is correct and available.

Inertial sensor assembly is a high accuracy measure assembly in the spacecraft attitude control system. It needs not only to detect and dispose lowpower level signal, but also to export the highpower signal to control temperature and to drive motor. Electromagnetic compatibility design is a key technique which determines the performance of the product. According to the development experience of products, the way of controlling inner and outer interferential sources in products is proposed by analyzing typical case.

Several typical thruster configurations of spacecrafts for rendezvous and docking task are introduced. Advantages and disadvantages of these thruster configurations are analyzed through comparison. Accordingly, some useful conclusions for the thruster configuration designs are obtained. In addition, some infectors that should be considered when designing a thruster configuration for rendezvous and docking task are also proposed.

In satellite electronic device testing systems, the ground testing software shall access testing system hardware, exchange information with users, and implement test data storage and management simultaneously, the concurrent processing capability of the testing software is the key technique to guarantees the reliability and the validity of the testing system. The sequential programming methods used in traditional software design can not meet the concurrence processing requirement in the satellite ground testing software. In order to resolve the  above mentioned problem,  a programming method combined with four concurrence mechanisms, multithread, message mechanism, shared buffer, and PV operation, is presented in this paper. Practical application shows that the multi concurrence mechanism based software design method can implement satellite ground testing software reliably, and resolve the parallelism problem in software development effectively.

Faulttolerant design is an important way to improve the reliability of spalewire CODEC. The faulttolerant ability of SpaceWire is analyzed and the faulttolerant design of SpaceWire CODEC is introduced, then validity of the design is demonstrated by fault injection and simulation.

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.