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 problem of singleaxis fast attitude maneuver for a flexible spacecraft is considered in this article. By converting the system into unconstrained model and applying the Minimum Principle, the solution for timeoptimal control is obtained, while the necessary and sufficient condition for the controlswitching time are also established under boundary value conditions. By analyzing the nonlinear equations of the switching time, it is found that if the damping coefficient can be negligible, the control input for timeoptimal attitude maneuver will be a symmetrical function of the maneuver time, and the complexity of solving the nonlinear equations could be reduced greatly. Based on this kind of symmetry, an analytic solution for the rigid body with only one flexible mode is obtained. To overcome the difficulty in solving the nonlinear equations with large number of flexible modes, the system is then transformed into a discretetime model and further simplified to a set of constrained leastsquares problems. Finally, numerical simulation is conducted to demonstrate the effectiveness of the method.

Minimumtime orbit transfers for solar sail spacecraft base on genetic algorithm (GA) are explored.The optimal control law is derived from the maximum principle, minimum time consumption is chosen as the optimization objective, and the launch date, the arrive date and the initial guess of the costates are optimized by using GA. In order to deal with the design of transfer orbit with many constraints，the penalty function is introduced into the GA.Orbit transfers for solar sail from Earth’s synchronous heliocentric orbit to Mars’ synchronous heliocentric orbit and from Earth to Mars have been designed and simulated as examples. Results of simulation show correctness and practicability of proposed orbit transfer algorithm.

A finitetime control scheme is proposed for satellites formation flight. Firstly, the nonlinear relative dynamics equation of formation flying satellites is established. Then, the sliding manifold with global fast convergence is designed, and a finitetime control technique is also presented, to guarantee global fast stability of the closedloop system and fast convergence of the tracking errors. The theoretical proof of stability and convergence is given subsequently. Finally, formation keeping scenarios are carried out to validate the effectiveness and advantage of the finitetime control method. Simulation results show that the developed finitetime control technique has faster convergence time and better control accuracy than traditional linear and terminal sliding mode controls.

According to the feature that satellite cables belong to flexible parts, the cable geometrical model and dynamics model based on physical characteristics are proposed by using a dynamic nonuniform rational BSplines (Dynamic NURBS). Cable stiffness, mass distribution, internal deformation energy and other physical quantities are taken into account in models, real shape and movement of the cable can be effectively reflected and the cable’s length can be estimated in the virtual environment, which provides theoretical support for virtual assembly of cable satellite.

In this paper, the FPGA configuration frame format and configuration frame type and distribution in the whole device is analyzed. Based on the analysis, the detail of bitstream read back fault detection method and reconfiguration fault tolerance method is presented. Finally, a circuit implementation with this paper’s fault tolerance method is illustrated. The research  contributes to the circuit fault tolerance design method of all satellite subsystems that use SRAMbased FPGAs.

In order to remedy some defects caused by only one regulating variable in the current DC/DC converters,a new concept to add a control variable in the converters is proposed. In this paper, the chosen control variable is the operating frequency f of the converter, and make f decrease according to a set law when the voltage of the primary power supply increases, so as to solve some problems such as severe heating of the switching transistors, etc. when the frequency is invariable. The main problem to realize the idea is to construct a variable frequency type of pulse width modulator ( PWM ) by a general PWM. Now taking the general modulator 1842/3/4/5 and 1842A/3A/4A/5A as examples to discuss how to construct the variable frequency type of PWM, and show the realizable frequency characters are given. Both of the linear transform circuits of voltages are designed to transform the voltage of the primary power supply to the input voltage of the PWM. The representations for calculating parameters of elements in these circuits are derived. The experiment results are consistent with designs.

The microwave responder is one of the important cooperative targets for spacecraft rendezvous and docking. To insure that the equipment has a good performance inorbit, a special simulator is needed in the spacecraft test before it is sent to orbit. A design of the simulator is introduced in this paper, and the corresponding function and performance requirments are analyzed. Based on the modularized design, a hardware and software design method is provided in this paper. Finally, the operating process and the test performance of the designed simulator are introduced.

The life of infrared earth sensor can not be validated because of its life limitation, weak parts are confirmed through reliability analysis and a life test is designed to carry out test verification on the ground. The combination of real life test and accelerated life test is taken to simulate the working environment in orbit, in order to assess the actual life of the product and know the real situation of operating performance and failure mode of tested pieces in the extreme condition product sealing function failure at the end of its life. It is shown that test pieces are working in a stable period, by analyzing key parameters of the life test and test data of shafting run in the extreme condition. It can be drawn from reliability analysis and life test data that the 3 years practical lifetime of infrared earth sensor for the target spacecraft has a greater margin and the life test can be used to verify the actual life of the product further.

Based on the research of the residual propellant gauging methods in 11th Fiveyears Plan, the application conditions such as propellant gauging methods of the gas injection, ultrasonic flowmeter, accelerometer, thermal, PVT(pressure, volume, temperature) and bookkeeping are analyzed. In this paper, the application of    propellant gauging methods in each period of the spacecraft life is researched, and three typical combinational applications of these methods are brought up.