Aiming at the landing problem of a no propulsion or only weak maneuverability probe in a weak gravity binary asteroid system, a low energy landing orbit design method based on unstable manifold and Poincaré section method is proposed. By searching for periodic orbits near the collinear libration points, calculating the eigenvalues and eigenvectors of the monodromy matrix, and solving the unstable manifold of the periodic orbits, the fuel consumption during the landing process is greatly reduced. For a lander with weak maneuverability, an orbit design method based on connecting splicing unstable manifold and descent orbit set is proposed, which obtains a low energy landing orbit that satisfies the terminal constraints. Numerical simulation verifies compared with the method of calculating the double impulse landing trajectory by genetic algorithm, the method proposed in this paper can flexibly add terminal constraints and design the target landing point, and does not need to design penalty function and perform iterative calculation, with high computational efficiency.

This paper focuses on the high accuracy navigation method based on the information fusion of star camera and inertial measurement unit (IMU) for aerial vehicles. Among the existing navigation methods for aerial vehicles, the integrated navigation of inertial navigation system (INS) and global navigation satellite system (GNSS) suffers from the risk of performance degradation in the radio signal denied environment. In the traditional INS/CNS (celestial navigation system) integrated navigation system, the drift of the gyroscopes in the inertial measurement unit can be suppressed, while the zero bias of the accelerometers cannot be eliminated effectively. In order to cope with this problem, a novel autonomous navigation approach based on the line of sight (LOS) measurements of space targets is presented, where the LOS vectors of space targets and background stars with known ephemeris are observed by the star camera, and the state of the vehicle is predicted by using the IMU. The position, velocity and attitude of the vehicle are estimated together with the calibration of the IMU measurement bias via the extended Kalman filter (EKF). In addition, an optimal selection strategy of observed targets based on the Cramer Rao lower bound (CRLB) is designed. The effectiveness of the presented method is illustrated through the visibility analysis of the space targets, the observability analysis of the navigation and the numerical simulation of the navigation filter.

The observability analysis and navigation filtering algorithms during Mars aerocapture are studied. The observability of the system is analyzed using a method based on Lie derivatives, with the effects of atmospheric density uncertainty on the system's observability degree being considered. The unbiasedness of the Schmidt Kalman Filter under the condition of time varying parameters is proven, in response to the system model nonlinearity and atmospheric density uncertainty. The Extended Schmidt Kalman Filter (ESKF) algorithm is introduced, which effectively improves the accuracy of state estimation during aerocapture. Through simulation verification, the ESKF algorithm shows better estimation performance compared to traditional methods, providing effective theoretical and methodological support for the execution of Mars aerocapture missions.

In order to realize the real time measurement and control of geosynchronous orbit satellite during the process of global deployment, a TT&C scheme with comprehensive utilization of satellite ground link and inter satellite link is proposed in this paper. Based on the two stages divided by the scheme, wherein one is in the view stage and the other one is out of sight stage, a time sharing heterogeneous architecture is constructed. In this architecture, satellite ground link is mainly used in stage one when ground station can control satellite directly, and inter satellite link is the only way to measurement the object when it is out of the ground station visual range. A variety of different inter satellite links, such as inter satellite link with observation data  relay satellite and TDRS/BD inter satellite link, are utilized. Furthermore, the communication link budget is calculated and the mission reliability is analyzed. The result shows that the scheme is feasible and reliable.

With the increasing number of space targets, the problem of orbit determination of the targets is becoming increasingly important for space security. Due to the large number and dynamic feature of the space targets that need to be observed, coupled with limited observation resources, it is necessary to dynamically adjust the collaborative observation scheme to efficiently utilize constellation observation resources and ensure that each target has better positioning accuracy. Thus, it is required to solve the mission planning problem of multiple targets using multiple observation satellites. This paper first establishes the orbit dynamic model of the flying targets, as well as the Kalman filter model of the collaborative positioning algorithm using the multiple line of sight information of different observation satellites. Then, a collaborative positioning accuracy estimation model and an observation priority model of the targets based on the Geometric Dilution of Precision (GDOP) is proposed. Based on the above models, a mission planning framework for collaborative observation based on reinforcement learning (RL) is developed. A policy network based on multihead selfattention mechanism is designed accordingly to calculate the planning results. The proximal policy optimization (PPO) algorithm is adopted to train the policy network in a training environment. Compared with the heuristic algorithm based on tracking priority, simulation results shows that the proposed RL method can effectively improve the overall tracking accuracy as well as the total tracking time of all the targets, and can provide faster computation speed compared to genetic algorithms.

An analytical finite time sliding mode controller is proposed for spacecraft attitude reorientation control under pointing constraints, which has the faster convergence compared with the existing asymptotic convergence control algorithms. Firstly, the navigation potential function is established based on the attitude maneuver target and the attitude pointing constraints. For further, the potential function and its gradient information are introduced on the sliding mode surface such that the attitude pointing constraints are satisfied during spacecraft attitude maneuver. To ensure the feasibility of attitude pointing constraints outside sliding mode surface rigorously, a new Lyapunov function is established by combining the sliding mode function and the artificial potential function. Then, based on the designed Lyapunov function and the finite time convergence theorem, a controller is designed to achieve finite time spacecraft attitude reorientation control under attitude pointing constraints. Finally, numerical simulation shows that the proposed controller can achieve fast attitude convergence and ensure the feasibility of the spacecraft attitude pointing constraints.

Aiming at the problem of attitude control for lateral channel of underactuated hypersonic vehicle, a control method based on the auto coupling PD(proportional differential) control theory is proposed. Firstly, the lateral attitude model of the underactuated vehicle is transformed into a second order dynamic model composed of two channels of velocity inclination angle and sideslip angle, and a virtual command of sideslip angle is introduced into the velocity inclination angle channel, and then two total perturbations are defined respectively for the internal dynamics and external perturbations of the two channels. The nonlinear underactuated perturbed system is equivalent to the linear perturbed system with virtual full actuated system. Then, the minimum speed factor and its adaptive speed factor are used to design the auto coupling PD controller of the outer ring velocity inclination angle, so as to obtain the virtual instruction of sideslip angle. According to the virtual instruction, the auto coupling PD controller of the inner ring sideslip angle is designed to obtain the control of aileron deflection angle. Finally, the robust stability and anti disturbance robustness of the auto coupling PD control system are analyzed. The simulation results show that the auto coupling PD controller designed in this paper not only has good anti disturbance robustness, but also has good dynamic quality and steady state performance, which has great application value in the control field of underactuated hypersonic vehicle.

A neural network nonsingular terminal sliding mode (NTSM) control method with block approximation of dynamics model parameters is proposed to address the problem of model uncertainty and unknown perturbations in the robotic arm. First, a nonsingular terminal sliding mode surface is used in order to accelerate the convergence of the system tracking error and avoid the problem of singularity in the traditional terminal sliding mode. Second, multiple selfrecurrent wavelet neural network (SRWNN) chunks are utilized to approximate the unknown dynamics model parameters of the system, and the weights are adjusted by using the adaptive update law. Meanwhile, the approximation error of the SRWNN is compensated by designing a robust control term, and the system stability is proved using Lyapunov stability theory. Finally, the simulation analysis using MATLAB shows that the average steady state error of the joint angle tracking error is reduced by 31.9% and 76.5% for the chunked SRWNN sliding mode control compared with the sliding mode control and the overall SRWNN sliding mode control, respectively, which demonstrates that this method is a reliable and effective trajectory tracking control method.

The reliability design and verification of spacecraft system is a key issue in spacecraft system design. With the increasing degree of complexity, comprehensiveness and intelligence of spacecraft system, the traditional reliability design based on static system architecture is becoming less and less able to meet the design requirements. The model based system design method has advantages in the integrated design of system structure and behavior. On the bases of decomposing the reliability design and analysis of spacecraft control system based on MBSE concept, a generalized spacecraft component level and system level reliability verification model is established using SysML modeling language. And an integrated cross level model and its design process of reliability architecture and behavior design and verification are further established. The parameter transfer and management method within and between models are designed and described via instance table, Opaque element and value property. The design practice shows that the whole solution proposed in this paper can meet the requirements of universalized spacecraft system reliability design and verification, and is conducive to the development of modeling and modularization of spacecraft system design.

A high precision eddy current displacement sensor for space application is developed to meet the demand of precision noncontact displacement measurement for ultra quiet stable device. Based on the working principle of sensor, the impedance model of the sensor coil is established, and the natural nonlinear characteristic is analyzed. Based on the high precision balanced bridge circuit, a nonlinear calibration method is proposed. Without extra calibrating circuit, the nonlinearity of sensor is reduced from 12% to 0.98%, by the design of balanced bridge circuit impedance and demodulation phase. The measurement results show that the nonlinearity of high precision eddy current displacement sensor is 0.98% in the measurement range of ±3mm. And the sensor has already been used in space application.

Demand for star sensor is increasing on near space vehicles. Compared with space environment, near space environment has new factors which influence star sensor’s detectability such as atmospheric extinction, sky backgroud light, high speed, high temperature, high angular velocity, etc. The factors which influence near space star sensor’s detectability are analyzed step by step. The influencing mechanism of detectability is obtained about FOV,sensitivity, SNR, image processing algorithm. The relation between SNR and environment/design factors is quantified. One method of parameter design for near space sensor is proposed. Finally one principle prototype for special requirement is designed， tested and verified. The results indicate the principle prototype can meet the application requirements.