The Chang′e-6 mission is the first lunar far-side sampling return mission in the world. The mission planning is an important assignment in the spacecraft system design. Mission planning problem is conventionally converted to the numerical constraints optimization problem in the traditional method. However, by using this method, the constraints could not be satisfied directly, and at the same time, the design variables dimension increased, the computing efficiency decreased, and the physical meaning was not apparent. A new mission planning approach using the mission geometry is proposed in this paper. According to the new approach, the periapsis constraints of the EarthtoMoon and MoontoEarth transfer trajectories, the relative geometry constraints on sampling point and lunar orbit plane, and the geometry relationship of the Sun-Earth-Moon are fully considered. The new approach reduced the design variables dimension, improved the computing efficiency. In addition, it also has the advantage of intuitive spatial geometric interpretation, and the result can be verified easily. The new approach has been applied to mission planning and trajectory design in the Chang′e-5 and Chang′e-6 missions successfully, and will be adopted in the future lunar exploration mission.

Queqiao-2 provides relay services in its overall life cycle for Chang′e-6, Chang′e-7, Chang′e-8, and Chang′e-4 missions, as well as lunar exploration missions from international cooperation. To meet the different relay needs of multiple launch windows and different flight phases of Chang′e-6 mission, the trajectory design method and process of the Queqiao-2 relay satellite to service the Chang′e-6 mission are introduced from the perspective of mission analysis. The adoption of a lunar elliptical sunsynchronous frozen orbit satisfies the requirements for communication range, coverage, launch windows, and orbital phase for the Chang′e6 mission. Then the trajectory scheme and the control strategy adopted on each flight phase are optimized with minimum fuel constraint. A four-impulse joint optimization trajectory control strategy is employed to significantly alter the orbital ascending node, inclination and argument of perilune with few velocity increments. Finally, the status of the in-orbit flight and relay support for Chang′e-6 mission are shown respectively. For the first time internationally, Queqiao-2 has achieved a significant change of right ascension of the ascending node in lunar orbit by using a lunar low energy capture orbit, and close range and long arc relay of target on the far side of the moon by using the lunar large elliptical sun-synchronous and frozen orbit. The result shows that the in-orbit flight of Queqiao-2 is consistent with that of the design, which ensures the successful implementation of the Chang′e-6 lunar far-side sampling and return mission. The design method can provide a reference for the orbit design of deep space exploration missions.

 According to the characteristics of China 's Chang 'e-5 unmanned autonomous lunar orbit rendezvous and docking mission, a millimeter-wave rendezvous and docking radar system is designed and implemented, which integrates angle acquisition and tracking measurement with single channel interferometer and rotating mechanism, double one-way pseudo-code ranging and carrier velocity measurement, and TURBO coded spread spectrum communication. The specific composition, working process, measurement and communication methods，design performance and verification of the new radar system are described. The new system has the characteristics of multi-information measurement and two-way communication integration, wide field of view, large distance dynamics, high measurement accuracy, strong anti-multipath capability, small weight and low power consumption. Verified by ground and actual on-orbit applications, the millimeter-wave radar system realizes the full-range high-precision measurement and two-way communication of lunar orbit unmanned rendezvous and docking for the first time. The performance is stable，and the dynamic range of ranging can reach 0.2m to 397km, the accuracy can reach ± 0.017m, the field of view of angle measurement can reach 360°× 130 °, and the accuracy is better than ± 0.07 °.The system can be applied to the Chang′e-6 mission.

Space solar power (SSP) system, a major type of space-based power-generating equipment, is an important infrastructure providing massive, continuous, and stable green electricity by utilizing solar energy in space. Many countries and organizations consider SSP to be one of the most promising clean energy sources. The historical activities of SSP in the world are summarized. This review focuses on the significant development of SSP during the last 10 years, which is the most important period for SSP. The latest international SSP development programmes in the United States, ESA, Japan, China, UK and Korea are presented. Some significant solar power satellite(SPS) concepts proposed in the decade, including typical SPS-ALPHA, MR-SPS, CASSIOPeiA SPS, et al., are introduced. The technical and non-technical challenges are also listed and several important in-space demonstration missions in recent years and in the near future are introduced. The conclusion is that the next 5 to 10 years will be an important period for rapidly developing the key technologies and conducting on-orbit demonstration and application. Controlling the mutual position relationship between the solar array and the transmitting antenna becomes a core issue to be considered in the innovative design of the SPS. Wireless power transmission technologies would be the demonstration focus for the first step. It is expected that the first commercial SPS would be constructed as early as 2040.

Most of the existing space solar power station (SSPS) schemes are conceptual designs with poor comprehensive performance. So a highly feasible engineering design scheme of SSPS via cylindrical modular concentrator array (SSPS-CMCA) is proposed. Firstly, the SSPS schemes proposed by scholars and organizations from various countries are analyzed. The scheme combination table is listed from the perspective of structure, control, shape, etc. A set of scheme standards for the design of SSPS is summarized. Then the SSPS-CMCA scheme is proposed according to this standard. The structural design of the scheme is described in detail, and the performance of each aspect and the system parameters of 1GW model are analyzed. Finally, the construction scheme of the whole process from ground experiment to space assembly and operation is proposed. 

Lenticular boom is used as the main supporting rod of the thin-film solar array of space solar power station. The envelope volume and weight of active driving mechanism commonly used to control the deployment of the boom are relatively large. The control approach of boom deployment is a decisive factor of the storage ratio, which can significantly reduce the envelope volume and weight of the solar array through optimization. It is regarded as a promising method for establishing the ultra-large and light-weight space solar power station. In view of this, the passive deployment of composite thin-walled lenticular boom (CTLB) with full utilization of the elastic recovery characteristics of its cross-section is investigated in this paper. With the explicit dynamics finite element (FE) method, the mechanical responses of CTLB during flattening and coiling are analyzed in terms of the elastic stress and strain. Then the peel force of the flexible fastener during detachment is numerically obtained. The flexible fastener is chosen as the control unit for CTLB deployment. By inheriting the deformed shape and material state from the proceeding FE analysis and simulating the flexible fastener using cohesive element, a dynamic model is established for the deployment of CTLB. Then the influence of the mechanical properties of flexible fasteners on the unfolding characteristic of CTLB is numerically studied. The results show that the orderly deployment of CTLB can be achieved by adjusting the detachment force of flexible fastener, which provides a solution for further reducing the complexity and weight as well as improving the storage ratio of ultra-large thin-film solar array.

 The compressed stacking structure has potential applications in space structures such as space solar power station and large antennas due to its high utilization of fairing space and easy separation and reassembly after entering orbit. The compressed stacking structure has multiple components and complex connection forms, and the modal testing method based on discrete acceleration sensors cannot monitor the full field modal information of the structure, making it difficult to evaluate the modal of locally connected structures. Therefore, a digital twin modal testing method for space compressed stacked structures was studied. Firstly, a high-precision simulation model of the compressed stacking structure was established and modal analysis was conducted to obtain the structural vibration mode; Then, modal tests were conducted using the hammering method and accelerometer data was collected; Finally, the simulation analysis data and experimental sensor data were fused to construct a compressed stacking structure digital twin, achieving real time monitoring of the entire field mode. The accuracy of predicting the vibration mode of the digital twin body reaches 99.7%, and the fundamental frequency error is 2.6%. It takes 2 seconds from sensor data to completing the calculation and cloud map display of the digital twin body. The results validate the high prediction accuracy and efficiency of proposed method, indicating effectiveness in improving the monitoring range of spatial structural modal experiments.

In order to control the heat source temperature and to improve the thermoelectric conversion efficiency of RTPV, a high-temperature thermal control method based on aerogel discrete spacer MLI is proposed, and the thermal model is established. The accuracy and applicability of the model are verified by experimental results. The thermal performance of the RTPV is analyzed by using the model. The results show that the heat source temperature is 1078 ℃ when the emissivity of the emitter is 0.145 and that the emissivity of the reflector is 0.162. The heat source temperature increases monotonously with the decrease of the emitter emissivity and the reflector emissivity. When the number of MLI layers is less than 20, increasing the number of layers can significantly improve the insulation performance of MLI. The results provide a basis for thermal design and optimization of RTPV.

With the rapid development of the space industry, traditional Entry, Descent, and Landing technologies are failing to meet the demands of modern space missions. Inflatable and mechanically deployable aerodynamic decelerators are expected to be used for the recovery of various space missions due to their foldability, small size and light weight. The research progress and current status are presented around inflatable and mechanically deployable aerodynamic decelerators. Firstly, the flight tests conducted by various countries are introduced on inflatable and mechanically deployable aerodynamic decelerators. Then, it presents the research progress of inflatable aerodynamic decelerator technologies. Four main areas are presented, including structural configuration, aerodynamics, aerothermodynamics and aeroelasticity. Thirdly, the technologies for mechanically deployable aerodynamic decelerators are presented, focusing on the areas of structural design and optimization, aerodynamics, aerothermodynamics, and attitude control. Finally, an insightful perspective on the future technological development of aerodynamic decelerator is presented, highlighting a potential innovative approach for the recovery of launch vehicle sub-stages, spacecraft, and so on.

Aiming at the problem of precision decline caused by orbital perturbations during the descending orbit interception of thin film spacecraft, an orbital control optimization algorithm based on optimal attack and sideslip angles is proposed. Firstly, the algorithm calculates the interceptable region based on the orbital dynamics characteristics of the thin film spacecraft, then determines whether the target is within the region to narrow down the search range for the optimal orbit. Drawing inspiration from the strategy used in reentry spacecraft which controls attack and sideslip angles, the algorithm seeks to minimize the miss distance as the objective function while subject to the constraint of fixed attack angle for thin film spacecraft. The particle swarm algorithm is employed to find the optimal solution for the attack angle. Lastly, to mitigate the effects of non-spherical Earth perturbations, a slight adjustment to the normal direction of interception orbit is achieved by introducing sideslip angles, thus enhancing the interception probability. Simulation demonstrate the thin film spacecraft can intercept the co-planar targets in low earth orbit with good terminal accuracy under both ideal gravitational filed and non-spheriacl perturbation gravitational filed. The work can provide a new idea for the removal and mitigation of space junk, and also has a certain reference significance for the development of new low-cost orbital interception.

Based on the requirement of flexible vibration measurement and dynamic parameter identification of the solar array in China space station, China has successfully applied a wireless accelerometer sensor network in orbit for the first time. Firstly, the architecture designs of the wireless sensor network in Tianhe core module and Wentian lab module are introduced. Then, for the requirement of multi-source data fusion of modal identification, the responsive TCP/IP time synchronization based on wireless delay self-optimization compensation and the broadcast UDP time synchronization based on wireless reference node are proposed. To address the issue of measurement interruption of wireless node communication links caused by the possible obstruction of the solar array, an autonomous wireless network topology dynamic management method based on relay transmission is proposed, achieving high reliability transmission of vibration measurement data. The in-orbit application of wireless accelerometer sensor networks in Tianhe module and Wentian module is summarized. The flight measurement data show that the time synchronization accuracy is better than 20ms for TCP/IP mode and 70ms for UDP mode. The highly reliable acquisition capability of wireless network is evaluated by using the measurement data of active excitation conditions at a 180° angle of the solar array, meeting the requirement for modal identification. Finally, based on the flight test results of space station solar array vibration measurement, several inspirations are proposed for the widespread application of future wireless sensor networks in orbit.

Aiming at the inflexiability of network traffic monitoring and tardy response to network emergencies for LS-LEO(large-scale low earth orbit) satellite networks, a method of network traffic monitoring for LS-LEO satellite networks based on INT(in-band network telemetry) is proposed. In order to reduce the overhead of LS-LEO satellite networks traffic monitoring, the proposed method combines the high dynamic topology characteristics of LEO satellite networks, and provides aggregation node selection of network traffic monitoring information within each domain of LEO satellite networks and leader node selection based on microburst traffic, which realizes the rapid monitoring and reporting for abnormal state information of network traffic. In the domain-based aggregation node selection strategy, the aggregation node is obtained by calculating the shortest path and delay. The microburst node is monitored by the telemetry information reporting strategy of the leader node, and the telemetry information is directly reported by the node to the aggregation node. The simulation results show that， compared with the traditional traffic monitoring method, the proposed method can effectively reduce the overhead of telemetry packets and realize flexible low-overhead monitoring of bursts in traffic. Taking the LEO satellite network scale of 500 nodes as an example, the average telemetry overhead for a flow with a path length of 7 hops can be reduced by 27.77%.

Since the medium and high orbit satellites have long on-orbit time and obvious temperature rise at the end of life, equivalent comprehensive degradation performance of the common coating OSR is studied through analyzing 5-year in-orbit temperature data of a MEO navigation satellite. Its thermal control design and orbit solar flux are taken into account, and how the radiating surface design influences the temperature variation in long term is also discussed. In this way, thermal control design can be optimized, and long-life as well as highly reliable design scheme can be obtained. OSR′s comprehensive degradation performance in this paper is appropriate for thermal control design and temperature prediction, which also shows that the method of solar flux normalization proposed here can be applied to study coating comprehensive degradation and in-orbit temperature variation. Analysis of the radiation surface design′s influence on long-term temperature rise and some design suggestions based on it have engineering reference value for the long-life operation of medium and high orbit satellites.

Aiming at the problems of short time and low precision of jet thrust control in air bearing table simulation test, an improved fuzzy cascade control scheme based on small thrust difference of rotor wing is proposed. Firstly, the characteristics of rotor wing thrust were modeled, the thrust output characteristics of single-channel rotor wing and jet were compared, and the feasibility of low-thrust rotor wing propulsion instead of jet were explained. Secondly, aiming at the dead zone problem of rotor wing thrust, a differential thrust method for rotor wing thrust distribution was proposed to solve the engineering difficulty of air bearing table control. Finally， a fuzzy adaptive cascade rotor wing thrust controller based on differential thrust method was designed. On the basis of this, the correction coefficient was introduced to adjust the intermediate volume ratio of inner and outer ring transfer online. The controllercan increase the position and attitude convergence speed and the control accuracy of the system, improve the robustness of the air float platform., and smooth the speed curve. The experimental results show that the fast displacement tracking adjustment time of the unit step signal is less than 40s, the control accuracy is less than 0.4mm, and the standard variance of error is 2.2×10^-4. This method can realize fast and accurate trajectory tracking control of the air-bearing platform under the action of the low thrust rotor, and provides a new way for the air bearing table platform to accurately simulate the space motion.

Micro-thruster is one of key payloads on high precision satellites. It is meaningful to pretest their thrusts in advance on ground. However, it is of difficulty to measure the thrust below micro-Newton on ground due to factors such as gravity acceleration and environmental disturbances. An integrated micro-thrust stand based on vertical pendulum (VP) structure is designed, to overcome above restrictions. The stand has an adjustable accuracy and can cover a wide dynamic range, in which the differential measurement method and integrated cable assembly are employed. Its performances are then characterized in common lab environment. The results show that the resolution in a fine mode can reach 10 nN,the spectral noise can reach 10nN/Hz^1/2  @ 10mHz and the range reaches 1mN, the measurement range in large range mode can reach 300mN, and the response bandwidth is DC-10Hz when the tested thruster has a mass of 330g. The stand can be calibrated whenever it is needed by both gravitational force and electromagnetic force. It can meet the on-ground test requirements of a variety of micro-thrusters.

Aiming at the security problems in the distributed satellite systems, two distributed secure beamforming algorithms based on eavesdropper location information are proposed to improve the secrecy performance. Considering the inaccurate location information of eavesdroppers, two optimization problems are formulated to maximize the achievable secrecy rate and minimize the transmit power, respectively. To solve these non-convex problems, the norm bound of channel state information is derived based on the eavesdropper location information, then, triangle inequality and Cauchy Schwartz inequality are adopted to deal with the channel error. After that, two distributed beamforming algorithms based on angle domain information are proposed to obtain the closed-form beamforming weight vectors with low computational complexity and fast convergence rate, providing theoretical basis for the deployment and application of distributed satellite communication system in the future. Finally, the effectiveness and superiority of the proposed scheme were confirmed through simulation results, where the secure communication can be achieved with lower complexity compared with traditional semidefinite programming algorithm, and the performance is improved by about 30% compared with the non-robust algorithm.