The solar array drive assembly is a key single machine for a three axis stable satellite, driving the solar array to orient towards the sun and transmitting electrical power to the entire satallite, whose reliability and lifespan directly affect the lifespan of the entire satellite. From the 1960s, after decades of development, significant progress has been made in transmission power, inorbit lifespan, reliability and low disturbance characteristics. In the past decade, with the development of low orbit satellite constellations and commercial aerospace, a strong demand for low cost, lightweight and dual axis configuration products has been displayed. This article summarizes the technological development direction and model application characteristics based on information from relevant fields, and proposes that the ultrahigh voltage and high power technology, as well as the the commercial aerospace application of low cost technology are the main development directions for the future. the future development trends.

In order to adapt to the unprecedented development of satellite development technology and scale, and achieve high reliability and low cost at the same time, the author reorganizes the traditional subsystem design methods according to the characteristics and disciplines, and puts forward the domestic initialtive subject chain design method, which has been successfully applied to the beidou3 experimental satellites, networking satellites and backup satellites, realizing the rapid mass production networking of Beidou3 (BDS3) satellite. This paper describes the design concept based on the subject chain, and focuses on the attitude and orbit control technology of beidou3 satellite based on the subject chain. The onboard operation results show that the satellites function continuously and reliably, and fully meets the high control accuracy requirements of Ka and laser intersatellite links and the requirements of navigation satellites for high continuity.

Lingxi03 satellite developed and launched by Galaxy Aerospace is the first flat panel satellite with a flexible solar wing of more than 20 square meters. Compared with traditional satellites, the vibration of flexible solar wings will bring low frequency disturbance to the attitude control system. Additionally, the satellite adopts a single wing asymmetrical structure to make the interference torque more than 50 times that of conventional satellites. According to the characteristics of the structure and layout of the Lingxi03 flat panel satellite, an attitude controller combining PIDT and structural filter is designed, which is based on frequency band isolation. The response at the modal frequency of the solar wing is targeted to suppress the vibration of the solar wing to ensure the control accuracy and suppress the vibration of the solar wing. The simulation results show that the controller can achieve the attitude control accuracy of 0.03°(3σ) and the stability of 0.005(°)/s (3σ).

The communication satellites deployed in geosynchronous orbit are expensive and the orbital resources are scare. The influence range of onorbit mission is wide and the reliability requirements are high. Based on the fact that the acutal life of some spacecraft in orbit is generally higher than the design life,it designs the theory life of geosynchronous orbit communication satellite offorbit backup scheme and studies the using mode,constraint conditions and backup efficiency of offorbit satellite, which can provide reference for the onorbit backup design other geosynchronous orbit satellites.

APSTAR 6E satellite is the first full electric propulsion satellite in China. It is launched by a low cost rocket, and the separation orbit is LEO. In response to the characteristics of small thruster and long orbit transition, in order to shorten the orbit transition, the satellite to use the Separate Propulsion System for orbit raising, after completing the task, it can be thrown away. Due to the short working time in orbit, the Separate Propulsion System adopts a low cost design. This scheme presents a scheme of using Separate Propulsion System for orbit control and an autonomous departure after separation from satellite. This scheme reduces the dead weight brought by the Separate Propulsion System, reduces the time of orbit raising by electric propulsion, and reduces the consumption of xenon propellant, it has great advantages in improving the capacity and the service life of satellite.

In recent years, benefited from the rapid development of platform technologies such as electric propulsion, power supply and thermal control, the major manufactures of foreign GEO communication satellites have successively completed the research, development and finalization of the new generation of GEO all electric propulsion satellite platforms. These new generation platforms have significant reference significance for the development of China GEO satellite platforms in terms of design concepts, application of new technologies and low costs. Combing the technologies and characteristics of the new generation common platform of various types of GEO communication satellites abroad like Spacebus NEO, Eurostar NEO, Space Inspire, OneSat and etc., the market trends, development and design ideas of new generation GEO communication satellite platforms are analyzed. At the same time, the design requirements and technical directions of the platform related electric thrusters, power processing units, thrust vector adjustment mechanisms and thermal control are analyzed. Development suggestions for the common platform of China next generation GEO communication satellites are proposed, which can serve as a reference for the development of China future GEO satellite platforms.

For electric propulsion geosynchronous satellites, orbit station-keeping can be carried out by electric propulsion layout with double robotic arms, and each robotic arm is equipped with one electric thruster. In this paper, a parsing-oriented approach is used to orbit station-keeping, and the East-West/North-South Station-Keeping maneuver is achieved by firing 1~2 times a day. The ignition direction remains unchanged for several days. Firstly, the station-keeping algorithm with analytical form is studied under the normal condition of double arms and thrusters. Then, a failure scheme based on the single arm with only one electric thruster is researched. Finally, a method to measure the efficiency of different vector mechanism capability and thruster specific impulse is given analytically. The simulation results show that the control accuracy can meet the requirements for geosynchronous satellites station-keeping, multi-GEO satellites collocation and fixed-point orbit injection.

High orbit spacecraft typically carry a large amount of liquid fuel, and during orbital maneuvers, the motion of the liquid can cause significant disturbances to the spacecraft's attitude. Developing a simple yet accurate dynamic model for liquid sloshing has important engineering value. This paper introduces several research advances related to equivalent dynamic modeling of liquid sloshing during the development of high orbit spacecraft with large fuel tanks. These include the equivalent modeling methods for arbitrary shaped 3D tanks applied to attitude maintenance modes, and time varying parameter equivalent mechanical modeling methods for large angle attitude maneuvers. The study presents a variable parameter equivalent modeling method based on the continuity of static moments and inertia, and numerical simulations verify the accuracy of the proposed methods.

In this paper, the modeling, control and simulation technology is studied to improve the control performance of spacecraft with complex structures in multi physical field. Based on the absolute nodal coordinate formulation (ANCF), the thin plate element unifying the temperature and displacement field is given, which can avoid mismatch problem in the process of mapping two physical fields. On this basis, considering the variation law of material Poisson's ratio with temperature, a rigid flexible thermal dynamic model of spacecraft with the bidirectional coupling effects between the displacement field and temperature field is established. A high order control algorithm for active suppression of multiband disturbance is designed, achieving the index of control stability of flexible spacecraft with a moment of inertia of  better than 5e-6°/s under 8°/25s attitude maneuvering condition in mathematical simulation. Based on the three axis air flotation motion simulator, the flexible load simulator and the thermal radiation environment simulator, a full physics attitude control test platform with rigid flexible thermal coupling is constructed to calibrate the dynamic model in time domain, which verifies the engineering feasibility of the integrated simulation of finite element modeling and control.

Pulsed Plasma Thruster (PPT) is the earliest electric thruster used for space missions. Among them, the Ablative Z-pinch Pulsed Plasma Thruster (AZPPT) can achieve high thrust to power ratio and specific impulse at a low volume and power due to its special front back coaxial configuration and anode spike. The effects of anode spike length, charging voltage, and discharge frequency on the performance of AZPPT was studied in this paper. The experiment showed that with the increase of charging voltage, the peak discharge current of the thruster, the electron temperature of the plume plasma, and the impulse bit increased, proving that the wide range adjustment of thrust can be achieved by changing the voltage. When the height of the anode spike increases, the impulse bit of the thruster decreases slightly, while the specific impulse increases significantly. Compared to the performance of thruster without the spike, the specific impulse increases by 23.7%, and the overall efficiency increases by 7.31%, indicating that anode spike can optimize the thrust performance of low power AZPPT significantly. Based on the above results, a 1 U miniaturized and modular propulsion system has been developed. The ignition of the propulsion system was successful, and the impulse bit reached 193.72 μN·s under a discharge energy of 5 J. It has the characteristics of low power, good overall performance, and compact structure.

The thermal propellant gauging system (TPGS) method is a method for measuring the residual amount of satellite propellant with active heat source input. It calculates the residual amount of propellant according to the temperature variation data of the measuring point when the tank heated by a specific power heating plate, and the measuring point is attached on the tank wall. In order to obtain an accurate thermal analysis model based on the gas-liquid distribution characteristics of plate tank in orbit in the microgravity environment, numerical simulation of the TPGS method is carried out with computational fluid dynamics (CFD) method to analyze the influence of different acceleration conditions, filling ratio and measuring point position on the accuracy of the TPGS method. The results show that the propellant climbs up along the deflector in the microgravity condition, and the wall area covered by the fuels larger, which results in the temperature variation in microgravity condition is obvious different compared with ground condition. It is found that, in the microgravity condition, when the measuring points are distributed away from the heat source, the temperature resolution of the measuring points is higher, and the measurement accuracy of the TPGS method is higher.