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Abstract: Pure gravity orbit is an orbit removed non-gravity disturbances, which can give a more precise relation between gravity and orbit, and can be used to navigation satellite, earth gravity felid measurement, gravity wave detection and other fundamental physical test mission. In this paper, the author propose a system concept of a new pure gravity orbit acquirement method named inner formation flying system, which treats the PM and the outer satellite as a pair of formation flying satellites. The inner formation flying system consists of an outer satellite and an inner satellite. The outer satellite is a small satellite with a cavity in the center. The inner satellite is a PM flying in the housing of the outer satellite''s cavity. So the inner satellite is shielded from non-gravitational forces of atmosphere drag and radiation pressure and flies along a purely gravitational path. From the view of formation flying, the outer satellite can be treated as a formation satellite with zero or small distance to the inner reference satellite. So the analysis of mechanic and dynamics of inner formation is made at first. Then a control architecture including feed forward and feedback strategies is presented and the controller is designed. The simulation is carried out in a high precision formation flying simulation environment at last to validate the inner formation flying methods. [Copyright &y& Elsevier]

There are various control laws for attitude stabilization and maneuvering such as magnetic torque control, solar torque control, momentum exchange control. Such controllers work in a linear continuous mode. But the torque that they can produce is in the range of 0.02–1 N m for momentum exchange control, 10−2–10−3 N m for magnetic torque control, and 10−5–10−6 N m for solar torque control. Therefore, the speed of attitude maneuvering is limited by the low-level maximal torques that can be delivered to the attitude control system. Thruster control used in attitude control is no linear continuous reaction thrust control. However, they can provide almost any torque level. Their range between 0.01 N m and 30 N m, which is very common in most satellite. It is practical and convenient to use thruster for the rate damping control of small satellite during initial separation from launch vehicle. The paper deals with the analysis and design of thruster attitude control of a small satellite OptimSat. Simulation results shows that rate damping control for small satellite using thruster is effective. [Copyright &y& Elsevier]

This paper innovatively proposes six scale-independent relative orbital elements, that can be used for dynamics analysis and control for relative motion with near-circular orbits and arbitrary inter-satellite separation. Scale-independent relative orbital elements are deduced from the first-order decomposed relative motion model. These elements reveal the twisted elliptical characteristics of large-range relative motion and standard elliptical characteristics of close-range relative motion. The error of the proposed relative orbital elements keeps zero for circular orbits and the error is irrelevant to inter-satellite range. Based on the scale-independent relative orbital elements, the configuration design equation is established. Given the configuration requirements, the initial orbital elements of the deputy can be calculated. The autonomous configuration control problem is studied by means of artificial potential functions. Scale-independent relative orbital elements are used as control variables, and the large-range reconfiguration and collision avoidance control are realized with low control consumption. • Six independent configuration parameters are proposed for near-circular and arbitrary separation relative motion. • The configuration design equation for large-range relative motion is formulated. • Reconfiguration and collision avoidance control for large-range distributed space systems are realized autonomously. [ABSTRACT FROM AUTHOR]

This paper proposes a novel electromagnetic separation system for the small spherical satellite Q-SAT, which is developed by Tsinghua University. Q-SAT was successfully launched on August 6, 2020 and works well now. Q-SAT requires the separation system no pollution, no impact, multi-point synchronous release and repeatable test. In order to meet these requirements, an electromagnetic separation method is proposed and the system is designed. The system consists of bearing structure, a pair of locking and release mechanism, four ejection mechanism and a drive power. The scheme of separation system can be summarized as "two locks, four ejections". The key to system is the lock and release mechanism using ball-lock principle driven by electromagnets. The analytical relationship of the mechanism driving force is given. How to balance the contradiction between lock and release, which is a problem bring by the system, is discussed. The prototype is made. The verification tests are fully carried out including function, vibration, shock, coldness and velocity measurement. The test results show that the system satisfies the launching requirement. The reliability of the system was proven by the successful separation. Systems based on the same electromagnetic separation method could be widely used on other satellites or payloads in the future. • A novel separation method for a small satellite based on electromagnetic mechanism is proposed. • An electromagnetic separation system is developed for the small spherical satellite Q-SAT based on the method. • The separation system has the characteristics of no impact, no pollution, low cost, repeatable test on the ground and high synchronization. • The separation system is fully verified by test and flight. [ABSTRACT FROM AUTHOR]

With the development of satellite cluster and giant constellations in recent years, the orbit design and long-term maintenance of these systems become more difficult with the restriction of fuel consumption and the influence of orbital perturbations. It's necessary to study the analytical model of relative motion which applicable to different inter-satellite ranges. Then, considering the perturbed evolutions of orbital elements, the decomposed relative motion model was proposed by separating the higher-fidelity relative motion model into two different parts, which reflect the circling motion around chief, and their overall motion relative to chief respectively. Eccentricity expansions had been carried out to provide a time-explicit analytical solution of the decomposed model. Based on this first order time-explicit analytical solution, the bound relative motions with different inter-satellite ranges were studied in Kepler orbit. The initial condition calculation formula for general chief centered circular relative motion was derived at the same time. The accuracy of first order analytical solution of the decomposed model was estimated both theoretically and numerically, to show the effect of inter-satellite range and chief's eccentricity. At last, the first order analytical solution using mean orbital elements was compared with high precision numerical solution under perturbations to validate it's effective in long-term stability analysis. • Decomposed relative motion model was proposed using classical orbital elements. • Analytical solution of decomposed model was provided without distance limitation. • Initial condition formula for general chief centered circular motion was derived. • Theoretical accuracy estimation equation directly reflects the key effect factors. • Simulations validate the effectiveness of analytical solution under perturbations. [ABSTRACT FROM AUTHOR]

An intelligent, person-following nanosatellite is under development for in-cabin astronaut assistance in the China Space Station. It is named the Intelligent Formation Personal Satellite (IFPS). The satellite weighs 2.0 kg and is shaped as a sphere of diameter 230 mm. Fans and MEMS flywheels are used for its position and attitude maneuvering as inside the Space Station cabin it is a weightless and standard atmospheric pressure environment. The RGB-D camera and IMU based visual-inertial SLAM method is used to support its localization and navigation. The on-board information processing and computing hardware primarily consists of an embedded AI microprocessor and an FPGA. The satellite is designed to fly autonomously and follow the designated astronaut to offer immediate assistance. Thus, efficient and robust astronaut visual tracking is the most important prerequisite for supporting its basic person-following operating mode. We achieved this by further improving our previously proposed tracking algorithm that consists of a deep convolution neural network (DCNN)-based detection module and a probabilistic-model-based tracking module. The DCNN in the detection module was further improved through optimizations of lightweight network architecture design, parameters model compression and inference acceleration. While maintaining the originally high detection accuracy, the DCNN was optimized significantly in terms of memory, computation and power consumption to quite meet the engineering constraints in the development of IFPS. The complete pipeline of the astronaut visual tracking algorithm was also designed and implemented in the embedded AI microprocessor for online tracking application. Experimental results demonstrated the effectiveness of the proposed efficient and robust astronaut detection and tracking algorithm. • An in-cabin nanosatellite for astronaut assistance is designed. Deep-learning-based astronaut visual tracking algorithm is proposed for supporting its astronaut-following operating mode. • An efficient DCNN was developed through lightweight network architecture design, parameters model compression and inference acceleration. • The DCNN provides high detection accuracy, and its memory requirements, computation burden and power consumption are engineering affordable. • The deep-learning-based tracking algorithm runs efficiently on embedded hardware. Robust detection and tracking performance is verified by experiments. [ABSTRACT FROM AUTHOR]

Q-SAT, a small spherical satellite designed and developed by Tsinghua university, was successfully launched on Aug 6, 2020 at Jiuquan satellite launch center in China. The mission of Q-Sat is the joint measurement of long-wavelength Earth gravity field and upper atmosphere density, which is of great significance for improving spacecraft orbit prediction and promoting the development of satellite gravity measurement technology. The satellite is designed in spherical shape innovatively and deploys high dynamic dual-frequency GPS as the main payload. On-orbit data shows that the atmospheric density detection precision is in the order of 10−14kg/m3 and the gravity recovery precision achieves the level of 30 orders. All mission success criteria and objectives were achieved. This paper presents the design, space mission, key technologies, science capabilities and preliminary results of the satellite. • Q-SAT, a spherical satellite for the joint measurement of long-wavelength Earth gravity field and upper atmosphere density, was launched in 2020. • An electromagnetic separation system with point connection-release structure is designed for the Q-Sat. • As secondary payload a powerful GPU module is implemented to validate artificial intelligence computing capabilities in orbit. • Modifications for the Jacchia-Roberts atmospheric density model are proposed in the paper. [ABSTRACT FROM AUTHOR]

Abstract Electronics installed in modern complex systems, like automotive and aircraft systems, are mainly subjected to mechanical vibrations in applications. The interconnection structure, as the mechanical fixing and electrical interconnection between the electrical device and the printed circuit board, is a key part in electronics. Presently, there are few researching articles on reliability of QFP interconnect structure under random vibration. In this paper, eight specimens of QFP100 with the self-testing real-time circuit have been tested to failure by subjecting to a narrow-band random vibration at its first natural frequency. The charging time of external capacitor in the self-test circuit has been extracted as the monitoring signal. Failure modes of interconnect structure have been analyzed in statistics. Then, the relationship between the monitoring signal and interconnect failure has been investigated. Finally, failure data, conforms to a two-parameter Weibull distribution, has been used to establish the life distribution of interconnect structure. Results show that there are four failure modes and three failure places in interconnect structure under random vibration, which is consistent with simulation results by FEM, and the charging time can characterize the failure and failure process of interconnect structure well. Highlights • QFP package with the self-testing real-time circuit has been tested to failure under narrow-band random vibration. • The charging time of external capacitor in the self-testing circuit has been extracted as the monitoring signal. • The charging time can characterize the failure and failure process of interconnect structure well. [ABSTRACT FROM AUTHOR]