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A novel guidance law that utilizes optimal control theory in combination with linear time-varying state-transition matrices to calculate the optimal path for a chaser spacecraft from its initial point to a desired final point and a desired amount of time is proposed in this paper. The optimal guidance law is formulated using more accurate second-order integral form of Gauss Variational Equations to map control accelerations to the dynamics in terms of relative orbital elements. When comparing the second-order form against the traditional Gauss Variational Equations, the guidance law results in a 15% reduction in control demand. In addition, this paper presents a new method that allows for the closed-loop testing of guidance, navigation and control systems with realistic software models of GPS receivers, RADAR systems for on-board relative navigation, and ground-station up-links of the target states. Specifically, utilizing the signal-to-noise ratio of various sensors, the errors that arise due to hardware noise and/or environmental factors can be modeled accurately based on the communication link. The guidance law is tested using the newly developed software-in-the-loop test-bed for a far-range formation flight scenario around the non-cooperative Alouette-2 spacecraft as well as on an arbitrary defined highly elliptical orbit.

This paper presents a guidance solution of relative motion between two spacecraft using relative classical orbital elements for on-board implementation purposes. The solution is obtained by propagating the relative orbital elements forward in time using a newly formulated state-transition matrix, while taking into account gravitational field up to the fifth harmonic, third-body effects up to the fourth order and drag, then calculating the relative motion in the local-vertical-local-horizontal reference frame at each time-step. Specifically, utilizing Jacobian matrices evaluated at the target spacecraft’s initial orbital elements, the solution proposed in this paper requires only a single matrix multiplication with the initial orbital elements and the desired time to propagate relative orbital elements forward in time. The new solution is shown to accurately describe the relative motion when compared with a numerical simulator, yielding errors on the order of meters for separation distances on the order of thousands of meters. Additionally, the solution maintained accurate tracking performance when used within a back-propagation, or terminal-point, guidance law.

Disturbance gravity is an important error source that affects the guidance accuracy of spacecraft, and the improvement of guidance accuracy is of great significance for improving the damage efficiency of spacecraft. Although this problem has become a hot topic in current research, the influence mechanism of disturbing gravity on guidance accuracy and its quantitative description still needs to be further summarized in typical guidance modes. In this paper, the influence mechanism of the disturbance gravity of the launch area on the guidance accuracy of the spacecraft in the perturbation guidance and closed-loop guidance modes is systematically analyzed, and the fall-point deviation of the spacecraft under the disturbance gravity of the launch area in two typical guidance modes is quantitatively calculated by simulation. The results show that disturbing gravity has a significant influence on the guidance accuracy of spacecraft, and the influence degree is different under different guidance modes. The research in this paper verifies the necessity of an in-depth study of disturbing gravitational fields to improve the guidance accuracy of spacecraft and lays a foundation for proposing effective compensation measures to improve the guidance accuracy of spacecraft.

To ensure the successful completion of the manned spacecraft flight missions during China Space Station phase, an analysis of the mission characteristics was conducted. In response to the challenges posed by extended in-orbit flight duration, complex and variable external thermal flux, docking at different ports of the space station, new emergency rescue missions, high precision landing requirements, and efficient earth-to-space transportation, the following coping strategies were proposed, including: implementing long-life design to accommodate extended missions, actively temperature controlling and passive thermal insulation for temperature-sensitive equipment, developing all-directional rendezvous and docking designs, adding emergency rescue spacecraft on the ground, upgrading the re-entry guidance and control methods, and adopting autonomous rapid rendezvous and rapid return technologies. These strategies have been successfully implemented during the space station phase, ensuring the successful completion of the Shenzhou-12 to Shenzhou-18 crewed spacecraft missions. The results demonstrate that the analysis of the mission characteristics was accurate, and the proposed countermeasures are reasonable and feasible, effectively ensuring the safety and reliability of the crewed spaceflight missions for China Space Station.

To probe the possibility of water sources on Mars and research the surface environment of Mars, a Martian spacecraft was designed in detail in this study. Camera PIXL was selected as the payload. The service platform includes an orbit control system, communication system, thermal control system, main computer, energy system, and structure system, which was designed to support the operation of the spacecraft to complete required space missions. The consequence of this research furnishes some guidance for the design of the Martian spacecraft, which discovers water sources on Mars.

This study investigates the modelling and attitude behaviors of a partial-spin spacecraft following mass separation. A dimensionless dynamic model of the spacecraft’s attitude motion is developed and validated against common dynamic simulation software. The effects of key configuration parameters—separated mass, eccentricity, separation arm length, and platform mass—on attitude stability are analyzed, revealing their influence on spin axis deviation angle. These findings provide guidance for the dynamic analysis and configuration design of such spacecraft.

The rapid developments of artificial intelligence in the last decade are influencing aerospace engineering to a great extent and research in this context is proliferating. We share our observations on the recent developments in the area of spacecraft guidance dynamics and control, giving selected examples on success stories that have been motivated by mission designs. Our focus is on evolutionary optimisation, tree searches and machine learning, including deep learning and reinforcement learning as the key technologies and drivers for current and future research in the field. From a high-level perspective, we survey various scenarios for which these approaches have been successfully applied or are under strong scientific investigation. Whenever possible, we highlight the relations and synergies that can be obtained by combining different techniques and projects towards future domains for which newly emerging artificial intelligence techniques are expected to become game changers.

The autonomous optimal guidance and control techniques serve as critical foundations for on-orbit failed spacecraft detumbling. Recently, the plumes have become an efficient detumbling medium, which can avoid direct contact to improve operation safety. However, the traditional molecular plume may lead to unaffordable fuel consumption, so it is necessary to propose a novel strategy to solve this problem. The Hall effect thruster is a common high specific impulse engine, which can generate high-speed plasma plume for contactless detumbling. The traditional plume models present difficulties for practical implementation due to the contradiction between the requirement of real-time computation and the limited computing capability of the spaceborne computer. To address the challenge of calculating the impact force, a fully analytical plasma plume model is proposed to substantially enhance the computing efficiency without loss of accuracy. A guidance law for optimal detumbling is proposed to mitigate the spin angular velocity while stabilizing nutation under multiple complex constraints. Moreover, in order to facilitate the development of the detumbling control, a terminal non-singular sliding mode controller based on the saturation function (TSM-sat) is designed, which can converge in finite time and avoid the chattering phenomenon. The numerical results indicate that the established plasma plume model can significantly improve the computing efficiency and the optimal guidance law can stabilize the target within a reasonable time.

The growing interest in Artificial Intelligence is pervading several domains of technology and robotics research. Only recently has the space community started to investigate deep learning methods and artificial neural networks for space systems. This paper aims at introducing the most relevant characteristics of these topics for spacecraft dynamics control, guidance and navigation. The most common artificial neural network architectures and the associated training methods are examined, trying to highlight the advantages and disadvantages of their employment for specific problems. In particular, the applications of artificial neural networks to system identification, control synthesis and optical navigation are reviewed and compared using quantitative and qualitative metrics. This overview presents the end-to-end deep learning frameworks for spacecraft guidance, navigation and control together with the hybrid methods in which the neural techniques are coupled with traditional algorithms to enhance their performance levels.

With the rapid growth of the number of spacecraft tasks and the complexity of the system, spacecraft software development is faced with the challenges of high-quality and rapid delivery, increased complexity, and frequent changes in requirements. The reuse rate of software assets needs to be improved urgently. In order to realize cross-platform and cross-project reuse of software design, this paper studies the design of electric propulsion software architecture. Carry out requirements decomposition and subsystem division of electric propulsion systems for different platforms, extract common requirements and variability requirements at the system level and subsystem level based on the Feature-Oriented Domain Analysis method, and establish a requirements model. Based on the requirements model, an electric propulsion software architecture integrating “event-driven + components” is proposed, and the classic event-driven model of “publish-subscribe” is adaptively improved. The architecture has been instantiated and applied to 16 projects on 4 platforms, and the reuse rate of core assets has reached 70%. This paper has certain guidance and reference significance for spacecraft software asset reuse and platform construction.