Under the grand goal of “Digital China”, the current development mode of spacecraft in China is upgrading from document driven to model driven. Model based systems engineering(MBSE) is the core method guiding this development mode upgrading. One of the core contents of MBSE is the construction and operation of digital twin model. Spacecraft digital twin model consists of multiscale models, such as spacecraft system level digital twin model, subsystem level digital twin model and component level digital twin model. At the same time, the spacecraft digital twin model is also composed of multidimensional models, such as structure model, software model, circuit model and so on. How to integrate multiscale model and multidimensional model into a whole digital twin model according to certain norms has become a key research content in the field of MBSE. A universal construction method of spacecraft digital twin model is presented in this paper, and a practical scheme based on the principle of operating system is also put forward for the design architecture and principle of the general operation platform of spacecraft digital twin model.

The sensor resources scheduling of LEO constellation is a mission planning problem with complex and multiple constraints. In order to implement the real time scheduling optimization of multiple maneuvering targets’ full process tracking and monitoring tasks, a mission planning model that is easy to quickly solve onboard is proposed. Based on the detailed modeling of the autonomous mission planning for LEO constellation sensor scheduling, the original problem is transformed into a simple 0 1 linear integer programming model with various constraints and optimization objectives. The proposed method is validated through a multiple targets tracking simulation scenario, with all targets tracked and monitored throughout the entire process.

With the increasing demand for satellite earth observation, the importance of missing planning for imaging satellites has become prominent. Currently, the autonomous task planning algorithm for imaging satellites is limited by the satellite’s own computational and storage capabilities, and its engineering application is still in the developmental stage. Imaging mission planning mainly relies on manual expertise from the ground. In this paper, based on the different motion characteristics of high speed and low speed targets in satellite observation, combined with the satellite platform’s attitude maneuvering performance, an imaging satellite adaptive observation planning algorithm based on orbit extrapolation is proposed. By continuously predicting changes in the target trajectory using Lagrange interpolation to construct an optimal objective function,intermittent pure pursuit is computed, and adaptive observations are implemented. This method utilizes the ground system to compensate for the limitations of onboard computation and storage capabilities, achieving rapid target tracking and trajectory prediction, improving response timeliness, and extending satellite imaging time.

The space solar power station is an ultralarge and lightweight space structure, due to the high flexibility, which may result in a notable spillover when treated as the classical modal based dynamics equations with artificial modal truncation. In response to this phenomenon, when modeling the flexible spacecraft dynamics, a continuous spatial domain method is used, to study the asymptotic behavior of the coupled rigid body and flexible system, to determine the dominated dynamical behavior and to provide a guiding principle of a reasonable reduced model for the controller design. The space solar power satellite is approximated by a planar nonlinear free-free beam with global large overall rigid body motions and local elastic vibration with finite extension and bending, to depict the strain hardening exactly. By Hamilton’s principle, a system of nonlinear dispersive equations in the quotient function spaces corresponding to an affine action of the rigid body motion which is described by a floating reference frame is obtained. Then, with the nonlinear dispersive equations, relative equilibrium and their linear stability, as well as the relation between the mode solutions or the dispersive solutions and structual parameters are analyzed in the case of small deformation. Nonlinear dynamical behavior including strain hardening, nonlinear normal modes and solitons are analyzed primarily in the finite deformation case. The principle of adopting a promising reduced model is summarized according to the structural parameters and initial conditions. Numerical simulations illustrate the validity of the key analytical results.

For the control design problems of spacecraft flight in the atmosphere, the aerodynamic angles are directly related to aerodynamic forces and aerodynamic moments, so the aerodynamic angle feedback is generally adopted. When studying the closed loop performance such as stability, it is necessary to use the aerodynamic angular kinematics formula. However, it is difficult to find the derivation method of aerodynamic angular kinematics, so there is a lack of flexibility in the definition of different coordinate systems in application. The kinematic forms of aerodynamic angles are complex, and their derivation is difficult. In this paper, the derivation method is put forward, the kinematic equations of aerodynamic angles are given, and their open loop performance is analyzed, providing the basis for control design. The derivation method and the equations of aerodynamic angular kinematics, and the methods and results of open loop performance analysis proposed in this paper can be applied to the control design and analysis of aerodynamic entry of target stars with atmosphere, thus providing theoretical support for them.

A finite thrust orbit and attitude control scheme is proposed to solve the problem of unadjustable orbit control thrust with large misalignment for a spacecraft with deviated centroid. The proposed orbit control scheme can control multiple orbit elements at the same time. The problem of unadjustable orbit control thrust is solved by taking the thrust angle as the control input of orbit control. When designing the attitude control strategy to track the desired attitude of orbit control, the large interference torque caused by the thrust misalignment is considered. To solve the problem of large disturbance, the traditional asymptotically stable attitude control law is improved. The coupling relationship of three axis attitude is used to counteract the influence of interference torque. The effectiveness of the orbit and attitude control strategy is verified by numerical simulation.

A novel active disturbance rejection control method is proposed to satisfy the high requirement of speed, accuracy and robustness of attitude control system for kinetic energy interceptor in space interception mission. Firstly, the attitude dynamics model of the kinetic energy interceptor is established and converted into the second order state equation of attitude tracking error. Secondly, based on the idea of small error, large gain and large error, small gain, a new nonlinear function is designed. Then, the tracking differentiator, the extended state observer and the nonlinear controller are respectively designed based on the nonlinear function. By taking the prescribed performance function that converges at a specified time as the expected signal of the tracking differentiator, the attitude error can converge to the expected value according to the prescribed performance function, and the convergence time can be predetermined. Finally, the pseudo rate pulse modulator is used to pulse the continuous control torque. Numerical simulations demonstrate that the proposed control method exhibits precise attitude control and robust resilience against external disturbances.

A method for designing simulation platform of the spacecraft Guidance Navigation & Control (GNC) system is proposed. During the in-space assembly and long term operation of the Chinese Space Station, complex flight missions are involved, including multi-spacecraft individual on-orbit flight, space rendezvous and docking, multi-spacecraft combination fusion control, separation process control and spacecraft re-entry. The integrate and realistic flight simulation is necessary and vital. Based on time synchronization, real-time data exchange, distributed and integrated simulation architecture，a universal multispacecraft collaborative Hardware-in-the-Loop (HIL) simulation platform of GNC system is proposed, on which integrate flight simulation with high fidelity is possible to be achieved. The platform consists of several universal simulators, which are highly flexible and expandable, and each can be configured as any specific spacecraft and perform the spacecraft’s whole mission process flight simulation independently. The platform synchronizes time via IRIG-B code and exchanges dynamics data on 1553B bus in real-time, and multi-spacecraft co-simulation can be implemented after coordination accomplished among simulators. This simulation platform has been successfully applied in a spacecraft flight control drilling program.

With the maturation and application of global satellite constellation flying technology, solving relative motion through the fundamental equations of satellite motion and designing various formation configurations have become a development trend. This paper establishes a relative motion model based on deviations in relative orbital elements, first theoretically derives the relative motion equations suitable for high eccentricity elliptical orbits under second order approximation conditions, and further analyzes the accuracy of the second order terms. Additionally, under the two body assumption, the true anomaly difference at perigee is used instead of the mean anomaly difference as the input for formation configuration design. Various typical formation configurations are designed and analyzed under different initial orbital element difference conditions. Finally, simulations are conducted to verify the correctness of the formation configuration design and analyze its applicability from the two body assumption to scenarios involving perturbations，the results show that the formation configuration design and the relative motion equations suitable for high eccentricity elliptical orbits are correct and effective.

The future propulsion system needs to adapt to large acceleration environment and have the refueling capacity with large filling volume. Therefore, it is very important for the overall design of aircraft to research the propellant management characteristic of spacecraft tank at a large acceleration environment. In this paper, the limit management performance of a vane type surface tension tank is studied by numerical methods. The gas liquid distribution in the tank and the liquid level curves of the propellant are analyzed. The diversion capacity of parallel vanes under a large acceleration is calculated, and the influence of vane width and clearance distance on the diversion performance is studied. The numerical simulation results show that increasing the vane width and decreasing the clearance distance between the vane and tank shell can improve the diversion effect at 10-2 g acceleration. The research results of this paper can provide theoretical reference for the optimal design of vane type surface tension tanks.