A field-of-view constrained three-dimensional impact time and angle guidance law is proposed in this paper. The challenging guidance problem formulated in the three-dimensional scenario while considering multiple constraints is addressed via the proposed dual-step joint approach. The guidance process consists of two steps. In the first step, the equiangular spiral shaping trajectory is conducted to achieve the desired impact time while adhering to the field-of-view constraint. In the second step, the field-of-view constrained impact angle guidance law is derived by employing the error dynamics and auxiliary function. The switching point between the guidance laws in the two steps needs to be determined quickly and accurately. The guidance parameters can be obtained efficiently by employing the three-point secant method, solving a nonlinear implicit equality with 2–3 iterations. Numerical simulation examples are provided to verify the performance of the proposed guidance law on different guidance constraints.

With the rapid changes of the flight environment and situation, there will be various unexpected situations while multiple missiles are performing the missions. To fast cope with the various situations in mission executions, the conventional sequential convex programming algorithm and the parallel-based sequential convex programming algorithm for multiple missiles fast trajectory replanning are proposed in this paper. The originally non-convex trajectory optimization problem is reformulated into a series of convex optimization subproblems based on the sequential convex programming method. The conventional sequential convex programming algorithm is developed through linearization, successive convexification, and relaxation techniques to solve the convex optimization subproblems iteratively. However, multiple missiles are related through various cooperative constraints. When the trajectory optimization of multiple missiles is formulated as an optimal control problem to solve, the complexity of the problem will increase dramatically as the number of missiles increases. To alleviate the coupled effect caused by multiple aerodynamically controlled missiles, the parallel-based sequential convex programming algorithm is proposed to solve the trajectory optimization problem for multiple missiles in parallel, reducing the complexity of the trajectory optimization problem and significantly shortening the computation time. Numerical simulations are provided to verify the convergence and effectiveness of the conventional sequential convex programming algorithm and the parallel-based sequential convex programming algorithm to cope with the trajectory optimization problem with various constraints. Furthermore, the optimality and the real-time performance of the proposed algorithms are discussed in comparative simulation examples.

This paper investigates the heliocentric time-optimal rendezvous performance of Sun-facing diffractive solar sails with various deflection angles and acceleration capabilities. Diffractive solar sails, which generate tangential radiation pressure force, are proposed and schematically designed to achieve diverse radiation pressure distributions. The radiation pressure force model and the time-optimal control problem for these innovative Sun-facing diffractive solar sails are established. Utilizing an indirect method and the optimal control law, we explore typical heliocentric rendezvous scenarios to assess the variational trends of transfer time in relation to different deflection angles and acceleration capabilities. The results for Sun-facing diffractive sails in specific rendezvous missions are compared to reflective sails with the same area-to-mass ratio, focusing on transfer trajectory and attitude control. Our findings reveal that diffractive sails exhibit significant advantages over reflective sails, particularly in the context of normal acceleration, paving the way for more efficient space exploration.

This paper proposed a new attitude determination method for low-orbit spacecraft. The attitude prediction accuracy is greatly improved by adding the unmodeled environmental torque to the dynamic equation. Specifically, the environmental torque extraction algorithm based on extended Kalman filter and series extended state observer is introduced, and the unmodeled part of dynamic is identified through the inverse dynamic model. Then, the collected data are analyzed and trained by a backpropagation neural network, resulting in an attitude-torque mapping network with compensation ability. The simulation results show that the proposed feedback attitude prediction algorithm can outperform standard methods and provide a high accurate picture of prediction and reliability with discontinuous measurement.

Real-time guidance is critical for the vertical recovery of rockets. However, traditional sequential convex optimization algorithms suffer from shortcomings in terms of their poor real-time performance. This work focuses on applying the deep learning-based closed-loop guidance algorithm and error propagation analysis for powered landing, thereby significantly improving the real-time performance. First, a controller consisting of two deep neural networks is constructed to map the thrust direction and magnitude of the rocket according to the state variables. Thereafter, the analytical transition relationships between different uncertainty sources and the state propagation error in a single guidance period are analyzed by adopting linear covariance analysis. Finally, the accuracy of the proposed methods is verified via a comparison with the indirect method and Monte Carlo simulations. Compared with the traditional sequential convex optimization algorithm, our method reduces the computation time from 75 ms to less than 1 ms. Therefore, it shows potential for online applications.

Active attitude control of solar sails is required to control the direction of the force generated by Solar Radiation Pressure (SRP). It is desirable to control the attitude through propellant-free means. This paper proposes a new method for attitude control of solar sails: A boom consisting of “smart” structural material can be deformed by the piezoelectric actuator, and Solar Radiation Pressure torque will be generated due to shape variation of sail membrane caused by boom deformation. The method has the advantages of simple structure, small disturbance and small additional load, and is not limited by the size of the solar sail. The case of rendezvous with the Asteroid 2000 SG344 is used to verify the attitude control around the pitch and yaw axes.