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This paper aims to provide a new method to design a fuel efficient control strategy such as J₂ perturbation for deploying a constellation into a specified configuration. The nonspherical perturbation, mainly J₂ perturbation, is the dominant perturbation for low-Earth-orbit (LEO) satellites of a constellation. This perturbation can be utilized in the control strategy to lower fuel consumption enormously.

The relationship of the coupled variables, the relative right ascension of ascending node (RAAN) and the relative phase (RP), are analyzed. First-order approximation expressions of the relative RAAN (RRAAN) and the relative phase (RP) with respect to the semimajor axis are derived. According to the Gauss’ variational equations, the reduced explicit functions of these variables in regard to each active control are established. Based on these functions, control strategy design methods, including the preliminary planning and iterative corrections, are proposed. The numerical simulation is carried out to verify the proposed method.

The results indicate that the constellation can be deployed accurately about the semimajor axis, the RRAAN and the relative phase (RP) by the developed fuel efficient control strategy.

The proposed control strategy is limited for the orbital altitude where the J₂ perturbation is dominant.

The proposed effective method is applicable for the engineers planning an orbital control strategy of deploying satellites of a constellation.

The new control strategy can realize utilization of J₂ perturbation and an accurate deployment, simultaneously. Further, this paper provides practical help for satellite engineers.

The purpose of this paper is to propose an attitude determination and control scheme for a low‐cost Micro‐satellite with defective inertia. Restricted by the payload design, the z‐axis inertia of this satellite is larger than the x and y axes, which is unstable for natural attitude dynamics.

An original operation mode is designed to avoid z axis from long‐time pointing to the sun during damping, which avoids some unexpected damage. In attitude determination design, EKF and UKF algorithms are compared on estimation accuracy, convergence time and computation complexity in attitude estimation design, which is referred to determine the final estimation scheme. A DSP‐based hardware solution is achieved and a semi‐physical testing and simulation system is built.

Simulation results show the 3‐axis stable mode can be built with the proposed scheme, and the unprotected facet of the satellite can be kept away from long‐time pointing to the sun.

The proposed ADCS scheme can be a reference for the future Micro‐satellite programs which share the similar configuration.

The purpose of this paper is to design free return trajectories launching at lower-latitude launch site Wenchang and landing at relatively high-latitude landing site Siziwang Banner tailored to human lunar missions for China, and in general demonstrate the feasibility of high-latitude landings with acceptable entry range.

Free return trajectories satisfying all basic constraints were generated directly by a high-fidelity model with multiple differential corrections. Suitable initial assumptions, control parameters, constraints and stopping conditions were set. Method was developed to automatically converge unlimited trajectories accurately to the same constraints, and their characteristics affected by the ephemeris were analyzed.

Launching into lower Earth inclination plus high-latitude landing with acceptable entry range requires asymmetric trajectories with high inclination Earth entry only from the south. Periodic trends of parameters at launch, injection and entry were found and analyzed. Nominal trajectory covering phases from launch to landing for China human moon flight with minimum entry range were designed.

Such trajectories can be used by China’s future manned lunar missions. Spacecraft capability and ground station distribution shall adjust accordingly.

Previous studies mainly concentrated on symmetric free returns using low-fidelity models first. This paper investigates asymmetric free returns skipping simplified gravity model approximation to simultaneously achieve high-latitude landing and acceptable entry range, and accurate automated generation of feasible trajectories daily across 19-year lunar nodal cycle within every monthly launch window without trial and error to reflect the actual effect by the ephemeris only. Others include landing accurately by controlling entry direction and range (and altitude), minimizing entry range and designing an effective scheme of differential correction for full convergence.

The purpose of this paper is to resolve complex nonlinear dynamical problems of the pitching axis of solar sail in body coordinate system compared with inertial coordinate system. And saturation condition of controlled torque of vane in the orbit with big eccentricity ration, uncertainty and external disturbance under complex space background are considered.

The pitch dynamics of the sailcraft in the prescribed elliptic earth orbits is established considering the torques by the control vanes, gravity gradient and offset between the center-of-mass (cm) and center-of-pressure (cp). The maximal torques afforded by the control vanes are numerically determined for the sailcraft at any position with any pitch angle, which will be used as the restriction of the attitude control torques. The finite/infinite time adaptive sliding mode saturation controller and Bang–Bang–Radial Basis Function (RBF) controller are designed for the sailcraft with restricted attitude control torques. The model uncertainty and the input error (the error between real input and ideal control law input) are solved using the RBF network.

The finite true anomaly adaptive sliding mode saturation controller performed better than the other two controllers by comparing the numerical results in the paper. The control torque saturation, the model uncertainty and the external disturbance were also effectively solved using the infinite and finite time adaptive sliding mode saturation controllers by analyzing the numerical simulations. The stabilization of the pitch motion was accomplished within half orbit period.

The complex accurate dynamics can be approximated using the RBF network. The controllers can be applied to stabilization of spacecraft attitude dynamics with uncertainties in complex space environment.

Advanced control method is used in this paper; saturation of controlled torque of vane is resolved when the orbit with big eccentricity ration is considered and uncertainty and external disturbance under complex space background are settled. Moreover, complex and accurate nonlinear dynamical model of pitching axis of solar sail in body coordinate system compared with inertial coordinate system is provided.