Search results
1 – 5 of 5Hongwei Yang, Yu Jiang and Hexi Baoyin
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…
Abstract
Purpose
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.
Design/methodology/approach
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.
Findings
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.
Research limitations/implications
The proposed control strategy is limited for the orbital altitude where the J2 perturbation is dominant.
Practical implications
The proposed effective method is applicable for the engineers planning an orbital control strategy of deploying satellites of a constellation.
Originality/value
The new control strategy can realize utilization of J2 perturbation and an accurate deployment, simultaneously. Further, this paper provides practical help for satellite engineers.
Details
Keywords
Jin Jin, Hexi Baoyin and Junfeng Li
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…
Abstract
Purpose
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.
Design/methodology/approach
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.
Findings
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.
Originality/value
The proposed ADCS scheme can be a reference for the future Micro‐satellite programs which share the similar configuration.
Details
Keywords
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…
Abstract
Purpose
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.
Design/methodology/approach
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.
Findings
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.
Practical implications
Such trajectories can be used by China’s future manned lunar missions. Spacecraft capability and ground station distribution shall adjust accordingly.
Originality/value
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.
Details
Keywords
Xiaobin Lian, Jiafu Liu, Chuang Wang, Tiger Yuan and Naigang Cui
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…
Abstract
Purpose
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.
Design/methodology/approach
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.
Findings
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.
Practical implications
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.
Originality/value
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.
Details
Keywords
Weijia Lu, Chengxi Zhang, Fei Liu, Jin Wu, Jihe Wang and Lining Tan
This paper aims to investigate the relative translational control for multiple spacecraft formation flying. This paper proposes an engineering-friendly, structurally simple, fast…
Abstract
Purpose
This paper aims to investigate the relative translational control for multiple spacecraft formation flying. This paper proposes an engineering-friendly, structurally simple, fast and model-free control algorithm.
Design/methodology/approach
This paper proposes a tanh-type self-learning control (SLC) approach with variable learning intensity (VLI) to guarantee global convergence of the tracking error. This control algorithm utilizes the controller's previous control information in addition to the current system state information and avoids complicating the control structure.
Findings
The proposed approach is model-free and can obtain the control law without accurate modeling of the spacecraft formation dynamics. The tanh function can tune the magnitude of the learning intensity to reduce the control saturation behavior when the tracking error is large.
Practical implications
This algorithm is model-free, robust to perturbations such as disturbances and system uncertainties, and has a simple structure that is very conducive to engineering applications.
Originality/value
This paper verified the control performance of the proposed algorithm for spacecraft formation in the presence of disturbances by simulation and achieved high steady-state accuracy and response speed over comparisons.
Details