Analysis Of Reachable Domain Of Spacecraft And Accurate Orbit Control Of Relative Motion With Perturbation

Orbit maneuver exhibits significant application in rendezvous and docking,orbit interception,formation flight etc. This paper completes some tasks in the following aspects: Spacecraft's absolute and relative reachable domain in the case of specified orbit transfer ability are analyzed; A kind of accurate guidance method for elliptical reference orbit with perturbation is given; Continuous thrust guidance is realized based on impulse transition principle.The foundation of relative motion model: Relative motion equation described by the state transition matrix is introduced to describe relative motion in elliptical orbit. Nonlinear relative motion equation of elliptical orbit with J₂ perturbation and atmospheric drag perturbation is established by the use of Lagrangian equation. Motion equation established in this paper could accurately describe relative motion between two spacecrafts, compared to the relative motion state in the orbit system of tracker from inertial system.The analysis of spacecraft's absolute reachable domain in the case of specified orbit transfer ability: The Gauss formula and the Gauss formula without odd point are introduced. The direction and position of thrust are analyzed using the Gauss formula when orbit elements of spacecraft take the extrame values. Orbit elements extrame values of spacecraft affected by impulse or continuous thrust and the thrust direction and position are confirmed by numerical method. The theoretical analysis results are proved to be correct by comparison to the numerical method results.The analysis of relative reachable domain in the case of specified orbit transfer ability: Impulse direction is analyzed with the longest distance between spacecraft and space platform, using relative motion equation described by the state transition matrix. The longest distance between spacecraft and space platform and the thrust direction at the moment affected by impulse or continuous thrust are confirmed by numerical method in given time. The analytical method is proved to be correct by comparison to numerical method results.Accurate guidance for elliptical reference orbit with perturbation: An iterative operation is given to optimize the initial velocity impulse which can achieve accurate rendezvous. This guidance low founds an orbit prediction method with nonlinear relative motion equation with J2 perturbation and atmospheric drag perturbation. This orbit prediction method is used to estimate the terminal error. Double-pulse guidance method for fuel optimal is obtained by optimizing transfer time. Singular conditions in solving double-pulse matrix is analyzed, and the guidance method is given in singular point. Continuous thrust guidance strategy is realized based on impulse transition principle. The method is proved to be feasible by numerical simulation.