Geostationary debris mitigation using minimum time solar sail trajectories with eclipse constraints

Abstract

SummaryMinimum time solar sailing trajectories are introduced using a combination of indirect and direct optimal control techniques. Here, large‐scale, multiphase optimal control problems are solved using a pseudospectral collocation technique applied to an orbital debris mitigation concept. These solutions are obtained for realistic sail dimensions, producing multirevolution, Earth‐centered trajectories while accounting for uncontrolled spacecraft dynamics in the eclipse regions. Specifically, minimum time solutions for orbit transfer and phasing maneuvers are obtained using only solar radiation pressure for propulsion and control. First, an optimal primer vector steering history is obtained through numerical approximation. Locally optimal, closed‐form solutions are then implemented based on the primer vector direction, resulting in minimum time satisfaction of desired terminal orbital conditions. In addition, a novel solution strategy is presented for multiphase optimal control problems characterized by uncontrollable dynamics with flexible phase boundaries. The maneuvers presented will be shown to enable efficient orbital debris mitigation for large‐scale debris in geostationary orbits.

A mesh refinement method for low‐thrust orbit transfer problems

“…Orbit transfer with a set of continuous low‐thrust burns has been researched for decades 5–8 . This problem is primarily resolved by simplification‐based techniques, for example, Edelbaum 9 developed a constant acceleration model and generated a series of suboptimal analytical solutions to the transfer between two non‐coplanar circular orbits.…”

“…Orbit transfer with a set of continuous low-thrust burns has been researched for decades. [5][6][7][8] This problem is primarily resolved by simplification-based techniques, for example, Edelbaum 9 developed a constant acceleration model and generated a series of suboptimal analytical solutions to the transfer between two non-coplanar circular orbits. Based on Edelbaum's method, Casalino and Colasurdo 10 obtained solutions in view of variable spacecraft mass, specific impulse, and magnitude of thrust, and Kluever 11 considered the effect of shadow and the decrease in spacecraft mass.…”