Transfer from Lunar Gateway to Sun-Earth Halo Orbits Using Solar Sails

被引:4
|
作者
Chujo, Toshihiro [1 ]
Takao, Yuki [2 ]
Oshima, Kenta [3 ]
机构
[1] Tokyo Inst Technol, Tokyo 1528552, Japan
[2] Japan Aerosp Explorat Agcy, Sagamihara, Kanagawa 2525210, Japan
[3] Hiroshima Inst Technol, Hiroshima 7315193, Japan
基金
日本学术振兴会;
关键词
Solar Sail; Halo Orbit; Near Rectilinear Halo Orbit; Lagrangian Point; Low Thrust Trajectory; Trajectory Optimization; Solar Electric Propulsion; Deep Space Gateway; Solar Radiation Pressure;
D O I
10.2514/1.A35559
中图分类号
V [航空、航天];
学科分类号
08 ; 0825 ;
摘要
To extend the usability of solar sails in the sun-Earth-moon system, we analyze the transfer trajectories from the 9:2 Earth-moon near-rectilinear halo orbit (NRHO) to halo orbits around the sun-Earth L1 and L2 points under the assumption of a future mission for a solar sail spacecraft equipped with a solar electric propulsion (SEP) system deployed from the Lunar Orbital Platform-Gateway. The dynamics are modeled using the bicircular restricted four-body problem, where the gravitational forces from the sun, Earth, and moon as well as solar radiation pressure (SRP) are considered. We propose a trajectory design method that utilizes both SRP and SEP. The method consists of initial guess generation and optimization steps. The initial guess generation comprises the forward propagation of the escape trajectory from the NRHO, the backward propagation of the stable manifold of the target halo orbits, and their apoapsis patching process. Optimization is conducted to minimize propellant consumption by effectively controlling SRP. We perform optimizations with various parameters, namely, the sail area-to-mass ratio (A/m), specifications of SEP, target sun-Earth halo orbit, and departure Delta V direction. The results validate the proposed trajectory design method and verify that solar sail acceleration can reduce the necessary amount of propellant, which indicates that such missions can be realized by small CubeSats.
引用
收藏
页码:1527 / 1540
页数:14
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