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

DOI Web Site 被引用文献2件 参考文献11件
  • Toshihiro Chujo
    Tokyo Institute of Technology, Tokyo 152-8552, Japan
  • Yuki Takao
    Japan Aerospace Exploration Agency, Kanagawa 252-5210, Japan
  • Kenta Oshima
    Hiroshima Institute of Technology, Hiroshima 731-5193, Japan

説明

<jats:p> 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 ([Formula: see text]), specifications of SEP, target sun–Earth halo orbit, and departure [Formula: see text] 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. </jats:p>

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