Optimization of space missions to near-earth asteroids using solar electric propulsion systems

The optimal trajectories of the spacecraft (SC) with electric propulsion system (EPS) to the near-earth asteroids are considered. Typically, within frame of the problem, the final SC mass is maximized or the active propellant mass is minimized using appropriate choosing of EPS thrust steering program and free trajectory parameters (launch date, magnitude and direction of departure velocity, etc.). Practically, however, more important is the problem of SC useful mass (i.e. mass of the payload and onboard service systems providing the payload operations) maximization. Last problem coincides with the problem of minimization of the total mass of EPS with the propellant storage system and the part of power supply system mass using for EPS operating. Maximum principle is used to derive necessary optimality conditions of thrust vector direction, EPS operation cycles, initial EPS thrust, EPS specific impulse, and free trajectory parameters. The boundary value problem of the maximum principle is solved using smoothing of discontinuous functions and homotopic approach to reduce the boundary value problem to the initial value problem. Both direct and round-trip trajectories are considered. The samples of the optimal trajectories to the near-earth asteroid 2003 GA are presented. © 2018 Author(s).