Small satellite architecture optimization: electric propulsion moon imaging mission

Abstract

This study underlies small satellite architecture optimization by using existing electric propulsion systems for the Moon missions. The estimated objective is panoramic imaging of the Moon accompanied with future in-situ applications. Edelbaum’s low thrust trajectory transfer with optimal control theory is used to calculate the required ΔV. During the journey, 1.5h eclipse duration effects the solar array design. The optimized xenon propellant density and pressure are 1350 kgm3 and 8.3 MPa within 300K. Two types of optimization process revealed based on hexagonal SC architecture. The iterative method with LEO departured ion thruster has 23 mN with minimum 213 kg total mass. Corresponding SC volume is 0.70 m3, propellant mass is 64 kg. This scenario cost $108.5M and takes 980 days. Same thruster level for GEO departure case takes 880 days with 58 kg xenon gas. The total cost reduces $2.5M. For HALL engine design, LEO departure case needs 0.8 m3, 247 kg SC including 82 kg xenon. 77 mN thrust operates 208 days towards the Moon that ends up with $121M total cost. If the SC to be launched from GEO, flight time reduces 45 days by consuming 65 kg propellant. Total SC mass, volume and power values are 230 kg, 0.71 m3 and 1351W which cost $115M. Results are compared with previous Moon or electric propulsion missions such as SMART-1, LADEE, Clementine and Hayabusa. For future applications of small satellites, innovative concepts are envisioned for in-space, Earth-independent exploration and space education. 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. .