IAC-15- B4.8.7 comprehensive study of small satellite moon missions: architecture design, electric propulsion system optimization and cost analysis

Abstract

This study underlies small satellite architecture optimization by using existing electric propulsion systems for the Moon missions. The estimated objective is 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 AV. During the joumey, 1.5h eclipse duration effects the solar array design. The optimized xenon propeliant density and pressure are 1350 kg m3 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, propeliant 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 S2.5M. 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 propeliant. 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.