Stand-alone CubeSat missions to Mars that escape Earth and experience a deep-space cruise require robust primary propulsion systems for orbit manoeuvring and precise trajectory control. Combined chemical{electric propulsion systems enable a hybrid high-thrust{ low-thrust transfer from a high energy Earth orbit to Mars. High-thrust chemical propulsion is used for Earth escape and the low-thrust electric propulsion is used in deep-space cruise, ballisitic capture, and final circularization to an operational orbit about Mars. This work focuses on the performance and design characterization of an iodine-propelled electric propulsion system, concomitant with low-thrust trajectory optimization of the heliocentric transfer and ballistic capture at Mars for a 16U stand-alone CubeSat. A performance model of an inductively coupled miniature ion thruster is implemented to calculate thrust, speci_c impulse, and e_ciencies. A power-constrained low-thrust optimal control problem utilizing the thruster performance is solved to calculate the trajectory, ight time, _V , and the propellant consumption for time-optimal and fuel-optimal strategies. @en

Stand-alone interplanetary CubeSats require primary propulsion systems for orbit maneuvering and precise trajectory control. The current work focuses on the design and performance characterization of the combined chemical-electric propulsion systems that shall enable a stand-alone 16U CubeSat mission on a hybrid high-thrust-low-thrust trajectory from a supersynchronous geostationary transfer orbit to a circular orbit about Mars. The high-thrust chemical propulsion is used to escape Earth and to initiate stabilization at Mars. The low-thrust electric propulsion is used in heliocentric transfer, ballistic capture, and circularization. For chemical propulsion, design and performance characteristics of a monopropellant thruster and feed system using ADN-based FLP-106 propellant are presented. For electric propulsion, a performance model of an iodine-propelled inductively coupled miniature radiofrequency ion thruster is implemented to calculate the variation of thrust, specific impulse, and efficiency with input power. A power-constrained low-thrust trajectory optimization using the thruster performance model is pursued to calculate the transfer time, ΔV, and the required propellant mass for fuel-optimal and time-optimal transfers. Overall, the combined chemical-electric systems yield a feasible propulsion solution for stand-alone CubeSat missions to Mars that balances propellant mass and transfer time.

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Mars Atmospheric Radiation Imaging Orbiter (MARIO) is a 16U stand-alone CubeSat mission that shall escape Earth, perform autonomous deep-space cruise, achieve ballistic capture, and enter an operational orbit at Mars to perform thermal radiation imaging. This work focuses on the systems design of MARIO. The design of combined chemical-electric propulsion systems, comprising FLP-106 based green chemical monopropellant thruster and the iodine-fueled RF ion thruster, for hybrid high-thrust-low-thrust Earth-Mars transfer is presented. Reflectarrays along with high-gain antennas are utilised to establish long-distance low-bandwidth X-band communication link with the Earth. Electrical power system design is pursued to provide steady power to the system during the transfer and science operations phases. A novel autonomous navigation strategy is proposed which includes horizon-based optical navigation near target bodies and deep-space line-of-sight navigation for accurate state estimation for autonomous operations. Details regarding on-board processing, attitude determination, and thermal control are delineated. Feasible budgets for mass and communications link are obtained. The structural composition of MARIO is detailed.

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