Low-Thrust Trajectory Optimization for Cubesat Lunar Mission: HORYU-VI

TL;DR

This paper develops a low-thrust trajectory optimization method for CubeSat lunar missions, enabling efficient orbit insertion and circularization within limited propulsion capabilities, verified through high-fidelity simulations.

Contribution

It introduces a phased optimization approach combining impulsive and finite-burn solutions for CubeSat lunar transfer, incorporating complex gravitational models and verified with NASA's GMAT.

Findings

Achieves lunar capture within 200 days

Establishes a stable science orbit at 280 days

Reaches near-circular 100 km orbit by 450 days with 710 m/s Delta-V

Abstract

This paper presents a low-thrust trajectory optimization strategy to achieve a near-circular lunar orbit for a CubeSat injected into a lunar flyby trajectory. The 12U CubeSat HORYU-VI is equipped with four Hall-effect thrusters and designed as a secondary payload on NASA's Space Launch System under the Artemis program. Upon release, the spacecraft gains sufficient energy to escape the Earth-Moon system after a lunar flyby. The proposed trajectory is decomposed into three phases: (1) pre-flyby deceleration to avoid heliocentric escape, (2) lunar gravitational capture, and (3) orbit circularization to the science orbit. For each phase, an impulsive-burn solution is first computed as an initial guess, which is then refined through finite-burn optimization using Sequential Quadratic Programming (SQP). The dynamical model incorporates Earth-Moon-Sun-Jupiter gravitational interactions and a high-fidelity lunar gravity field. All trajectories are independently verified with NASA's General Mission Analysis Tool (GMAT). Results demonstrate that HORYU-VI achieves lunar capture within 200 days, establishes a stable science orbit at 280 days, and can spiral down to a near-circular 100 km orbit by 450 days, using a total Delta-V of 710 m/s, well within the capability of the electric propulsion system.