Low-Energy Lunar Transfer with Specific Landing Conditions

抄録

This paper presents the design methodology and results of a low-energy lunar transfer trajectory with specific landing conditions. Future space exploration missions require a precise landing technique that employs autonomous navigation based on terrain pictures obtained by an onboard camera to estimate relative states to the target landing point. Picture-based autonomous navigation requires an appropriate sunlight condition during the landing sequence, and thus determines the relationship between the landing orbital plane and sun direction. To achieve low-energy transfer, solar tidal force is utilized for perigee raising, thereby making sun direction during coasting prior to lunar orbit insertion (LOI) a key parameter for low-energy transfer trajectory design. This study employs the bi-circular restricted four-body problem to model the spacecraft dynamics. The spacecraft trajectory is solved by the combination of forward propagation and backward propagation. Forward trajectory begins from trans lunar injection (TLI), and is propagated for the specified time period along with a lunar flyby. Backward trajectory begins from LOI and targets the end point of the forward trajectory. The required delta-v for the deep space maneuver is calculated so as to match the end states of both trajectories in velocity. This study identifies three types of transfer trajectory characterized by lunar flyby conditions and Jacobi integral. When the optimized result is compared to a basic Hohmann transfer trajectory, the total delta-v from TLI to LOI is reduced by 124.8 m/s, 175.9 m/s, and 158.6 m/s, but the time of flight increases by 65.2 days, 82.3 days, and 64.6 days, respectively. All types of trajectory are feasible in terms of mission design to meet the required landing conditions, and effectively reduce the required total delta-v with an acceptable increase in the total time of flight.