Flight Times to the Heliopause Using a Combination of Solar and Radioisotope Electric Propulsion

Kurzfassung

We investigate the interplanetary ight of a low-thrust space probe to the heliopause, located at a distance of about 200AU from the Sun. Our goal was to reach this distance within the 25 years postulated by ESA for such a mission (which is less ambitious than the 15-year goal set by NASA). Contrary to solar sail concepts and combinations of ballistic and electrically propelled ight legs, we have investigated whether the set ight time limit could also be kept with a combination of solar-electric propulsion and a second, RTG-powered upper stage. The used ion engine type was the RIT-22 for the �rst stage and the RIT-10 for the second stage. Trajectory optimization was carried out with the low-thrust optimization program InTrance, which implements the method of Evolutionary Neurocontrol, using Arti�cial Neural Networks for spacecraft steering and Evolutionary Algorithms to optimize the Neural Networks' parameter set. Based on a parameter space study, in which the number of thrust units, the unit's speci�c impulse, and the relative size of the solar power generator were varied, we have chosen one con�guration as reference. The transfer time of this reference con�guration was 29.6 years and the fastest one, which is technically more challenging, still required 28.3 years. As all ight times of this parameter study were longer than 25 years, we further shortened the transfer time by applying a launcher- provided hyperbolic excess energy up to 49 km^2/s^2. The resulting minimal ight time for the reference con�guration was then 27.8 years. The following, more precise optimization to a launch with the European Ariane 5 ECA rocket reduced the transfer time to 27.5 years. This is the fastest mission design of our study that is exible enough to allow a launch every year. The inclusion of a y-by at Jupiter �nally resulted in a ight time of 23.8 years, which is below the set transfer-time limit. However, compared to the 27.5-year transfer, this mission design has a signi�cantly reduced launch window and mission exibility if the escape direction is restricted to the heliosphere's "nose".