This paper presents study results quantifying the benefits of higher voltage, electric power system designs for a typical solar electric propulsion spacecraft Earth orbiting mission. A conceptual power system architecture was defined and design points were generated for several system voltages using state-of-the-art or advanced technologies. A 300-V "direct-drive" architecture was also analyzed to assess the benefits of directly powering the electric thruster from the photovoltaic array without up-conversion. Computational models were exercised to predict the performance and size power system components to meet spacecraft mission requirements. Pertinent space environments were calculated for the mission trajectory and an electron current collection model was developed to estimate photovoltaic array losses due to natural and induced plasma environments. The secondary benefits of power system mass savings for spacecraft propulsion and attitude control systems were also quantified. Results indicate that considerable spacecraft wet mass savings were achieved by the 300-V and 300-V direct-drive architectures.