The feasibility of an electric propulsion, nuclear thermal rocket, in-space propulsion system was evaluated by considering both the full range of theoretical possibilities and the likely performance derived from empirical arcjet performance. Joule heating was assumed for the electric propulsion device. The figures of merit were specific impulse and vehicle total initial mass, which quantify engine and mission performance, respectively, with a goal of achieving 1,200 sec specific impulse. Results were compared with state of the art nuclear thermal and chemical propulsion systems to estimate relative performance using a reference mission – a Hohmann transfer from low Earth orbit to low Mars orbit. The results indicate that while performance gains are realized relative to a nuclear thermal rocket, those gains are small – a specific impulse increase of approximately 5% – while the mass penalty is high – an increase of approximately 25%. The 1,200 sec specific impulse goal was not realized. Further, the addition of electricity generating equipment and additional heat loads pose additional mission risks. For all of these reasons, further analysis of the proposed hybrid system is not recommended.