The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (-850-1000 s) and engine thrust-to-weight ratio (-3-10), the NTR can also be configured as a "dual mode" system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTA ranging from an expendable, "single burn" trans-lunar injection (TLI) stage for NASA's "First Lunar Outpost" (FLO) mission to all propulsive , "multi-bum," NTR-powered spacecraft supporting a "spl it cargo/piloted sprint" Mars mission architecture . Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capabil ity, fuel operating temperature and lifetime, cryofluid storage and stage size. Two solid core NTR concepts are examined--one based on NEAVA (Nuclear Engine for Rocket Vehicle -Ph .D.lNuciear Engineering, Member AIAA --Aerospace Engineer, Member AIAA Application) -derivative reactor (NOR) technology , and a second concept which utilizes a ternary carbide "twisted ribbon " fuel fo rm developed by the Commonwealth of Independent States (CIS) . The NOR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NOR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration . This paper provides descriptions and performance characteristi cs for the NOR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a reference 240 t-class heavy lift launch vehicle (HLL V) and smaller 120 t HLL V option. Attractive performance characteristics and "high leverage" technologies associated with both the engine and stage are identified, and supporting parametric sensitivity data is provided. The potential for "commonality" of engine and stage components to satisfy a broad range of lunar and Mars missions is also discussed .
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