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Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and MarsThe 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 (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 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 NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split 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 capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR 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 NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR 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 (HLLV) and smaller 120 t HLLV 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.
Document ID
19960001947
Document Type
Conference Paper
Authors
Borowski, Stanley K. (NASA Lewis Research Center Cleveland, OH, United States)
Corban, Robert R. (NASA Lewis Research Center Cleveland, OH, United States)
Mcguire, Melissa L. (Analex Corp. Brook Park, OH., United States)
Beke, Erik G. (Dayton Univ. OH., United States)
Date Acquired
September 6, 2013
Publication Date
September 1, 1995
Subject Category
SPACECRAFT PROPULSION AND POWER
Report/Patent Number
E-9935
NAS 1.15:107071
AIAA PAPER 93-4170
NASA-TM-107071
Meeting Information
Space Programs and Technologies Conference and Exhibit(Huntsville, AL)
Funding Number(s)
PROJECT: RTOP 242-10-01
CONTRACT_GRANT: NAS3-25776
Distribution Limits
Public
Copyright
Public Use Permitted.

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