By using an innovative test facility at NASA’s Marshall Space Flight Center in Huntsville, Ala., researchers are able to use non-nuclear materials to simulate nuclear thermal rocket fuels — ones capable of propelling bold new exploration missions to the Red Planet and beyond.
The Nuclear Cryogenic Propulsion Stage team is tackling a three-year project to demonstrate the viability of nuclear propulsion system technologies. A nuclear rocket engine uses a nuclear reactor to heat hydrogen to very high temperatures, which expands through a nozzle to generate thrust. Nuclear rocket engines generate higher thrust and are more than twice as efficient as conventional chemical rocket engines.
The team recently used Marshall’s Nuclear Thermal Rocket Element Environmental Simulator, or NTREES, to perform realistic, non-nuclear testing of various materials for nuclear thermal rocket fuel elements. In an actual reactor, the fuel elements would contain uranium, but no radioactive materials are used during the NTREES tests. Among the fuel options are a graphite composite and a “cermet” composite — a blend of ceramics and metals
A first-generation nuclear cryogenic propulsion system could propel human explorers to Mars more efficiently than conventional spacecraft, reducing crews’ exposure to harmful space radiation and other effects of long-term space missions. It could also transport heavy cargo and science payloads. Further development and use of a first-generation nuclear system could also provide the foundation for developing extremely advanced propulsion technologies and systems in the future — ones that could take human crews even farther into the solar system.
ABSTRACT – An initial pre-conceptual CERMET Nuclear Thermal Propulsion reactor system is investigated within this paper. Reactor configurations are investigated where the fuel consists of 60 vol.% UO2 and 40 vol.% W where the UO2 consists of Gd2O3 concentrations of 5 and 10 mol.%.Gd2O3. The fuel configuration consisting of 5 mol.% UO2 was found to have a total mass of 2761 kg and a thrust to weight ratio of 4.10 and required a coolant channel surface area to fueled volume ratio of approximately 15.0 in order to keep the centerline temperature below 3000 K. The configuration consisting of 10 mol.% Gd2O3 required a surface area to volume ratio of approximately 12.2 to cool the reactor to a peak temperature of 3000 K and had a total mass of 3200 kg and a thrust to weight ratio of 3.54. It is not known yet what concentration of Gd2O3 is required to maintain fuel stability at 3000 K; however, both reactors offer the potential for operations at 25,000 lbf and and at a specific impulse which may range from 900 to 950 seconds.
A basic model was produced which consolidated the capabilities of MCNP5 and the STAR-CCM+ code to simulate a Nuclear Thermal Rocket propulsion reactor. The MCNP5 code was used to determine the excess reactivity within the reactor as well as energy deposition profile within the reflector and fission core. The energy deposition profile was imported to STAR-CCM+ code which then determined the steady state temperature profile within the core assuming a hydrogen flow of the appropriate mass flow rate. The combined codes were not fully coupled, but were still successfully used to determine the appropriate coolant channel surface area to fueled volume ratio required to cool the core to a maximum temperature of 3000 K. If the fuel elements consist of 60 vol.% UO2 and 40 vol.% W, with the UO2 consisting of 5 mol.% Gd2O3, then only 6 lattice rows of fuel hexes are required assuming a flat-to-flat distance of 3.51 cm, and 127 coolant channels per fuel hex adding up to a surface area to volume ratio of 14.8. In a configuration where 10 mol.% Gd2O3 is used, then the reactor has a larger volume, requiring 7 lattice rows of fuel hexes and 91 coolant channels per fuel hex. Each of the configurations used W-25%Re cladding tubes which were 0.009 cm thick tungsten cladding sleeves 0.005 cm thick. The reactor masses ranged from 2761 kg at 5 mol.% Gd2O3 to 3200 kg at 10 mol.% Gd2O3, which yields a thrust to weight ratio which ranges from 3.543 to 4.107 and a power density which ranges from 6228 GW/m^3 to 9002 GW/m^3.
The 1960s NERVA rocket design had an ISP of 850 seconds.