Encyclopedia Astronautica
Nerva Gamma


DoE nuclear/lh2 rocket engine. 81 kN. Study 1972. Isp=975s. The final Nerva Gamma flight engine was an improved version of the Alpha, a small engine that could be launched together with its stage and a payload in a single space shuttle launch.

The Gamma incorporated the growth potential of the Alpha engine to a 975-s specific-impulse version incorporating carbide fuel elements was studied (the Gamma engine), as was the potential of the reactor to double as a heat source for a long-term electrical power supply.

Three primary improvements could be designed into future nuclear engines:

  • A fuel-element change would allow a higher gas temperature and provide a higher specific impulse.
  • A dual-mode system, if incorporated, would use the reactor to generate electric power for long periods of time when propulsion is not needed.
  • The heat energy in the reactor would be used for such purposes as attitude-control thrust and payload heating.
The design features listed above were logical second-generation improvements to the state-of-the-art Alpha Engine. Increases in specific impulse could be anticipated by increasing the temperature level at the nozzle exit although many unknowns, technologically, would be involved in the design. A specific impulse of 975 s might be achievable in a core containing carbide fuel elements. This carbide-core engine was referred to as the Gamma Engine.

It was known that corrosion loss rates can be greatly reduced and temperature limits extended through the use of pure carbide, i.e., UC-ZrC, fuel elements. This material was studied experimentally at LASL from 1969, and some preliminary work was done on the design of a reactor core based on such fuel. By using carbide fuel elements the temperature limit of the exit gas could be extended from the 2730 K to roughly 315O K. This material was more brittle and cracked more easily under thermal stress than composites, and the core design had to take this into account. Also, the carbide fuel was denser than composite fuel, which would increase the engine mass by roughly 260 kg.

In striving for higher temperatures, components other than fuel elements would also be changed. The low-melting-point aluminum pressure vessel would be changed, and the nozzle would change from one incorporating a thin-walled cooling jacket to an un-cooled structure. The latter change might not be advisable depending on engine and mission lifetime requirements, all of which would be affected by Alpha Engine flight experience and by the developing United States space program.

A nuclear propulsion system possessed the singular advantage of using the reactor as a thermal-energy source for a long-life electrical generating system (dual-mode system). With the nuclear rocket engine as a thermal-energy source, power levels of 10 to 25 MWe for durations of two to five years could be achieved for more ambitious missions than were possible with the Pioneer or Mariner probes, at a mass expense less than that of then-envisioned space-power systems, such as SNAP-8 and the Brayton cycle. Some applications requiring these power levels were data transmission at high bit rates and side-looking radar mapping. An additional advantage for dual-mode operation arose for outer-planet missions where the nuclear stage was used for orbit injection; namely, by operating the power system during reactor cool-down, the additional mass of the power system could be compensated by savings in cooldown propellant. Also, the electric-generating system could be operated during the transit for payload conditioning, vehicle attitude control, communications, and data transmission for scientific experiments such as mapping of the asteroid belt.

The power-conversion system was based on the Rankine-cycle using thiophene (C4H5S) as the working fluid. Thermal energy was removed from the reactor by circulating hydrogen through the core-support structure (tie tubes) and through the boiler of the power-conversion system (primary loop). The waste heat was rejected via a radiator encircling the propulsion module by circulating hydrogen through the radiator and through the condenser of the power-conversion system. For a typical 10-KWe electric generating system, the total additional mass (including the radiator and engine modifications) would be roughly 600 to 700 kg.

Only minor modifications to the Alpha Engine would be required to accommodate the primary loop. Such modifications included the substitution of a higher-temperature material (e.g., titanium) for aluminum engine parts that were in contact with the primary-loop hydrogen and the addition of valves (to isolate the primary loop during the propulsion mode and to isolate the tie-tube circuit during the electric-generation mode). Other modifications would provide cooling of engine parts, e.g., of the control-drum actuators during the electric-generation mode.

Other uses, in addition to the dual-mode use of reactor energy, would be made of this ready source. For example, it would provide redundancy in attitude control, which was important for overall mission reliability. Hydrogen heated within the reactor would be an excellent working fluid for an attitude-control system (ACS) because it was a simple monopropellant with a good Isp. Warm hydrogen could also keep certain payload components at a desired temperature, which would be especially important if manned missions were reactivated.

Gamma-Engine component development, especially of fuel elements, would proceed gradually throughout the Alpha development period. By 1980, Alpha flight data would be available to guide the Gamma Engine design studies. Dual-mode and ACS hydrogen supply would be added to the program. Because Alpha development experience would be available, engine development could begin in 1980, leading to a flight-qualified Gamma Engine as early as 1984.

Engine: 3,410 kg (7,510 lb). Area Ratio: 100. Propellant Formulation: Nuclear/Slush Hydrogen. Restarts: 20.

Status: Study 1972.
Height: 4.46 m (14.64 ft).
Diameter: 1.24 m (4.06 ft).
Thrust: 81.00 kN (18,209 lbf).
Specific impulse: 975 s.
Burn time: 1,200 s.
First Launch: Designed 1972.

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Associated Countries
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Associated Manufacturers and Agencies
  • AEC American agency overseeing development of rocket engines and rockets. Atomic Energy Commission, USA. Responsible for development of nuclear weapons and nuclear power applications. Became part of the newly-created Department of Energy in 1971. More...

Associated Propellants
  • Nuclear/LH2 Nuclear thermal engines use the heat of a nuclear reactor to heat a propellant. Although early Russian designs used ammonia or alcohol as propellant, the ideal working fluid for space applications is the liquid form of the lightest element, hydrogen. Nuclear engines would have twice the performance of conventional chemical rocket engines. Although successfully ground-tested in both Russia and America, they have never been flown due primarily to environmental and safety concerns. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was adopted for the core of the space shuttle, and Centaur stages still fly today. More...

Associated Stages
  • Nerva Gamma Nuclear/LH2 propellant rocket stage. Loaded/empty mass 18,643/5,829 kg. Thrust 81.00 kN. Vacuum specific impulse 975 seconds. Improved version of the Alpha nuclear stage designed to fit into the space shuttle payload bay. Additional propellant modules could be added in orbit. Such propellant modules would have a mass of 23,181 kg, including 21,265 kg of usable propellant. Given an Alpha engine development program, it would have been flight tested by 1984. In addition to propulsion, it would provide 10 to 25 MWe power for missions of two to five years duration. More...

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