Testing and Reliability

Nuclear reactor power and propulsion systems for human exploration missions must be qualified to a much higher level of reliability than power and propulsion systems that are intended for use by robotic missions. In the past, chemical systems such as the Space Shuttle Main Engines were tested more than 400 times on 30 engine systems to establish operational safety margins. These tests included operation at full power or greater, and for full lifetime duration or greater.

Nuclear systems will likewise need to be validated for reliable operation at full power and for mission lifetime. The latter factor is especially important because some potential applications—e.g., missions using nuclear-electric propulsion (NEP) to the outer solar system—require continuous operation without maintenance or repair for periods as long as 10 to 20 years. No high-power-density reactor has, however, ever been operated on Earth, without maintenance shutdowns, for any period longer than one order of magnitude or more below such a duration.

CANDIDATE SOURCES OF ELECTRICAL POWER AND PROPULSION

Nuclear-Electric Propulsion

Currently, NASA’s Project Prometheus is pursuing the development of NEP systems. In these systems, the heat from the reactor is carried away by a coolant in a closed loop to a power-conversion system, where it is used to generate electricity. The electricity then powers one of a variety of different electric propulsion technologies that accelerate ions and eject them from the thrusters at high velocities. JPL has demonstrated ion thrusters with a specific impulse of some 3,100 seconds in space on the Deep Space 1 mission and more than 6,000 seconds in the laboratory, which greatly reduces the amount of propellant required for the mission as compared to chemical propulsion systems. The electric power from an NEP system is also available for spacecraft instrumentation. Typically the power requirements for the mission may range from ~100 kWe, the nominal requirement for JIMO, to many megawatts of electricity (MWe) that might be required for human and cargo missions to Mars and beyond.

The leading reactor concepts employ liquid-metal coolants (either actively pumped or circulated in heat pipes), liquid-metal-to-gas heat exchangers, and Brayton-cycle conversion systems. An inherent difficulty with NEP systems is that only 10 to 20 percent of the thermal energy generated in the reactor is converted into thrust. Thus, massive radiators are required to reject the waste heat into space. In addition, electric propulsion systems provide low thrust, so NEP systems must operate at near 100 percent duty for long durations to provide high delta-V. The low thrust also complicates navigation, maneuvering the spacecraft and trimming its orbit in a complex gravitational field (e.g., when orbiting a satellite of one of the giant planets). In addition, the NEP systems currently under development for robotic missions may not scale up to provide the tens of megawatts necessary for expedited human missions to Mars.

In 2004, NASA conducted a limited set of parametric studies to examine the utility of the basic JIMO spacecraft design for other missions. The results of such studies are highly dependent on the assumptions made about the system and mission performance, which are, in turn, tightly linked. Figure 2.1 shows the results of a typical set of calculations to determine how the flight time and launch mass for a Neptune mission vary for different power values for the NEP system and the specific impulse of its ion thrusters. Representative solutions from such calculations can be combined and plotted to show the scaling relationships between mission metrics such as transit time, total launch mass, reactor power, propellant mass, and delta-V (Figure 2.2).a

These parametric studies had mixed results. The JIMO design envisioned when these calculations were performed in 2004 did appear suitable for some missions—e.g., a multiasteroid sample-return mission and a Saturn-Titan mission. But missions to other important solar system destinations—e.g., Neptune—did not appear feasible because the transit times were excessively long and launch masses excessively large. Studies performed in

a  

The committee cannot verify the results of these calculations because specific assumptions—e.g., mass-to-thrust ratio, spacecraft mass, power-conversion efficiencies, and payload mass—were not provided.



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