Anthony C. Zuppero1,Thomas K. Larson1, Bruce G. Schnitzler1,  John W. Rice1,
James E. Werner2, Thomas J. Hill1, William D. Richins1, Garland L. Parlier1
1. Lockheed Martin Idaho Technologies,
Idaho National Engineering and Environmental Laboratory
Idaho Falls, ID 83415
2. Department of Energy, Idaho Field Office,
Idaho National Engineering and Environmental Laboratory
Idaho Falls, ID 83415
September 1998,
This is a pre-conference version of the paper for the
      "Space Technology & Applications International Forum (STAIF-99)"
                              Jan 31, 1999 to Feb 4, 1999 in Albuquerque:
             "16th Symposium on Space Nuclear Power and Propulsion"
                          Session: "Affordable Space Fission Power and Propulsion"



The Lunar Prospector has recently returned data consistent with multi-billion ton, pure veins of water ice at the forever dark poles of the moon (reference 1, 2). Public statements asserting the value of this water specify splitting it through electrolysis to make liquid oxygen (LOX) and liquid hydrogen (LH2). LOX and LH2 make a premier performance rocket fuel; LH2 is an exceptionally premier performance, thermal-rocket propellant). The moon could serve as a rocket fuel station, as in Figure 1.

The Idaho National Engineering and Environmental Laboratory (INEEL), a United States Department of Energy (DOE) national laboratory, proposes an alternative to electrolysis:

direct use of the water as a propellant for steam rocket space transportation (reference 3, 4).

A steam rocket delivers payload by using a heat source such as a nuclear reactor to convectively heat water to steam, and produces thrust by expanding the steam in a rocket nozzle, as in Figure 2.

A nuclear heated steam rocket would launch payloads from the lunar surface to low lunar orbit (LLO). The rocket would then return to the lunar surface to launch more payloads. A separate facility located at an ice formation on the moon would use a nuclear heater to melt ice, as in Figure 3. The facility would condense pure water that would be piped to the NSR to use in its rocket propellant tank.
Some of the payload delivered to LLO would be water for use as propellant (fuel) for steam rockets. While a nuclear-heated steam rocket must be used to lift off from the lunar surface, either a solar-heated or nuclear-heated steam rocket that is already in orbit could take the payloads from LLO to a Lunar Escape Orbit, which is a nearly ideal location. Note that a payload at lunar escape can be lowered to Low Earth Orbit (LEO) by an incremental aerobraking and orbit decay, which consumes only a small fraction of the propellant.
Every known paper involving steam rockets showed performances that seemed to exceed that of competing architectures by orders of magnitude (reference 5-11). Based on work presented at the 1997 Joint Propulsion Conference, the steam propulsion and chemical propulsion systems were compared. The reference system (steam propulsion) used 6 tons of steam rocket hardware to deliver 14,400 tons per year to lunar escape. The system included the launch rocket to deliver payloads. A more recent analysis estimated 20 tons would represent a more conservative system mass estimate (reference 3).

The chemical propulsion system would need to do the same job, namely to deliver 14,400 tons per year to lunar escape. That system required between 1080 and 8100 tons of electric generators; the water splitters and liquefiers required between  420  and 4,400 tons.  The total required mass for the electrolysis system was between 1,480 tons and 12,500 tons (totals for upper and lower limits). This chemical propulsion system would require 100 to 1000 times more hardware mass on the lunar surface than the steam system while delivering the same yearly payload. The hardware mass required for electrolysis is about 1000 times heavier than the steam rocket system..

In addition, analyses showed that steam rocket water containers weigh so little that engine-tank-structure assemblies weighing only tens of tons could propel payloads with masses between 1000 and 10,000 tons (reference 6, 10).

In addition to savings in hardware, the steam rocketís propulsion system offers cost savings over electrolysis systems. Systems where electricity provides the energy for propulsion all have a common step: they must reject heat radiatively to space (Figure 5). Radiative heat transfer to the vacuum of space limits the electric power produced per unit hardware mass. The available electric power limits the LOX/LH2 production rate, which is orders of magnitude lower than that of the steam rocket architecture.

In contrast, the steam rocket architecture could, in principle, operate entirely without electricity. The results of these analyses suggest fundamental reasons for the disparity in cost between electrolysis of water and using it directly for steam rockets.

These costs are significant when put in the context of manning and maintaining a permanent space system. As with other systems, the number of processes that must operate in sequence and with extreme reliability significantly increase the cost to operate a space system. The chemical propulsion option has many such elements, contrasting the few elements for the steam propulsion option (Figure 5).

The issue of practicality is crucial when investigating the potential of utilizing lunar ice for space exploration. One needs both a nearly limitless source of ice and people to man the operations. Only missions that are short enough for humans to endure and use affordable hardware are candidates for practicality. Moon missions were practical three decades ago in duration (~days) and affordability (~$20 B or less). Recent literature indicates that missions to a comet or asteroid are indeed practical today (e.g. have reasonably achievable mission delta-V), but the duration of four or more years makes exploration unfeasible (reference 14).

Only the moon offers sufficiently short manned round trip times to make exploration a reality. The use of lunar ice to provide steam propulsion could provide the answer to solving the issue of cost and practicality and lead to the next generation of space exploration.

3.0 Steam Propulsion Advantages

Steam propulsion can offer clear advantages when the alternatives meet two conditions: 1) the energy in the rocket exhaust is derived entirely from electricity; 2) the mission delta-V is less than about 6 km/s, or about 3 times the specific velocity of steam rocket exhaust (~2 km/s).

All architectures using LH2 and those using LOX, derived from electrolysis of water, chemical fueled rockets, liquid hydrogen propellant, nuclear thermal rockets or solar thermal rockets satisfy the first condition. All architectures using electricity to accelerate masses including all ion engines, spark jets, arc jets, electric thrusters, rail guns launching mass off the surface of the moon or asteroids also satisfy the first condition.

The second condition (mission delta-V less than about 6000 m/s) is satisfied by missions from the Lunar Surface to LLO, from LEO to GEO, from LLO to Lunar Escape, from Lunar Escape to GEO; missions returning payloads from about 10% of the periodic comets using propulsive capture to orbits around Earth itself; and, 100 day missions from Lunar Escape to Mars. The first and second conditions are satisfied for most missions between Venus and Jupiter.

3.1 Architectural simplicity

Architectural simplicity provides a relative advantage for steam propulsion. Figure 5 shows the relative simplicity of the steam rocket architecture compared to that of the alternative. The steam architecture needs only to extract pure water from ice and store it in a tank. Its principal disadvantage is that it must typically use 3 times as much water as the alternative; this is not a problem due to the vast amount of lunar ice available.

The alternative, chemical propulsion, must convert the heat to electricity, dump waste heat doing so, operate an electrolysis unit to split water, manage separate oxygen and hydrogen gas streams, operate mechanical refrigerator-compressor systems to liquefy oxygen and hydrogen gasses, reject heat for these processes, operate separate storage and transfer systems for cryogenic liquids. Rejecting heat to the vacuum of space requires relatively massive radiators, because the vacuum of space provides an environment like that of a thermos jar. The tanks for the cryogenic liquids must contain approximately one atmosphere and must be insulated and refrigerated to cryogenic temperatures.

3.2 Electricity
Electric generators in space are limited by the Second Law of Thermodynamics to operate no more than Carnot efficiency: they must reject heat to generate electricity. All known space electric generators have a thermal efficiency less than 20%. They reject at least 5 times more energy to space, radiatively, than electricity they generate.

Space electric generators can only reject heat by radiation to space, unless they are in contact with a thermal heat sink. The Stephan-Boltzman relation (fourth power of temperature law) limits the heat loss rate for practical radiators. Further, all radiators in space must be armored to mitigate the effects of micro-meteors. For example, a radiator operating at 1100 Kelvin with an emissivity of 0.5 using a 2 mill thick membrane to contain zero grams of thermal conducting fluid (extremely conservative), with structure and armor of 20 mills of radiating material, with density 3 (ceramic), and operating as a heat dump for a 20% efficient electric generator would achieve no more than 5.5 Megawatts per ton (0.18 kg/kw).

Space electric generators are characterized by the system mass in kilograms to generate one kilowatt of power, expressed as the reciprocal of megawatts per ton. The most optimistic electric generator proposed to be practical was quoted at 7.8 kg/kw (0.12 Megawatts per ton) (reference 15). Typical space electric generators cannot do better than 15 to 200 kg/kw. (0.066 to 0.005 Megawatts/ton).

In comparison, a steam rocket does not require electricity, and it generates 200 Megawatts per ton..

3.2.1 Thermal Heat Sinks

Thermal heat sinks are located at the lunar poles and on the surface of ice moons such as Callisto, Ganymede and Europa. The 90 Kelvin, permanently dark lunar crater basin containing the ice formation could be used as a thermal heat sink. An electric generator would exhaust vapor directly into space at that crater. In so doing, the system would not lose the water because almost all will condense back on the crater. With this heat sink, a minimum mass electric generator would consist of a turbine, a dynamo and a heat source. Based on system masses of similar terrestrial turbine-dynamo combinations, such generators may achieve 2 Megawatts per ton. A nuclear heat source would, in principle, achieve 200 Megawatts per ton and be a small relative fraction of the system mass.

The convective thermal heat sink properties of ice fields of moons may offer a way to produce LOX and LH2 competitive with steam rocket systems. The resulting increase in specific impulse (400 sec for LOX/LH2 systems, 800 sec for LH2 nuclear heated thermal rockets) could open the entire solar system to human exploration.

3.4 Solar Electric Generators
Solar electric generators must remain well below 200 Celsius, or they become resistors instead of electric generators. They are less than 30% efficient and typically less than 15% efficient. Thus, 85% of the input solar energy must be radiated to space, at less than 200 C. Commercial, unarmored solar photovoltaic materials designed for use in orbit have achieved nearly 300 watts per kilogram, at beginning of life (reference 16). This is an upper limit power factor of 0.3 Megawatts per ton.

In comparison, a steam rocket does not require solar electric generators.

3.5 Tank Mass

Lower tank mass for steam rockets provides an absolute advantage over cryogenic systems. The tanks for a steam rocket can be at least an order of magnitude less massive than tanks for cryogenic fluids such as liquid oxygen, liquid hydrogen and liquid methane. The water vapor pressure for steam tanks is about 1 kPa (1/100 atmosphere, 7 mm Hg ) at 1 Celsius, which is about 100 times lower than that of the cryogenic liquids. The water for a steam tank could be kept warm by just the waste heat from a small electric generator. Research suggested that a collapsible water propellant bladder tank could be made from 5 mill Polybenzoxazole (reference 17). Such a tank would hold hundreds of tons of water per ton of tank, which is at least an order of magnitude more than the best tanks for cryogenic liquids. A tank in orbit designed to contain water during a 100 milli-G acceleration maneuver could in principle hold thousands of times its mass in propellant.

3.6 Mass in Space

Mass already in space provides an absolute advantage for steam rockets. The cost to launch payload into LEO from Earth can be reduced by the payback factor. The lunar ice/nuclear heated steam rocket architecture has the unique potential to provide high payback. Estimates suggest 20 to 100 tons of mass for this architecture, launched from earth, would be able to return 14,000 tons per year, and 140,000 tons of mass to LEO during the operational lifetime (reference 3,4). Thus, a mass payback of approximately 1000 times the launched mass appears to be possible. This is equivalent to orders of magnitude drop in the cost of rocket propellant and mass in space. In sharp contrast, calculations suggest a electrolysis system would deliver a mass payback to LEO of 3 to 10 per year from the lunar surface, or between 30 and 100 times its mass during operational lifetime.


The discovery of abundant, pure veins of ice in permanently dark craters of the moon offers a new degree of freedom in the design of space transportation architectures. Comparison of two competing schemes, a nuclear heated steam rocket and a chemical propulsion rocket, both using water as the raw material, demonstrated the nuclear heated steam rocket to be orders of magnitude better. The result could be a dramatic drop in the cost of space transportation for missions around Earth itself and to nearby objects such as the Moon, Mars, earth-crossing asteroids, and near-earth objects.

The origin of the "orders of magnitude" lies predominantly in the hardware needed for electrolysis for chemical propulsion. The steam rocket architecture could in principle operate entirely without electricity. All other architectures derive all the energy for the rocket from electricity. Electrolysis of water and refrigeration of the resulting gases into Liquid Oxygen and/or Liquid Hydrogen cause a requirement for a massive electric generator. The hardware mass to produce electricity and manufacture LOX and LH2 requires about 1000 times heavier systems than the steam rocket system (1,400 to 12,000 tons for the electric system, Vs tens of tons for the steam system).

Other factors also provide potentially dramatic performance gains. The tank to hold water can be a collapsible bladder that can hold between hundreds and thousands of times its mass in propellant, compared to tanks for cryogenic fluids such as LH2 and LOX, which hold at most only tens of times their mass in rocket fluids. The complexity of the electrolysis system relative to the steam system has a dramatic effect on the reliability required for the space hardware. Reliability in complex systems increases cost exponentially.

Next generation space exploration is feasible when one reduces costs and duration of flight. The discovery of lunar ice and the utilization of that ice for steam propulsion can result in significantly reduced costs in comparison to cryogenic techniques.


1.Binder, A. B., "Lunar Prospector: Overview," Science 281, 1475 (1998)

2. Feldman, W. C., Maurice, S., Binder, A. B., Barraclough, B. L., Elphic, R. C., Lawrence, D. J. (1998), "Fluxes of Fast and Epithermal Neutrons from Lunar Prospector: Evidence for Water Ice at the Lunar Poles," Science 281: 1496-1500

3. Zuppero, Anthony, George Zupp, Bruce Schnitzler, Thomas K. Larson, John W. Rice, "Lunar South Pole Space Water Extraction and Trucking System,", SPACE 98 AND ROBOTICS 98 CONFERENCES, American Society of Civil Engineers, April 26-30, 1998, Albuquerque, New Mexico, U.S.A.

4. Zuppero, Anthony, Bruce Schnitzler, Thomas K. Larson, Idaho National Engineering And Environmental Laboratory, Idaho Falls, Idaho, AIAA 97-3172, "Nuclear-Heated Steam Rocket Using Lunar Ice" 33rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, July 6-9, 1997, Seattle, WA;

5. Landis, G. and A. C. Zuppero, "Optimum Rocket Transportation with In-Situ Propellants", Presentation at Joint NASA - Univ. of Arizona Conference on Resources of Near Earth Space, Univ. of Arizona, Tucson, AZ, 7-10 Jan 1991.

6. Zuppero, Anthony C, "Simple Propulsion to Mine Rocket Fuel from Near Earth Comets", Tues 2 July 1991, Missions to NEA's and Utilization, at "First International Conference on NEAR-EARTH ASTEROIDS," 30 June - 3 July 1991, San Juan Capistrano Research Institute, San Juan Capistrano, California, USA

8. Zuppero, Anthony C & Michael G. Jacox, "Near Earth Object Fuels (neo-fuels): Discovery, Prospecting and Use," 43rd Congress of the International Astronautical Federation, 28 August thru 5 Sept, 1992, Washington DC, paper # IAA-92-0159

9. Zuppero, Anthony Editor, Patrick Whitman and Mark Sykes, "Report Of The Space Energy Resource Utilization Initiative Workshop", 29 through 31 March 1993 at the University of Southwestern Louisiana at Lafayette, Louisiana; held for and funded equally by Ballistic Missile Defense Organization (BMDO) of the U.S. Department of Defense, and the Office of the Deputy Assistant Secretary of Energy for Space and Defense Power Systems, U.S. Department of Energy (DOE).

10. Zuppero, Anthony C, Olson, Timothy S., and Redd, Lawrence R, "Manned Mars Missions Using Propellant From Space," CONF 930103, 10 TH Symposium On Space Nuclear Power And Propulsion, 10 - 14 January, 1993, Albuquerque, New Mexico, Part 1, pp. 501 - 513

11. Powell, J. R. H. Ludewig, and G. Maise, "Nuclear Thermal Propulsion Engine Based on Particle Bed Reactor Using Light Water Steam as a Propellant", CONF 930103, American Institute of Physics Conference Proceedings 271, Tenth Symposium on Space Nuclear Power and Propulsion, Albuquerque, NM, 10-14 Jan 1993, pp. 579-583.

12. Zuppero, Anthony C., "Rocket Fuel to Earth Orbits From near-Earth Asteroids and Comets," American Society of Civil Engineers, Third International Conference on Engineering, Construction and Operations in Space, May 31, - 4 June, 1992, Denver Colorado, Volume II, pp 2271-2281, "Space 92",

13. Zuppero, Anthony C, Michael G. Jacox and Mark Sykes, American Nuclear Society, Regional Jackson Hole Conference, "Bootstrapping Spacebased Infrastructure with Recently Observed Water Objects in the Space Near Earth," Nuclear Technologies for Space Exploration, Jackson Hole Wyoming, 17-21 August 1992, Sponsored by American Nuclear Society, Idaho Section, NTSE-92, Vol. III, pp 625 - 634

14. Lewis, John S. and Ruth A. Lewis, "Space Resources, Breaking the Bonds of Earth," ISBN 0-231-06498-5, Columbia Univ. Press, New York 1987

15. Gilland, James, NASA LeRC, 1992, electric propulsion analyses quoted to INEEL asserted 7.8 kg/kw as the lowest mass believable for missions to Mars, at 10 Megawatt electric power level.

16. Sanyo Morton, 1994, bare photovoltaic device achieved 275 watts per kilogram.

17. Joseph Lewis, Jet Propulsion Laboratory, California Institute of Technology, 1998, as part of paper: Zuppero, Anthony, and Joseph Lewis, "Ice as a construction Material, " Workshop on Using In-situ Resources for Construction of Planetary Outposts", Lunar and Planetary Institute, sponsored by Space í98, Lunar and Planetary Institute, NASA, April 30- May 1, 1998, Albuquerque Hilton Hotel, New Mexico

Images of the Moon's South Pole
A complete electronic version of the Clementine article has been seen at:
Lunar South Pole full images (NRL Naval Research Lab) Master Pointer list to URL for Clementine
(NASA JSC Johnson Space Center) collection of images
(Naval Research Lab)      "The Deep Space Program Science Experiment (DSPSE), the first of a series of Clementine technology demonstrations jointly sponsored
     by the Ballistic Missile Defense Organization (BMDO) and the National Aeronautics and Space Administration (NASA ), launched in
     early 1994. Its principal objective is to space qualify lightweight imaging sensors and component technologies for the next generation of
     Department of Defense (DoD) spacecraft."