A Solar Electric Propulsion Mission for Lunar Power Beaming
- Paper number
IAC-06-C3.4.-D3.4.02
- Author
Dr. Henry Brandhorst, Auburn University, United States
- Coauthor
Mr. Mark J. O'Neill, ENTECH, Inc., United States
- Coauthor
Mr. Ron Spores, Aerojet, United States
- Coauthor
Mr. Christian Carpenter, Aerojet, United States
- Coauthor
Mrs. Julie Rodiek, Auburn University, United States
- Coauthor
Mr. Michael Crumpler, Auburn University, United States
- Year
2006
- Abstract
As the NASA Vision for Space Exploration takes shape, one of the key issues that will affect the lunar exploration is the ability to provide electric power to various surface locations. This power should be available through daylight times as well as night times. While nuclear reactors and radioisotope power sources may be the choice for continuous power, it is the purpose of this paper to explore the advantages of an electric propulsion spacecraft plus laser power beaming to provide power to any location on the lunar surface.
The major benefit of solar electric propulsion (SEP) is that more payload can be delivered to the moon for less cost than by chemical means. In addition, SEP allows for high delta-V orbital adjustments to permit a range of orbital locations that best fit the application. One disadvantage is that it takes longer to reach the moon, but this is not a limiting factor for this case. Various trajectories to the moon have been calculated for an SEP spacecraft using either continuous thrust or a non-continuous thrust profile1. In the non-continuous profile, thrusting is only done when the spacecraft is in sunlight. A non-continuous thrusting profile was used for the ESA Smart-1 electric propulsion mission to the moon. One important condition is to have the spacecraft pass through the L1 point to enable low delta-V trajectories to any location on the lunar surface (as well as return to earth trajectories).
Many options exist for lunar orbits that can be used for power beaming. All have various system impacts. Stopping the spacecraft at the L1 point requires a beaming distance of about 56,000 km to provide continuous beaming to near-equatorial lunar locations. The constraints on laser power beaming over this distance have been described previously2. If a Molniya-type highly elliptical orbit is chosen, the apogee may be only about 12,000 km, hence beaming issues are eased. Now the length of time the lunar surface site is in view by the satellite is important for keeping the amount of energy storage on the surface small. In the same way, circular orbits of varying heights above the moon encounter the same site view time issue. Thus, maximum elevation of the beaming spacecraft, the angular distance, the precession of orbits around the moon and the perturbations of lunar gravity all combine to complicate the analysis. The results of the analyses on the various orbits will be presented.
Testing of GaAs solar cells in the SLA with a 800 nm laser has yielded efficiencies over 45
However, in a lunar crater or at night when beamed power is the most needed for a habitat they may be somewhat equivalent. The lunar conditions will be modeled in the paper. A dual-use version of SLA has been designed which uses two different narrow cell circuits next to one another in the focal plane of the stretched lens. During sunlit periods, the focal line of the stretched Fresnel lens is maintained on one circuit using triple-junction cells for maximum broad-spectrum conversion efficiency. During dark periods, for beamed laser power conversion, the focal line is maintained on another circuit using single-junction GaAs cells for maximum monochromatic conversion efficiency. This dual-use SLA will be described and compared to a single-GaAs-cell version for both solar (at reduced efficiency) and laser conversion. It is also important to note that as cell efficiency increases, the amount of waste heat decreases thus leading to an overall benefit. The impact of slowly traversing the Van Allen radiation belts on the solar array output will also be presented.
Many fundamental issues have been uncovered in this effort. Orbital location, the type of orbit, design of the vehicle and the size of the receiving array and pointing accuracies are all factors in the mission design and will be reported.
References: 1 R. Spores, et al, “A Solar Electric Propulsion Cargo Vehicle to Support NASA Lunar Exploration Program” (IEPC-2005-320), 29th International Electric Propulsion Conference, October 31 – November 4, 2005, Princeton, NJ
2 F. Little and H. Brandhorst, “An Approach for Lunar Power – 24/29”, International Conference on Solar Power from Space – SPS-04, June 30 – July 4, 2004, Granada, Spain
- Abstract document
- Manuscript document
IAC-06-C3.4.-D3.4.02.pdf (🔒 authorized access only).
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