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Figure 9.2 The lunar solar power system.

atmosphere and its water vapor, rain, dust, and smoke. Also, microwaves in this general frequency range can be received and rectified by the planar antennas, called rectennas (bottom right of Figure 9.2), into alternating electric currents at efficiencies in excess of 85%.

The beams will be 2 to 20 times more intense than recommended for continuous exposure by the general population. The beams will be directed to recten-nas that are industrially zoned to exclude the general population. Microwave intensity under the rectenna will be reduced to far less than is permitted by continuous exposure of the general population through adsorption of the beam power by the rectenna and by secondary electrical shielding. The beams will be tightly focused. A few hundred meters beyond the beam, the intensity will be far below that permitted for continuous exposure of the general population. Humans flying through the beams in aircraft will be shielded by the metal skin of the aircraft, or by electrically conducting paint on composite aircraft. Of course aircraft can simply fly around the beams and the beams can be turned off or decreased in intensity to accommodate unusual conditions. The low-intensity beams do not pose a hazard to insects or birds flying directly through the beam. Active insects and birds will, in warm weather, tend to avoid the beams due to a slighter higher induced body temperature. See Glaser et al. (1998: Ch. 4.5).

The Earth can be supplied with 20 TWe by several thousand rectennas whose individual areas total to —10 X 104 km2. Individual rectennas can, if the community desires, be located relatively close to major power users and thus minimize the need for long-distance power transmission lines. Individual rectennas can be as small as —0.5 km in diameter and output —40 MWe or as large in area as necessary for the desired electric power output. Note that existing thermal and electric power systems utilize far larger total areas and, in many cases, such as strip-mining or power line right-of-ways, degrade the land or preclude multiple uses of the land.

An "average" person can be provided with 2 kWe, for life, from —10 m2 of rectenna area or a section —3 m, or 10 feet, on a side. This "per capita" section of the rectenna would have a mass of a few kilograms and be made primarily of aluminum, semi-conductors, glass, and plastics. This is a tremendous reduction in resources to supply each person with adequate commercial energy. In contrast, coal fired power systems will use —517000 kilograms of coal to provide 2kWe to an energy-rich "average" person for a lifetime of 80 years. This is 160 kWe-y or —4210000 kWt-h and is now done in developed nations.

Rectenna areas can be designed to reflect low-quality sunlight back into space and thereby balance out the net new energy the beams deliver to the biosphere. Space/lunar solar power systems introduce net new commercial energy into the biosphere that allows humankind to stop using the energy of the biosphere. Space/lunar solar power enables the production of net new wealth, both goods and services, without depleting terrestrial resources (Criswell, 1994, 1993).

Beams will be directed to commercially and industrially zoned areas that the public avoids. Power outside the tightly collimated beams will be orders of magnitude less than is permissible for continuous exposure of the general population. Considerable "knee-jerk" humor is directed at the concept of beaming microwave power to Earth. However, the essential microwave technologies, practices, environmental considerations, and economic benefits are understood. Microwaves are key to radio and television broadcasting, radar (air traffic control, weather, defense, imaging from Earth orbit), industrial microwave processing, home microwave ovens, cellular and cordless phones and other wireless technologies. Planetary radar is used to observe the Moon, asteroids, Venus, and other planets. It should be noted that medical diathermy and magnetic resonance imaging operate in the microwave. Medical practices and lightning associated with thunderstorms produce microwave intensities in excess of those proposed for beaming of commercial power.

The core space/lunar solar technologies emerged from World War II research and development. These technologies are the drivers of economic growth in the developed nations. The space/lunar components are technologically similar to existing solar cells, commercial microwave sources (e.g., in cellular phones, microwave oven magnetrons, and klystrons), and solid-state phased-array radar systems. These are commercial and defense technologies that receive considerable research and development funding by commercial and government sources worldwide. The essential operating technologies for space/lunar solar power receive more R&D funding than is directed to commercial power systems and advanced systems such as fusion or nuclear breeder reactors.

Space solar power systems output electric power on Earth without using terrestrial fuels. Few, if any, physical contaminants such as CO2, NOx, methane, ash, dust, or radioactive materials are introduced into the biosphere. Space/lunar solar power enables the terrestrial economy to become fully electric while minimizing or eliminating most cost elements of conventional power systems (Nakicenovic et al., 1998: p. 248, p. 103). Eventually, the cost of commercial space/lunar solar power should be very low. The commercial power industry and various governments are starting to acknowledge the potential role of commercial power from space and from installations on the Moon (Trinnaman and Clarke, 1998; Deschamps, 1991; Glaser et al., 1998; ESA, 1995; Stafford, 1991; Moore, 2000; World Energy Council, 2000).

Geosynchronous space solar power satellites (SSPS) deployed from Earth

Following the "petroleum supply distribution" crises of the early 1970s, the United States government directed the new Department of Energy and NASA to conduct environmental impact studies and preliminary systems analyses of SSPS to supply electric power to Earth. The studies focused on construction of a fleet of 30 extremely large satellites deployed one a year over 30 years. Each satellite, once positioned in geosynchronous orbit, would provide 0.01 TWe of baseload power to a rectenna in the United States for 30 years (Glaser et al., 1998).

An SSPS reference design (Ref-SSPS) was developed for the 0.01 TWe satellite and used to conduct full life-cycle analyses of engineering, operations, and financing. Smaller 0.005 TWe electric SSPS were also studied. The 0.01 TWe Ref-SSPS was approximately 10 km by 20 km on a side, had a mass of ~ 100 000 tons, and required —1000 tons/y of supplies, replacement components, and logistics support. In addition, a facility in geosynchronous orbit, with a mass of —50000 tons, was used to complete final assembly and testing. The one assembly facility and the 30 Ref-SSPS were to be partially constructed in orbit about Earth from components manufactured on Earth and shipped to space by very large reusable rockets. The assembly facility components and first Ref-SSPS self-deploy from low Earth orbit to geosynchronous orbit using solar power and ion propulsion. Ion propulsion requires a propellant mass in low Earth orbit of approximately 20% of the initial mass in low Earth orbit. These numbers allow an approximate estimate of the total mass, —160000 tons, required in low orbit about Earth to deploy one Ref-SSPS and maintain it for 30 years.

Weingartner and Blumenberg (1995) examined the energy inputs required for the construction and emplacement of a 0.005 TWe SSPS. They considered first the use of 50 micron (50 X 10~6 m) thick crystalline solar cells. The following comments assume they included the GEO construction facility, make-up mass, and reaction mass for the ion propulsion in their calculations. Specific energy of production of the satellite at geosynchronous orbit and its operation over 30 years was found to be 3044 kWh/kg. Details are given in the annotations to the reference.

One 0.01 TWe Ref-SSPS that outputs 0.3 TWe-y of base-load electric energy on Earth over 30 years delivers —24000 kWe-h per kilogram of SSPS in geosynchronous orbit. This is approximately the same as for a 0.005 TWe SSPS. The SSPS delivers a net energy of —21000 kWe-h/kg. In principle, the SSPS components can be refurbished on-orbit for many 30-year lifetimes using solar power. In this way the "effective energy yield" on Earth of a given SSPS can approach the ratio of energy delivered to Earth divided by the energy to supply station-keeping propellants, parts that cannot be repaired on orbit, and support of human and/or robotic assembly operations. Assuming a re-supply of 1% per year of Ref-SSPS mass from Earth, the asymptotic net energy payback for Earth is — 60 to 1 after several 30-year periods. Eventually, the refurbished SSPS might supply — 88 000 kWe-h of energy back to Earth per kilogram of materials launched from Earth. This high payback assumes that solar power in space is used to rebuild the solar arrays and other components.

For comparison, note that burning 1 kg of oil releases 12 kWt-h or —4 kWe-h of electric energy. The first Ref-SSPS equipment has a potential "effective energy yield" —5000 times that of an equal mass of oil burned in air. If the Ref-SSPS can be refurbished on-orbit with only 1000 tons/y of make-up mass (components, propellants, re-assembly support) then the Refurbished-SSPS yields — 22000 times more energy per kilogram deployed from Earth than does a kilogram of oil. By comparison, the richest oil fields in the Middle East release —20000 tons of crude oil through the expenditure of 1 ton of oil for drilling and pumping the oil. However, one ton of oil is required to transport 10 to 50 tons of oil over long distances by ship, pipeline, or train (Smil,

1994: 13). Unlike the energy from burning oil, the SSPSs add high-quality industrially useful electric energy to the biosphere without the depletion of resources or the introduction of material waste products into the biosphere. These estimates must be tempered by the observation that practical terrestrial solar cells are the order of 500 microns in thickness and take considerably more energy to produce than is estimated above.

NASA-DoE developed life-cycle costs for a small fleet of 30 Ref-SSPSs of 0.3 TWe total fleet capacity. The calculations were done in 1977 US dollars. In the following cost estimates, 1990$/1977$ = 1.7 is assumed and all costs are adjusted to 1990$. NASA-DoE estimated that the power provided by the Ref-SSPS would cost approximately 0.10 $/kWe-h. This corresponds to ~ 1300 T$ to supply 1500 TWe-y. The National Research Council of the National Academy of Sciences reviewed the NASA-DoE program in the late 1970s and did concede that the basic technologies were available for the Ref-SSPS in the 1980s for both construction and operations (Criswell and Waldron, 1993 and references therein). However, the NRC projected energy and overall costs to be approximately a factor of 10 higher. In particular, solar arrays were estimated to be 50 times more expensive. The NASA-DoE estimated launch cost of 800 $/kg was increased three times to approximately 2400 $/kg. The NRC estimates of cost were consistent with —13000 T$ to supply 1500 TWe-y. Significantly, the NRC accepted the estimated costs for constructing and maintaining the rectennas (—26 M$/GWe-y).

Row #20 of Table 9.6 summarizes the characteristics of a fleet of Geo-SSPS, located in geosynchronous orbit, to supply commercial electric power by the year 2050. A Geo-SSPS supplies baseload power. This power is supplied via one or multiple beams to one or a set of fixed rectennas that can be viewed by the SSPS from geosynchronous orbit. A geosynchronous SSPS will be eclipsed a total of 72 minutes a day for 44 day periods twice a year during the equinoxes. The eclipse occurs near local midnight for the rectennas. Adjacent un-eclipsed SSPSs might provide power to the rectennas normally serviced by the eclipsed SSPS.

Most rectennas will need to output a changing level of power over the course of the day and the year. "Stand-alone" SSPSs must be scaled to provide the highest power needed by a region. They will be more costly than is absolutely necessary. The alternatives are to:

• Employ a separate fleet of relay satellites that redistribute power around the globe and thus minimize the total installed capacity of SSPS in geosynchronous orbit;

• Construct and employ an extremely extensive and expensive set of power lines about the Earth, a global grid, to redistribute the space solar power;

• Provide expensive power storage and generation capacity at each rectenna;

• Provide expensive conventional power supplies that operate intermittently, on a rapid demand basis, as excess power is needed; or

• Provide a mixture of these systems and the SSPS fleet optimized for minimum cost and maximum reliability.

These trades have not been studied. A fleet of geosynchronous SSPS does not constitute a stand-alone power system. A 20 TWe SSPS system will either be over-designed in capacity to meet peak power needs or require a second set of power relay satellites. Alternatively, the order of 10 to 100 TWe-h of additional capacity will be supplied either through power storage, on-Earth power distribution, or other means of producing peak power.

As noted earlier, the rectennas will output the order of 200 We/m2 of averaged power. This is 10 to 200 times more than the time-averaged output of a stand-alone array of terrestrial photovoltaics and associated power storage and distribution systems.

It is highly unlikely that Geo-SSPS can supply 20 TWe by the year 2050 or thereafter. Major issues include, but are not limited to, total system area and mass in orbit, debris production, low-cost transport to space, environmentally acceptable transport to space, and the installation rate. Extrapolating a fleet of Ref-SSPS to 20 TWe implies 220000 km2 of solar collectors and support structure, 3100 km2 of transmitting aperture, and an on-orbit mass of 200000000 metric tons. If the 2000 to 3000 Ref-SSPS were co-located at geosynchronous altitude, they would collectively appear 1.7 to 2 times the diameter of the Moon. The individual satellites would be distributed along the geosynchronous arc with concentrations above Eurasia, North America, and South America. Few would be required over the Pacific Ocean. They would be highly visible, far brighter than any star under selected conditions, and sketch out the equatorial plane of Earth across the night sky.

Each of the 2000 to 3000 satellites would have to be actively managed, through rockets and light pressure, to avoid collisions with the others. If evenly distributed along geosynchronous orbit, they would be 80 to 130 km apart or separated by 4 to 7 times their own length. Allowing for clumping over Eurasia and North and South America, they would almost be touching (Criswell, 1997).

Micrometeorites will impact SSPSs and generate debris. Much of this debris will enter independent orbits about Earth and eventually impact the SSPSs. It is estimated that over a 30-year period a small fleet of 30 SSPS with 0.3 TWe capacity will convert 1% of the fleet mass into debris (Glaser et al., 1998: p. 8). A 20 TWe fleet would eject —6 X 105 tons/y of debris. By contrast, in 1995 the 478 satellite payloads in geosynchronous orbit had an estimated collective surface area of — 0.06 km2 (Loftus, 1997). There were also 110 rocket bodies.

The estimated collision rate is ~ 10~2 impacts/km2-y (Yasaka et al., 1996; Table 2). For the 20 TWe SPS fleet, a minimum initial rate of 2000 collisions/y is implied against existing manmade objects. Nature poses inescapable hazards. Meteor storms exist with fluxes 104 times nominal background. A large SSPS fleet in geosynchronous orbit may, under meteorite bombardment, release sufficient debris that the accumulating debris re-impacts the arrays and destroys the fleet. Special orbits about Earth that are located within the "stable plane" may minimize self-collisions of SPS debris (Kessler and Loftus, 1995). However, satellites in these orbits do not remain fixed in the sky as seen from Earth. Far more artificial debris is present in low Earth orbit. A major fleet of LEO-SPS could generate sufficient debris to make travel from Earth to deep space extremely hazardous, perhaps impossible.

Ref-SSPS in geosynchronous orbit, or lower, will be the dominant source of radio noise at the primary frequency of the microwave power beam and its harmonics (higher frequencies) and sub-harmonics (lower frequencies). The preferred 12 cm microwave wavelength, ~2.45 GHz, for power beaming is inside the "industrial microwave band" that is set aside by most nations for industrial usage. Combinations of new active filtering techniques and reallocation of existing communications bands will be required for delivery of beamed power to rectennas on Earth. Neither national nor international agreements for the allocation of the industrial microwave band for power transmission are now in place. Personal communications and wireless data transmission systems are now being used without license in this frequency range.

Fleets of massive Earth-to-orbit rockets were proposed to deploy Ref-SSPS components and construction equipment to low orbit about Earth. Very large single-stage and two-stage-to-orbit launch systems were designed that could place ~300 tons of payload into orbit. The objective was to reduce launch costs to low Earth orbit to ~250 $ per pound (~500 $/kg). Analyses were restricted to hydrogen-oxygen launch vehicles. Launch noise would be a serious problem unless operations were moved from populated areas, such as the east coast of Florida, to remote areas. Also, the water vapor deposited in the upper atmosphere might deplete ozone and affect other aspects of atmospheric chemistry in the stratosphere and above.

Approximately one launch a day was required to deploy 0.01 TWe of electric capacity each year. This implies <0.4 TWe could be deployed between 2010 and 2050 for the scale of the industry and investments assumed for the Ref.-SSPS.

Freshlook study

In 1996 the United States Congress directed NASA to reexamine space solar power. Approximately 27 million dollars was expended through the year 2000.

A publicly available summary of the first part of the Freshlook Study is provided by Feingold et al. (1997) and NASA (1999). All resources continue to focus on versions of power satellites deployed from Earth to orbits about Earth. Contractor and community studies explored a wide range of low- and medium-altitude demonstration satellites and finally converged again on two designs for geosynchronous satellites - the solar "power tower" aligned along a radius to the Earth and the spinning "solar disk" that directly faces the Sun. The systems were projected to provide power at —0.1 to 0.25 $/kWe-h. Costs are similar to those for the 1970s NASA-DoE Ref-SSPS. However, recent costing models are far more aggressive and project wholesale electricity cost as low as 50/kWe-h supplied to the top of the rectenna. Low projected beam costs are achieved through:

• Attainment of launch costs of —120 - 150 $/kg, a factor of 3 to 5 lower than the 1970s Ref-SSPS studies and a factor of 100 lower than current practices;

• Avoiding the need for large assembly facilities in low or geosynchronous orbit through the use of SSPS components designed to "self-assemble" in low and geosynchronous orbits;

• Extensively utilizing "thin-film" components and minimal structural supports; and

• Assuming 40 years operational lifetime for satellites versus 30 years.

Costs for the complete system are not included. Estimates of major systems costs were reduced through:

• Minimizing up-front research and development through use of highly standardized components;

• Minimizing time between first deployment of a satellite and start of first power delivery;

• Providing power initially to countries that now use high cost power; and

• Other investors paying at least 50% of the costs of all ground facilities (launch facilities, rectennas, component manufacturing and testing, ground assembly and transportation, etc.).

The above conditions raise serious concerns. NASA, the US Air Force, and several major launch services companies have the goal of reducing launch costs to the order of 1000 $/lb. or approximately 2200 $/kg early in the 21st century. The "power tower" was projected to be —20% more massive per TWe-output than the Ref-SSPS. A very simple model of SSPS mass and power output and launch costs can be adjusted for these two factors. For total electric cost to be 0.1 $/kWe-h, including the cost of rectennas, the mass of the "power tower" or "solar disk" and its make-up mass over 30 years has to decrease from —160000 tons to —12000 tons. The original SSPS and Freshlook designs pushed photoconversion, electrical, and structural limits. Another factor of ten reduction in mass per unit of power is extremely challenging and is likely to be physically impossible. Preliminary reports from the final "Freshlook" studies indicate that space solar power satellites deployed from Earth will not be competitive with conventional power systems (Macauley et al., 2000)

Conversely, consider the challenge of deploying a space-based power system into orbit from Earth that delivers busbar electricity, 90 percent duty-cycle, at 1 0/kWe-h. Including all the mass elements associated with the Ref-SSP (satellite, make-up mass and components, assembly facility and supplies, ion-engine reaction mass), each kilogram of Ref-SSPS related mass launched to orbit is associated with the delivery to Earth of 17000 kWe-h over 30 years. Selling the energy at 1 0/kWe-h yields —165 $/kg. This return must cover launch costs and all other investments and expenditures on both the space components and the construction and operation of the rectennas on Earth. In this model the recten-nas on Earth will be the dominant expense, —60%, of a space power system that delivers inexpensive energy to Earth. It is necessary to invest less than 50 $/kg (@ 0.9 0/kWe-h) in the space components. Allowable space expenditures might increase to <170 $/kg for satellite systems that are —3 times less massive per kWe than the Ref-SSPS. This is an extremely difficult target, probably impossible, with financing, the load-following SSPS is impossible 1 0/kWe.

LEO/MEO - solar power satellites (#21)

As an alternative to Geo-SSPS, several groups have proposed much smaller solar power satellites, 10 to 100s MWe. A wide range of orbital altitudes above Earth have been proposed, from low altitude (LEO <2000 km) to medium altitude (MEO <6000 km), and orbital inclinations ranging from equatorial to polar.

Communications satellites are the core of the most rapidly growing space industry. The satellites provide transmission of television and radio to Earth, and radiotelephony and data transmission between users across the globe. Hoffert and Potter (see Glaser et al., 1998: Ch. 2.8) propose that LEO and MEO solar power satellites be designed to accommodate communications and direct transmission capabilities for the terrestrial markets. The primary power beam would be modulated to provide broadcast, telephony, and data transmissions to Earth. For efficient transmission of power from a satellite, the diameter of its transmitting antenna must increase with the square of the distance from the receiver on the Earth. Also, larger transmitting antennas are required on the satellite as the receivers on Earth decrease in diameter. Thus, attention is restricted to LEO and MEO orbits to enable efficient transmission of power to Earth. Otherwise, the power transmitter dominates the entire mass of the satellite and makes synergistic operation with communications functions far less attractive. Engineering and economics of these satellites will be generally similar to experimental LEO-SSPS units proposed in Japan.

The Japanese government, universities, and companies have sponsored modeling and experimental studies of commercial space solar power. These have focused on the proposed SPS 2000. SPS 2000 is seen by its developers as an experimental program to gain practical experience with power collection, transmission, delivery to Earth, and integration with small terrestrial power networks (Matsuoka, 1999; Glaser et al., 1998: see Nagatomo, Ch. 3.3). This satellite is to be in equatorial orbit at an altitude of 1100 km above Earth. Studies indicate a satellite mass of — 200 tons. Power output on orbit is to be — 10 MWe. Approximately 0.3 MWe is delivered to a rectenna immediately under the equatorial ground path of the unit satellite. Power will be transmitted by the satellite to a given ground receiver 16 times a day for — 5 minutes. This implies a duty cycle (D) of the satellite and one rectenna to be — (1/12 hr) X 16/24 hr = 0.056 — 6%. Thus, —18 (1/0.056) unit satellites would be required to provide continuous power to a given rectenna. Power users would be restricted to equatorial islands and continental sites. A given satellite would be over land and island rectennas no more than —30% of its time per orbit. This reduces the effective duty cycle for power delivery to —2%.

It is highly unlikely that LEO and MEO satellites can provide low-cost solar electric power to Earth. They essentially face the same burden of launch costs as described in the foregoing Ref-SSPS. However, the low duty cycle (0.01 < D < 0.3) increases the cost challenges by at least a factor of three to 100. In addition, orbital debris is far more of a concern in MEO and LEO orbits than in GEO. More debris is present. Relative orbital velocities are higher and collisions are more frequent.

The supply of 20 TWe from LEO and MEO is an unreasonable expectation. A factor of 10 increase in satellite area over GEO, due to a low duty cycle, implies >2000000 km2 area of satellites close to Earth with a total mass >2000000000 tons. The area would be noticeable. Collectively, it will be >20 times the angular area of the Moon. The components will pose physical threats to any craft in orbit about Earth. The heavy components will pose threats to Earth. For comparison, Skylab had a mass of —80 tons. The International Space Station will have a mass of —300 tons.

Space solar power satellites using non-terrestrial materials (#22 and #23)

O'Neill (1975; also see Glaser et al., 1998, Ch. 4.10) proposed that SPS be built of materials gathered on the moon and transported to industrial facilities in deep space. These are termed LSPS. It was argued that without redesign at least 90% of the mass of an SSPS could be derived from common lunar soils. Transport costs from Earth would be reduced. Design, production, and construction could be optimized for zero gravity and vacuum. NASA funded

Figure 9.3 Facilities and transportation for construction of lunar-derived LSPS.

studies on the production of LSPS. MIT examined the production and design of LSPS and factories for LSPS in geosynchronous orbit (Miller, 1979). Prior to these studies a team at the Lunar and Planetary Institute examined the feasibility of producing engineering materials from lunar resources (Criswell etal., 1979, 1980).

General Dynamics, under contract to the NASA-Johnson Space Center, developed systems-level engineering and cost models for the production of one 0.01 TWe LSPS per year over a period of 30 years (Bock, 1979). It was compared against a NASA reference model for a 0.01 TWe SSPS to be deployed from Earth that established the performance requirements and reference costs (Johnson Space Center, 1977, 1978). General Dynamics drew on the studies conducted at MIT, the Lunar and Planetary Institute, and others. The General Dynamics studies assumed there was no existing space program. New rockets and a spaceport were constructed. New space facilities were built in low orbit about the Earth and the Moon and in deep space. Note the annotations to the Bock (1979) reference. A ten-year period of R&D and deployment of assets to space and the Moon was assumed. The General Dynamics studies explicitly estimated costs of research and development, deployment over 30 years of a fleet of 30 LSPS with 0.3 TWe capacity, and operation of each LSPS for 30 years. They also included the establishment and operation of rectennas on Earth.

Figure 9.3 illustrates two of the three major facilities and transportation concepts (C and D) developed by General Dynamics for the systematic analysis of lunar production options. Study Case D assumed extensive production of chemical propellants (Al and O2) derived from lunar materials. The lunar base was sized for the production of 90% of the LSPS components from lunar materials. Most of the components were made in deep space from raw and semi-processed materials transported to deep space by chemical rockets and electrically driven mass drivers. General Dynamics projected a base on the Moon with ~25 000 tons of initial equipment and facilities, 20 000 tons of pro-pellants, and ~4500 people. Approximately 1000 people were directly involved in production of components for shipment to space. The rest supported logistics, upkeep, and human operations. People worked on the Moon and in space on six-month shifts. The space manufacturing facility (SMF) in GEO had a mass of —50000 tons and a crew of several hundred people. The lunar base and space manufacturing facility were deployed in three years. This fast deployment required a fleet of rockets similar to that required to deploy one 0.01 TWe Ref-SSPS per year from Earth, —100000 tons/year to LEO at a cost of —500 $/kg. NASA-JSC managers required this similarly sized fleet to ease comparisons between Ref-SSPS systems deployed from Earth and those constructed primarily from lunar materials. Hundreds of people crewed the logistics facilities in low Earth orbit (40000 tons) and tens of people the facility in lunar orbit (1000 tons).

General Dynamics concluded that LSPS would likely be the same or slightly less expensive than Ref-SSPS after production of 30 units. LSPS would require progressively smaller transport of mass to space than SPS after the completion of the second LSPS.

LSPS production could not be significantly increased without an expansion of the lunar base, the production facility in deep space, and the Earth-to-orbit fleet. In the context of the Ref-SSPS studies it is reasonable to anticipate by 2050 that total LSPS capacity would be no more than 1 TWe and likely far less.

These systems, engineering, and costs studies by General Dynamics provided the core relations used to model the Lunar Solar Power System. Thus, the LSP System studies, described in the following section, build directly on two million dollars of independent analyses that focused on utilization of the Moon and its resources.

Over the long term power satellites can be located beyond geosynchronous orbit, #23 in Table 9.6, where sunlight is never interrupted and SSPS power capacity can be increased indefinitely. The satellites will constitute no physical threat to Earth and appear small in the terrestrial sky. These remote SSPS will be beyond the intense radiation belts of Earth but still exposed to solar and galactic cosmic rays. Two favorable regions are along the orbit of the Moon in the gravitational potential wells located 60° before and after the Moon (L4 and L5). Power bases on the Moon and relays and/or LSPS at L4 and L5 can provide power continuously to most receivers on Earth. Advanced power satellites need not be restricted to the vicinity of Earth or even the Earth-Moon system. For example, there is a semi-stable region (L2 ~1.5 million kilometers toward the Sun from the Earth) where satellites can maintain their position with little or no use of reaction mass for propulsive station keeping. A power satellite located in this region continuously faces the Sun. The aft side continuously faces the Earth. It can continuously broadcast power directly back to Earth and to a fleet of relay satellites orbiting Earth. Such power satellites can be very simple mechanically and electrically (Landis, 1997). Asteroid and lunar materials might be used in their construction (Lewis, 1991a)

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