EARTH-TO-ORBIT TRANSPORTATION FOR SOLAR POWER SATELLITES Gordon R. Woodcock Boeing Aerospace Company and Gerald Hanley Rockwell International Transportation of solar power satellites to space will require cargo transport capability much greater than any other space technology application thus far investigated. The cost of space transportation operations represents, in the reference SPS system, more than one fourth of the total production cost of the SPS's, even though the unit cost in dollars per kilogram is projected to be much less than that presently foreseen for the Space Shuttle. Three-fourths of the cost is contributed by the launch systems (including launches delivering orbit transfer propellant) with the remainder contributed by orbit transfer systems. Further, developing the vehicles required and acquiring the operational vehicle fleet is the largest single element of SPS nonrecurring cost. Consequently, the design approach for these vehicles and their ability to achieve the projected cost is of great importance to the economic practicality of solar power satellites; commensurate importance has been given to the concept definition for space transportation in the SPS Systems Definition studies. The history of SPS launch vehicle evolution is shown in Figure 1. Early studies of SPS launch vehicles examined ballistic systems shaped like large Apollo spacecraft; these were to return to Earth engines-first by aerobrakmg and land at sea for recovery by ship. Single-stage and two-stage options were examined. The performance of the two-stage systems was enough better to more than offset their greater operational complexity. Later? comparison of winged and ballistic launch vehicles concluded that the winged systems were preferred. Although more expensive per unit, shorter turnaround time permits a smaller vehicle fleet, effecting overall savings. This trade resulted in selection of the two-s€age winged vehicle now represented as the SPS reference launch vehicle. The size of the vehicle was somewhat arbitrary. The only specific consideration was selection of a payload bay large enough to accommodate a fully-assembled electrical slip ring, 16 meters in diameter. The payload capability of the reference vehicle was estimated as 420 gross tonnes, with an effective net payload of about 360 to 380 tonnes after accounting for mass of payload pallets, propellant containers, and similar factors. This vehicle design was based on ''normal” technology growth. The second stage engine was the Space Shuttle Main Engine (SSME) and the first stage engine was assumed to be a new-development gas-generator oxygen-hydrocarbon engine. Modest use of composite materials in the dry structure was assumed, limited to areas not subjected to high temperatures as a result of aerodynamic or plume heating. The booster is a heat-sink design for reentry heating: the orbiter assumes an advanced Shuttle-type RSI, with improved durability and serviceability. Subsystems masses were based on extrapolations from the Shuttle subsystems. The reference vehicle is shown in Figure 2. Figure 3 presents a mass distribution, and Figure 4 shows the corresponding first unit cost. Figure 5 shows the schedule estimates for vehicle turnaround upon which the fleet size is based. Alternative vehicle designs have been created by other studies. The most important are (1) A parallel-burn, crossfeed configuration developed by Rockwell International on their SPS studies; (2) A single-stage-to-orbit airbreathing/rocket runway takeoff vehicle concept developed by Rockwell, and (3) A smaller HLLV concept developed by Boeing. The parallel-burn configuration yields about 10$ improvement in payload capability at a given liftoff mass, but involves increased operational complexity. An adequate tradeoff to select between series and parallel burn has not been conducted. The airbreather concept was representative of vehicle designs that might be attainable with highly advanced propulsion and structures technology.
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