VI-B-3 WINGED LAUNCH VEHICLES Robert B. Bristow Engineering Analysis Division SUMMARY The purpose of this study was to define a heavy lift launch vehicle for the solar power satellite system. Each solar power satellite has been estimated to weigh nearly 91,000 metric tons (200,000,000 pounds) in geosynchronous earth orbit, and would thus require the launch vehicle fleet to deliver close to 227,000 metric tons (500,000,000 pounds) to a low earth assembly orbit for each power satellite. The mission profile used in sizing the launch vehicles was an unmanned launch from the Kennedy Space Center to an elliptical earth orbit. The launch vehicle injected the payload into a 92.5 x 500 kilometer (50 x 270 nautical mile) orbit; the orbit injection point was near the 92.6 kilometer (50 nautical mile) perigee. After injection the payload was separated from the launch vehicle's second stage. The payload was then placed into a circular orbit at the 500 kilometer (270 nautical mile) apogee by a propulsive system in the payload. Payload at 500 kilometer (270 nautical mile) circular was 453.59 metric tons (1,000,000 pounds). The second stage of the launch vehicle would reenter the earth's atmosphere at an altitude of 121,920 meters (400,000 feet after one orbit. The mission profile required each spent stage of the launch vehicle to be returned to the launch site for reuse. Candidate concepts for the launch vehicle were sized based on a predicted technology level expected in 1995. In addition, 20 percent of the vehicle's booster stage dry weight was established for growth allowance. The configurations investigated were two stage launch vehicles; the propulsion sequence was a series burn. Recovery systems were defined and sized for each stage. The systems studied were: (1) winged, powered flyback; and (2) winged, nonpowered glideback systems. A 3 degree-of-freedom launch trajectory program was used to simulate vehicle flight for each of the configurations studied. The analysis program optimized the launch vehicle tilt profile and exo-atmospheric portion of the flight. A parametric study was made of stage mass fractions, wing configuration, and booster fuel. The launch vehicles considered were sized for payloads of 453.59 metric tons (1,000,000 pounds). The launch vehicle defined in this study was a two stage, series burn, winged vehicle, with a payload capability to the target orbit of 453.59 metric tons (1,000,000 pounds). Vehicle gross lift off weight was 13,738 metric tons (30,286,651 pounds) and the thrust to weight ratio at lift off was 1.3. The booster stage was a heat sink powered flyback stage that required 20 LOX/C-Ho high chamber pressure engines for boost, and o o
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