types of rocket engines. The first stage supplies approximately 2/3 of the delta-V requirements, after which it is separated and returns to the LEO staging depot. The second stage completes the boost from LEO to GEO and also provides the thrust for returning the stage to the LEO staging depot for reuse. The vehicle gross weight at start burn of 1,290,000 kg of which 400,000 kg is payload and 830,000 kg is propellant (Figure A-20). More than two out of every three HLLV flights would be required to carry fuel for the COTV. The chief advantage of the chemical COTV is the short trip times of 30 hours for the round trip from LEO to GEO and back to LEO. The nuclear COTV concept analyzed combined the desirable features of the chemical COTV and the electrical COTV - high thrust and high specific impulse, respectively. The stage, shown on Figure A-21 , has a nuclear gas core, light bulb-shaped engine with a theoretical specific impulse of 2250 seconds and a thrust level of 890,000 newtons. The component mass breakdown is given in Table A-3. Although such a system could meet the short trip time requirement for personnel transfer and the high performance requirement for cargo transfer, the development risks and the presence of nuclear materials in LEO eliminated this system from further consideration. The electric approach utilizes low thrust engines with high Isp and round trip time measured in months rather than hours. Studies were conducted to determine the optimum thruster, propellant type, and trip time. Thruster types considered were the nuclear-thermal, resistojet, thermal arcjet, MPD arcjet and ion bombardment. Fuels considered were mercury, argon, cesium and xenon. Of these, the ion bombardment thruster had the most satisfactory performance in terms of thrust and specific impulse technology readiness. Argon, stored as a cryogenic, appears to be the best propellant because of its ready availability as a by-product of LOX production and its consequent low cost (about $0.50 per kg). Furthermore, experience with the development of 8 and 30 cm diameter mercury ion thrusters is sufficient to analytically predict the performance of argon ion thrusters as large as 120 cm.
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