9. 6 to the elimination of the need for reactor shielding, and the fact that the engine is already man-rated when so needed. When carrying a crew transport module, it is probably desirable,at least in regard to crew confidence, to have the capability to fly the OTV manually; however, this capability already exists due to the choice of shuttle avionics for the OTV, and the only necessary modifications would be in the construction of a flight crew station within the crew transport module. The first factor of interest in the design of a space transport system is the mass ratio M, or the ratio of initial mass to final mass. This is found by the standard "rocket" equation: 6V M = e(Ig) Sp In this case, g is 9.78 meters/sec 2 , the acceleration of gravity at the Earth's surface; Isp is 455 seconds, the specific impulse of the SSME; and 6V is the transfer velocity change. Using 6V from Earth orbit to LS of 3918 m/sec, and multiplying it by 10% to give a velocity reserve for rendezvous, the necessary 6V becomes 4310 m/sec. The equation then gives a value for M of 2.634. Since the difference between the initial mass and the final mass is the mass of propellant expended, a more useful mass ratio to have is: y= M - 1 The new mass fraction y represents the ratio between fuel expended and final vehicle mass,and is equal to 1.634 for the OTV transfer case. Since each OTV makes a round trip, the vehicle must carry propellant for the return trip as part of the payload during the initial transfer. For convenience, let low earth orbit represent the "inbound" destination, and LS or lunar orbit,the "outbound" destination. The final mass at the outbound destination consists of the fixed structure mass MF, which includes the masses of the engine, thrust structure, and fixed equipment; propellant tank mass MT; mass of the inbound propellant required, MPI; a tank mass
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