thruster concept, holding the SPS payload mass constant. This is explained by the notations on Figure 7-20 that a number of power processing units (PPU's) are required equal to the number of 100 cm ion thrusters needed. The PPU's for this thruster concept are quite heavy, weighing nearly 1600 kg each, while the relative mass of the needed PPU's for the MPD thruster concept is negligible by comparison. A significant mass reduction in the PPU concept for the 100 cm ion thruster will reduce the demonstrated performance advantage of the MPD thruster concept. This mass penalty places severe limitations on the performance of a 100 cm ion propulsion system. For example, Figure 7-20 shows that the propulsion power requirements for a 100 day trip time will equal the power producing capability of the entire SPS solar array. This means that the degradation factor of 50 percent will result in staggering and unacceptable power and economic losses. The conclusion is that the 100 cm electric ion thruster concept does not appear feasible for SPS application as presently envisioned. Figure 7-21 also shows that the mass fraction ratio of total system mass to payload mass for the MPD 60 to 100 day trip times has a range of 1. 23 to 1.197, respectively. The argon propellant requirement for this trip range is approximately 15. 55 x 106 kg, assuming a specific impulse of 7000 s for the suggested baseline MPD thruster. The number of MPD thrusters required for a 60 day LEO to GEO orbital transfer is 30 000, each thruster having an assumed thrust level of 6. 85 N. Figure 7-22 shows the expected initial system acceleration as a function of the number of electric thrusters required, which can also be equated to trip time in days when used with Figure 7-20. Earlier in-house structural analysis for the photovoltaic SPS concept indicated that a system acceleration level of 1. 0 x 10“4 g for orbital transfer was a near limit for the specific structure design. This acceleration limit corresponds to approximately 20 000 baseline MPD thrusters and a trip time of 90 days. This specific acceleration limit is compatible with the present mission profile range of trip times. The initial system acceleration requirements shown contain a small margin for steering losses. For a strictly orbit raising application, thrust vector pointing is nearly perpendicular to the radius vector for an initially circular orbit. Thrusting for inclination changes by SPS, however, requires considerable out-of-plane thruster steering and necessitates a considerable gimbaling capability for the orbital transfer thrusters. The mission profile, as
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