28.5° inclination is assumed, a chemically augmented propulsion system will be needed to provide an acceleration of approximately 10“3 g for approximately 21 h to boost the 1. 4 x 108 kg SPS payload (chemical system mass not included) to an altitude of approximately 2037 km to intercept a totally sunlit orbit. Thus, the trade is between the cost and complexities of the chemically augmented propulsion system versus the increase cost because of performance penalties of HLLV launches to a 55° assembly orbit. First iteration conclusions for a smaller scale SPS indicate that these relative costs nearly cancel each other. If, however, it can be demonstrated that HLLV launch penalties can be reduced to less than 10 percent, then an economic advantage is evident. For this study, an SPS assembly orbit of 435 km and 55° inclination is assumed, utilizing low thrust electric propulsion for orbital transfer from LEO to GEO and taking from 60 to 100 days to arrive on station. A second important assumption for SPS LEO to GEO orbital transfer is that a low thrust sortie through the Van Allen radiation belt will cause an exposed (nominal coverslide assumed) photovoltaic array to degrade to a significant degree. Previous in-house studies have provided considerable insight into the problem of degradation by charged particle radiation with regard to design impacts and operating procedures for SPS transits through the ''hot" portion of the radiation belts. Analyses have shown that at altitudes of 5556 to 7408 km, . , , . , ,, , r . „ electrons an equivalent intensity on the order ot 10iO 1 MeV -^'2 ^ay will penetrate a 3 mil cover-slide and reach the vulnerable portion of a solar cell. The fastest expected SPS transit time through this region using low thrust propulsion is on the order of 4 days, which will result in an almost instantaneous power degradation of exposed solar arrays on the order of 25 percent. Although this altitude range describes only the ''hot" portion of the belt, significant array degradation occurs in an altitude range from as low as 1852 km to approximately 13 890 km. The implications are apparent. Figure 7-20 shows that trip times on the order of 60 days from LEO to GEO require many thousands of magnetoplasmadynamic (MPD) thrusters requiring very large quantities of power. A significant portion of the total spacecraft power generation capability must be available for the electric thrusters during the orbital transfer phase. If solar cells are the energy collecting devices, then those exposed are subject to greater degrees of degradation. Referring to Figure 7-20, extending the allowable trip time to 100 days significantly reduces the power requirement of the propulsion system and subsequent degradation; thus, extending the mission trip within the constraints
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