## Info

Volume

1.7 m3

Table 10-28

Linear dimensions

1.4 m

Table 10-28

Body cross-sectional

2.0 m2

Table 10-28

Moment of inertia

SOkg-m2

Table 10-28

In Sec. 7.5.1, we decided to try eliminating a kick stage and flying the spacecraft up using low-thrust chemical propulsion. In order to maintain reasonable efficiency, we initially assume a metered bipropellant system with an of 300 s (Sec. 10.4.1, Chap. 17). Using the rocket equation (Eq. 17-7), we can compute the propellant mass as 28 kg and then add small amounts for attitude control and margin as given in Table 10-30. Here our knowledge of the propellant mass as a fraction of the spacecraft mass is good, although the spacecraft mass itself is not yet well known. Because the propellant mass is small, we may choose later to go to a simpler monopropellant system or to have the launch vehicle put FireSat directly into its end orbit.

Given an approximate mass for die whole system we can estimate the size and moments of inertia from Table 10-28. This, in turn, can tell us something about the solar array configuration. We estimate the body area at 2.0 m2 and the required solar array area at 1.7 m2. So we can probably avoid solar panels altogether and use an omnidirectional array consisting of solar cells mounted on the non-nadir facing sides of the body. This will be compact, economical, and easy to control.

Finally, Table 10-31 presents two ways of developing a preliminary weight budget for FireSat We can estimate the mass of each subsystem as a percentage of spacecraft dry mass or as a percentage of the payload mass. Column (1) lists die average percentage of spacecraft dry mass devoted tp each subsystem based on the historical data for spacecraft listed in Appendix A. The resulting mass distribution and margin are shown in column (3). Column (2) lists die same data expressed as the average percentage of payload mass devoted to each subsystem. The resulting FireSat mass distribution is shown in column (4). We recommend using the mass'distribution shown in column (4). A weight margin of at least 25% at this stage of development is appropriate. The column (3) approach resulted in a "margin" of only 112 kg or 8% of the spacecraft dry mass. This approach prematurely divides die available margin among the subsystems. We recommend maintaining the margin at the system level and then allocating it to the payload or other subsystems as necessary throughout the development

Element of Weight Budget |
(1)' Est % of Spacecraft Dry Mass |
<2)t Est % of Payload Mass |
Est. Mass Based on Col.(1) (kg) |
(kg) |
Comments |

Payload |
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