Pre-Phase A Report - Lisa - Nasa

136 Chapter 6 Mission Analysis
fairing and launch capability of a Delta II 7925 H. The Earth-escape orbit has an excess
normalized energy of C3 =1.1km 2 /s 2 so that the three spacecraft will slowly drift behind
the Earth. After launch and injection to the Earth-escape trajectory, the three spacecraft
will be separated and individually targeted to their desired operational orbit.
At launch, each spacecraft will be attached to a propulsion module. The propulsion
modules will provide the capability to maneuver the spacecraft into the desired operational
orbits. After reaching the operational orbits, about 13 months after launch, the propulsion
modules will be separated from the spacecraft to avoid having excess mass, propellant,
and/or moving parts near the proof masses within the spacecraft. After reaching the final
orbits, the spacecraft positions will evolve under gravitational forces only. Micronewton
ion thrusters will be used to keep the spacecraft centered about the shielded proof masses
within each spacecraft. The thrusters could be used to perform small (∆V < 1cm/s)
maneuvers if required.
Maneuvers with total ∆V ≈ 1000 m/s are needed to reach the desired operational orbits
after launch. If a conventional chemical propulsion system was chosen, each spacecraft
would perform two maneuvers of approximately 500 m/s each. The first orbit change
maneuver would be a plane-change maneuver to incline the spacecraft orbit by i 1◦ with respect to the ecliptic. The plane-change maneuvers would take place at different
times for each spacecraft since the three final orbits are shifted 120◦ from each other along
the ecliptic. For example, one spacecraft might perform a plane-change maneuver shortly
after launch, the second spacecraft about 4 months after launch, and the third spacecraft
about 8 months after launch. Approximately 13 months after launch, each spacecraft
would perform a maneuver to stop the slow drift with respect to the Earth.
The mass of chemical propellant needed to perform the orbit-change maneuvers, combined
with the current masses estimated for each spacecraft, would total more than the
capability of the Delta II 7925 H launch vehicle. The mass of the propellant can be considerably
reduced by use of ion-engines with their much larger specific impulses (velocity
with which propellant is ejected) compared with chemical systems. The solar-powered
ion engines have an efficiency such that the required propellant mass can be reduced by a
factor of approximately 10. This is offset partly by the mass of the additional solar panels
needed to provide power for the ion engine.
The required ion-engine thrust is rather small, about 20 mN, if the engine is on during
most of the orbit transfer phase. This is much smaller than the thrust of engines designed
for interplanetary missions. However, ion-engines developed for station keeping of
geosynchronous communications satellites are of an appropriate size. In particular, the
Hughes XIPS thruster has a thrust of 18 mN. This engine is currently being tested in orbit
on a Hughes communications satellite. The mass of propellant (xenon) needed for use
with this engine to provide the required ∆V for the mission is only 20 kg per spacecraft
compared with the ≈ 180 kg per spacecraft of propellant needed for the traditional chemical
system. This reduction in mass is a major factor that enables the current mission
design to be launched on a Delta-II class launch vehicle.
The spacecraft transfer trajectories using the ion-engines have not yet been fully optimized.
A set of candidate trajectories has been found to establish the engineering
feasibility. The candidate trajectories require that the ion-engines be on during approximately
80 % of the 13 month transfer phase. The angle between the thrust direction and
3-3-1999 9:33 Corrected version 2.08