Pre-Phase A Report - Lisa - Nasa

7.2 Spacecraft subsystem design 149
and DC. The operational attitude control signals for pointing of the Z-axis (telescope axis)
will be provided by the main signal detecting diodes, the difference between the signals
from their quadrants giving information on the wavefront tilt.
Initial beam acquisition will rely on the startrackers which are co-aligned with the main
telescope. A pointing accuracy of about 100 µrad is sufficient. Once this is established,
the laser beam will be defocussed from its diffraction-limited divergence and imaged in
the receiving spacecraft on the quadrant diodes. The resulting signal will be used to
iteratively repoint the spacecraft until the laser beam divergence can be reduced to the
minimum value.
The control torques and forces for the attitude and drag-free control during the operational
phase are provided by the Field Emission Electric Propulsion (FEEP) subsystem. The
thrusters of this subsystem are composed of an emitter, an accelerator electrode and a
neutralizer and use liquid cesium as propellant. Their specific impulse is of the order
of 10 000 s and they can provide a controlled thrust in the range of 1 to 100 µN, with
a noise below 0.1 µN. Six clusters of four thrusters each are mounted on the spacecraft
equipment panels. The major force to be compensated is the solar radiation pressure,
which, if the spacecraft is completely closed on the top (thermal shield), has a magnitude
of about 50 µN.
As the drag-free and attitude control is so intimately related to the experiment and
payload, a more detailed description of these subsystems is given in Sections 2.8, 3.2
and 7.3, also including more details on the FEEP thrusters. Because of this close relation,
the drag-free control system is a PI-provided item and described in Section 3.3 .
During the early orbit and transfer phase, the primary attitude sensors to be used are the
startrackers (ST) and Fine Sun Sensors (FSS) of the top spacecraft of the composite. As
the startrackers cannot measure large rates, an Inertial Reference Unit (IRU) isrequired
as primary sensor during some phase of the early orbit and transfer phase, e.g. during
rate reduction after separation from the launcher, during slew manoeuvres and during
V manoeuvres. The IRU’s will not be required during the operational phase and are
mounted on the propulsion module.
During the initial Sun acquisition phase and during emergency safe modes, Sun Acquisition
Sensors will be used.
7.2.4 On-board data handling
The spacecraft controller consists of two identical units operating in a String A and
String B fashion. String B acts as a warm backup and receives state data from String A
at specified intervals. String B will contain a watchdog timer to monitor String A. If this
timer runs out, String B will take over as the primary spacecraft controller.
The spacecraft controller will perform the command and data handling functions, attitide
determination, and control functions as well as processing science data. These functions
include science and engineering data collection and data storage. Power to the controller
will be supplied by the spacecraft.
A RAD 6000 flight computer as used in the Mars Surveyor Program is suggested for the
spacecraft. It contains 128 Mbytes of DRAM and 3 Mbytes of PROM. The relatively low
Corrected version 2.08 3-3-1999 9:33