Pre-Phase A Report - Lisa - Nasa

148 Chapter 7 Spacecraft Design
7.2 Spacecraft subsystem design
7.2.1 Structure
The spacecraft structure is composed of the following elements:
• An exterior cylinder with top and bottom plate to stiffen the cylinder and provide
mounting points for the subsystems. Made from CFRP honeycomb with 0.3 mm
CFRP skins and a 10 mm CFRP core.
• Interface rings at each side of the cylinder. These interface rings are made in AFRP
(Aramid Fibre Reinforced Plastic) or Carbon/Carbon material.
• Two CFRP support structures for mounting the high-gain antennas and their pointing
mechanisms.
The payload module, which on the outside includes a thermal shield made of CFRP with
a thickness of 1.5 mm, is attached to the central cylinder, with a system of Kevlar straps.
The two startrackers are supported by suitable interface brackets on the PLM thermal
shield.
7.2.2 Thermal control
The thermal control subsystem of the spacecraft is basically a passive system with heaters
and their associated controls as the only active elements. All units will have primary and
redundant survival heaters controlled by thermostats. These heaters will only be used
during the transfer phase or in contingency situations. During the routine operational
phase, operation of the heaters will not be required. All units attached to the side panels
and the internal surfaces of the side panels will have low emissivity surfaces so that most
of the thermal energy is radiated to space, via radiators on the side panels.
A detailed thermal design and thermal analysis has not been performed and especially the
very demanding requirements of the payload module and the optical bench in particular
might require additional measures, such as a thermal shield on top of the central cylinder,
thus protecting the PLM against direct Sun radiation.
Also the amount of heater power during the transfer phase, in order to ensure that the
temperature of units that are switched off does not fall below their minimum survival
temparature, has not been properly assessed yet and requires more detailed analysis. It
is currently estimated at 17 W.
7.2.3 Coarse attitude control
The main requirements on spacecraft drag-free and attitude control derive from payload
constraints. The drag-free control system must force the spacecraft to follow the proof
mass to 1 nm/ √ Hz . The control signals are derived from the payload-provided electrostatic
accelerometer as described in Section 3.2 . The attitude control system points each
spacecraft towards the spacecraft at the other end of its optical path. The pointing tolerance
is 5 nrad/ √ Hz for frequencies above about 10 −4 Hz and 30 nrad for lower frequencies
3-3-1999 9:33 Corrected version 2.08