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Instrument Host Overview
========================
System Configuration
--------------------
Dictated by the long distances from the Earth and the Sun at
which the spacecraft operates, the configuration of the
spinning spacecraft (5 rpm) is dominated by the large
diameter (1.65 m), Earth-pointing High-Gain Antenna (HGA)
providing the communication link and by the Radioisotope
Thermoelectric Generator (RTG) supplying the spacecraft's
electrical power. Experiment requirements for
electromagnetic cleanliness (EMC) and for minimisation of the
RTG radiation environment resulted in a 5.6 m long radial
boom which carries several experiment sensors and is mounted
on the opposite side of the spacecraft to the RTG. A 72.5 m
tip-to-tip dipole wire boom and a 7.5 m axial boom serve as
electrical antennas for the Unified Radio and Plasma Wave
Experiment ([STONEETAL1992A]). Most of the scientific
instruments are mounted on the main body, as far as possible
removed from the RTG, and in compliance with the
field-of-view requirements of the experiment sensors. The
spacecraft mass at launch was 367 kg including 55 kg of
payload and 33.5 kg of hydrazine for orbit, attitude and spin
rate adjustments.
The HGA meets the radio-link requirement to Earth with 20 W
X-band and 5 W S-band transmitters. The uplink S-band
carries commands and ranging code. The downlinks in X- and
S-band carry telemetry and turnaround ranging code,
respectively. (S-band telemetry could only be used in the
early orbit phase at close distance to Earth). This
simultaneous ranging and telemetry is a basic feature of the
spacecraft communication system.
Spacecraft mass properties and balance have been a driver in
the spacecraft design to meet the requirements both for the
launch configuration and for the deployed boom configuration
with the HGA pointing towards Earth. The spin axis of the
launch configuration was the geometric centre line. The
theoretical spin axis in deployed configuration is aligned
with the electrical axis of the HGA.
Near-continuous data throughout the mission is a prime
scientific requirement. Since continuous coverage by ground
stations is impossible for such a long-duration mission, data
are stored on-board and replayed, interleaved with real-time
data, during periods of coverage. The nominal tracking
coverage is 8 h in every 24 h.
A variety of downlink bit rates up to 8192 bit/s is
selectable, which can provide real-time data rates between
128 and 1024 bit/s and stored data rates between 128 and 512
bit/s. The prime data rates are 1024 bit/s for real-time
data ('tracking mode') and 512 bit/s for stored data
('storage mode'). The reason for the disparity between the
maximum downlink rate of 8192 bit/s and the real-time data rate
of 1024 bit/s has to do with Ulysses' science objective of
providing 'continuous' data. With a single 34-m DSN pass per day
most of the data downlinked during a DSN pass is not real-time,
but playback data, which is sometimes collected at a lower rate.
A strong scientific requirement was to have an
electromagnetically and electrostatically clean spacecraft;
EMC considerations have therefore driven the mechanical
configuration design. The spacecraft is divided into a
'quiet' and a 'noisy' zone. The former comprises an
electromagnetically shielded compartment of sensitive
experiments, whereas the latter contains the less-
susceptible but more emissive electrical spacecraft
subsystems. The preamplifiers for the wire booms and the
axial boom are mounted outside the overall spacecraft
compartment, which acts as a Faraday cage against fields
generated inside. A unipoint grounding concept has been
implemented in which the main platform constitutes an
electrical ground reference with only one area which serves
as grounding starpoint for the electrical system. All units
that produce significant magnetic fields are removed as far
as possible from the magnetometers. Low-impedance ground bus
bars are isolated from the platform and connected to it only
at the starpoint. Strict control of the magnetic properties
of all subsystems and experiments was exercised. For
example, the RTG was magnetically compensated.
The electrostatic cleanliness requirement for low-energy
particle measurements ([BAMEETAL1992A]) has been achieved by
making all external surfaces of the spacecraft electrically
conducting. This measure will also prevent differential
charging of parts of the spacecraft in the Jovian
magnetosphere.
The Jupiter gravity assist necessitates the spacecraft's
passage through the Jovian radiation belts. All subsystems
and experiments have therefore been designed to survive this
environment and radiation-hardened parts (design dose rate 60
krad) have been used throughout the spacecraft.
The spacecraft also provides autonomous system capabilities
for failure-mode detection and for safe spacecraft
reconfiguration. This is required during unexpected and/or
predicted periods of nontracking and because of the long
signal travel time between ground and spacecraft. The
preprogrammable functions include search-mode initiation to
reacquire the Earth if no command is received after a
preselectable time of up to 30 d, switch-over to redundant
units, and preprogrammed attitude manoeuvres at superior
conjunctions.
Structure and Mechanisms
------------------------
The Ulysses spacecraft has a box-type structure with two
overhanging 'balconies' and a single aluminium honeycomb
equipment platform. All electronics units of the scientific
instruments and spacecraft subsystems, most of the sensors
and the propellant tank are mounted on this platform. The
RTG is mounted on an outrigger structure to minimise its
radiation effects and to isolate the main subsystems and the
experiments from excess heat.
The two-section radial boom carries the two magnetometer
sensors ([BALOGHETAL1992A]), the solar X-ray/cosmic gamma-ray
burst sensors ([HURLEYETAL1992]) and the magnetic search-coil
sensor of the wave experiment ([STONEETAL1992A]). Because of
the radiation pattern of the RTG, it was necessary to have
the gamma-ray sensor lying as closely as possible along the
RTG centre axis. The related boom configuration achieves
this whilst maintaining the maximum length of the boom
consistent with a two-hinge system and satisfying the
requirement for the spacecraft to be balanced in both the
stowed and deployed boom configurations. The boom section
material is carbon-fibre-reinforced plastic (CFRP) tubing
with a 50 mm diameter and 1 mm wall thickness.
The electrical antennae of the wave experiment
([STONEETAL1992A]) consist of a pair of radially extending
wire booms in the spin plane and an axial boom deployed along
the orbital spin axis. The wire booms consist of 5 mm wide
and 0.04 mm thick Cu-Be ribbon stowed during launch on two
identical drive units. The wires were deployed to a length
of 72.5 m tip-to-tip by centrifugal forces acting on tip
masses after the second trajectory correction manoeuvre
(TCM-2). Each wire boom has a passive tubular root damper
which reduces relative motions between the boom and the
spacecraft by natural material damping with a time constant
of 3.5 h. The axial boom element is formed by a
pre-stressed, coilable elastic Cu-Be tube anchored in the
axial-boom drive mechanism located on the rear face of the
spacecraft. The boom element was deployed to a length of 7.5
m by a traction force through a set of rollers driven by a
stepper motor one day after the wire booms.
Several experiment sensors had protective covers on ground
and during launch; these were all successfully opened in the
first month of operation.
Thermal Subsystem
-----------------
Thermal control of the spacecraft, its subsystems and of most
of the experiments is achieved by passive means in
conjunction with a commandable internal/external power dump
and heater system. This involves an optimised layout of
subsystems which avoids hot spots on the spacecraft platform,
an efficient thermal-blanket design in order to minimise the
solar input, the compensation, by the power dump system, of
heat fluxes which are caused by the varying solar input and a
heater system for individual critical units. The most
stringent requirements on the thermal subsystem are to
guarantee a temperature above +5 degrees C at all times for
the hydrazine of the Attitude and Orbit Control Subsystem
(AOCS) and a temperature below +35 degrees C for all
experiment solid-state detectors. All spacecraft walls are
covered with thermal multilayer blankets, which are closely
fitted around the experiment-sensor apertures. The blankets
consist typically of 20 layers of aluminised mylar. The
outermost layer is kapton, coated with a transparent
conductive coating (Indium Tin Oxide) to provide an
electrically conductive outer spacecraft surface. Heat
rejection is performed by a thermal radiator, located on the
rear of the spacecraft and covered by a 2 mil kapton foil.
All units external to the spacecraft (e.g. several
experiments) are thermally decoupled from the interior.
RTG and Power Subsystems
------------------------
Electrical power is provided by the RTG at a level of about
280 W at the beginning of the mission, decreasing to about 250
W at nominal mission end. The RTG, which generates 4500 W of
thermal energy, has two major components: a heat source and a
converter. The General-Purpose Heat Source (GPHS) consists
of several elements containing the isotopic fuel 238 Pu, in
the form of PuO2. The radioactive decay energy is absorbed
at the heat source-converter interface where heat is
produced. The Si-Ge converter contains thermoelectric
elements which convert the heat into electrical energy.
Power is delivered to the experiments and subsystems at 28V
+/-2%.
Communication Subsystem
-----------------------
The communication subsystem provides capabilities for
telemetry with bit rates up to 8 kbit/s, ranging, telecommand
and radio science. It operates in X-band (downlink) and
S-band (up- and downlink). The subsystem includes two
redundant transponders (each consisting of an X-band exciter,
a modulator, an S-band receiver and an S-band power
amplifier), two redundant 20 W X-band Travelling-Wave-Tube
Amplifiers (TWTA), a TWTA Interface Unit and an S-band
Radio-Frequency Distribution Unit. A considerable amount of
cross-coupling capability exists within the subsystem
([BIRDETAL1992A], [EATON1990]).
The parabolic HGA, with both X-band (8.4 GHz) and S-band (2.3
GHz) capabilities, is the prime communications link.
Telemetry is provided in X-band, with a 2 degree beamwidth
(3 dB); downlink S-band is used for ranging, and
radio-science investigations. S- or X-band ranging
operations can be performed with or without telemetry
transmission. Both transponders can be operated
simultaneously, one in X-band and the other in S-band.
A special feature of the HGA is its ability to measure the
offset of the spin axis from the direction of the ground
station by the CONSCAN (conical scan) system. This is
accomplished by a tilt of 1.8 degrees between the S-band
antenna pattern and the spin axis which results in a spin
modulation in the uplink signal strength as the satellite
rotates. Processing within the Attitude and Orbit Control
Subsystem (AOCS) gives the offset magnitude and direction
which is either transmitted for ground analysis or employed
in a closed loop control system to minimise the offset.
Attitude adjustments are made by operating hydrazine
thrusters.
Command and Data-Handling Subsystem
-----------------------------------
The command and data-handling subsystem provides capabilities
for ground commanding, a variety of telemetry formats,
on-board data storage, and, in combination with the AOCS,
safe automatic manoeuvring.
The telecommand decoder checks commands for validity and
distributes them to the experiments and subsystems. There
are directly executed commands and memory-load commands. The
latter are stored as block commands that are validated prior
to execution of critical operations. A command-time-tagging
capability over a range of 32 s to 24 d is also available.
The Central Terminal Unit (CTU) processes command messages
received from the decoder, provides on-board timing
information, and performs formatting and encoding of data to
be sent to the ground. It also controls all on-board
automatic functions. The CTU contains a provision to
auto-check its own functioning and to switch over to the
redundant CTU in the event that a failure is detected,
assuring that important spacecraft information is maintained.
The CTU contains a master crystal oscillator from which all
synchronisation and timing signals for subsystems and payload
are derived. 32-bit timing information with a resolution of
2 s is included in every telemetry format, ensuring
unambiguous identification of the telemetry data throughout
the mission lifetime. A spin reference ('Sun pulse') and
spin segment clock (16384 pulses per spin period) are also
supplied by the CTU, based on Sun-sensor data provided by the
AOCS subsystem.
The Data Storage Units consist of two redundant tape
recorders for storage of data during those periods when the
spacecraft is not in communication with the ground
(nontracking periods) for subsequent playback and
transmission during the next tracking period. The 45 Mbit
capacity of each tape recorder is sufficient to provide
continuous storage at 512 bit/s for 22 h or 256 bit/s for 44
h.
Telemetry formats are built up of successive frames. There
are three data formats:
- Scientific format consisting of 32 scientific frames
- Interleaved format consisting of a block of 32 frames,
interleaving real-time and stored frames with a selectable
ratio (1:1, 1:3, 1:7). Formats are played back in reverse
time order, but the frames within each format and each
individual frame are in forward order
- Engineering format consisting of two frames of spacecraft
housekeeping data and containing no scientific data.
Telemetry channels are sampled in a time-ordered fashion and
allocated to specific words (8 bit), which are arranged into
frames of 128 words. In the scientific and interleaved
formats one frame consists of 110 digital science words, nine
analogue science words, two subcommutated experiment
housekeeping words, four subcommutated spacecraft
housekeeping words and three synchronisation and
identification words. Analogue channels are sampled and
converted into 8-bit words with an accuracy of 1% full scale.
There are also datation channels which contain accurate time
information on an event with a resolution of 0.488 ms (32 s
range) or 3.9 ms (256 c range). Datation channels are used
by the gamma ray-burst instrument (high resolution) and by
the magnetometer and wave experiments (low resolution).
Attitude and Orbit Control Subsystem (AOCS)
-------------------------------------------
The primary operational functions of the AOCS are to maintain
the spacecraft spin axis Earth-pointing and control the spin
rate. Additional operational functions are dictated by
trajectory control requirements, nutation damping, and by the
measurement of the attitude for scientific reasons.
The spacecraft Earth-pointing attitude is measured and
controlled by the CONSCAN system with spin-rate, spin-phase
and solar-aspect-angle information determined from redundant
Sun sensors. The Sun-sensor outputs are processed in the
data- handling subsystem to provide the spin reference pulse
and the spin segment clock. These signals and the Sun-sensor
data are then used in the AOCS electronics to determine the
spacecraft spin rate and solar aspect angle for the purpose
of closed loop on-board control, failure detection and
recovery. Hydrazine thrusters are actuated either by
telecommand or automously. These are fed from a single tank,
mounted at the launch centre of gravity, and arranged in two
blocks of four thrusters each, providing complete redundancy.
Another AOCS operation is the periodic precession manoeuvring
for correction of the apparent Earth drift with respect to
the spin axis. These can be performed in closed loop
on-board or in open loop via time-tagged command.
A special manoeuvre strategy is required for conjunctions,
since the proper spacecraft attitude depends on the operation
of the Sun sensors with a safe operational limit of the solar
aspect angle greater than 1.25 degrees.
The spacecraft carries three nutation dampers containing a
fluid whose viscous motion dissipates energy. In the
operational spin rate range the nominal damping time constant
for nutation cone angles from 2.0 to 0.02 degrees is less than
one hour.
The AOCS also includes failure-mode-detection and protection
functions which result in fail-safe operation and a
reacquisition capability in both automatic and
ground-initiated recovery sequence.
In-Flight Performance
=====================
Following launch and orbit injection of Ulysses on 6 October 1990,
the initial flight phase consisted of checking out all spacecraft
subsystems (including redundancy), deploying the radial, axial and
wire booms and switching on all experiments. The latter were
commanded on one-by-one and thoroughly checked over an extended
period between 19 October and 16 November 1990. Early in January
1991 the spacecraft was formally declared to be commissioned.
Anomalies
---------
The following is a list of (major) spacecraft anomalies to have
occurred since launch:
Nutation. Shortly after deployment of the axial boom, build-up of a
nutation-like disturbance was observed. This is now believed to be
the result of an oscillation induced by non-uniform solar heating
of the axial boom coupling into the spacecraft motion, together
with under-performance of the passive nutation dampers on board the
spacecraft. The onboard CONSCAN system has been successfully
employed to control subsequent episodes of nutation, which occurred
in 1994 and 1995. Nutation is predicted to return in 2001. More
details on the nutation anomaly are discussed in the Ulysses
MISSION.CAT.
Disconnect Non-Essential Loads (DNEL). The DNEL condition is a
spacecraft safing mode, and is known to be associated with
operation of the latching valve when coinciding with unpredictable
peaks in payload current demand. Overcurrent criteria are violated,
and the onboard protection logic correctly operates, placing the
spacecraft in a minimum current demand mode by disconnecting the
scientific payload. To date, 8 DNEL events have occurred.
CTU-2 Anomaly. During check-out of the Data Handling Subsystem
following Jupiter flyby, telemetry from the redundant Central
Terminal Unit (CTU-2) was found to be partially corrupted. CTU-2 is
a redundant unit, however, and will only be used in case of CTU-1
failure. If this occurs, extra data processing will be implemented
to minimise the impact of the CTU-2 malfunction.
References
==========
Spacecraft systems descriptions were adapted from [WENZELETAL1992].
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