Instrument Host Information
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

      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

    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

    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

    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

      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.

    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

     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.

    Spacecraft systems descriptions were adapted from [WENZELETAL1992].