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].