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Pioneer Vehicle Descriptions

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Pioneer Venus (Pioneer 12) Vehicle Description

SPACECRAFT:  PIONEER VENUS ORBITER


  Spacecraft Information
  ======================
    Launch Date                    : 1978-05-20
    Instrument Host Name           : PIONEER VENUS ORBITER
    Instrument Host Type           : SPACECRAFT

  Mission Information
  ===================
    Mission Start Date             : 1978-05-20
    Mission Stop Date              : 1992-10-07
    Mission Alias Name             : P12

  Targets
  =======
    VENUS

  Instruments
  ===========
    PVO PLASMA WAVE ANALYZER
    FLUXGATE MAGNETOMETER

  Spacecraft Description
  ======================

    Extracted from:
    `Pioneer Venus Spacecraft Design and Operation'
    IEEE Transactions on Geoscience and Remote Sensing,
    vol. GE-18, No. 1, January 1980

    By George J. Nothwang

    I.  Introduction
    The Pioneer Venus mission objectives dictated the
    requirement for two spacecraft designs designated the
    Orbiter and the Multiprobe. (The Multiprobe is defined as
    the Bus with the one Large Probe and three identical Small
    Probes attached in the launch/cruise configuration.) The
    conceptual designs of these spacecraft resulted from Phase B
    studies conducted from October 1972 to July 1973, and after
    selection of the spacecraft contractor, Hughes Aircraft
    Company, in February 1974, a spacecraft conceptual design
    review was conducted in November 1974.
    The Orbiter and Multiprobe utilized the same designs to the
    maximum extent possible to minimize costs. In addition,
    designs of subsystems or portions of subsystems from
    previous spacecraft designs (such as OSO and Intelsat) were
    utilized to the maximum extent possible with little or no
    modifications. This commonality in the two spacecraft
    designs also resulted in certain amounts of commonality in
    ground test equipment and test software as well as
    commonality in spacecraft flight operations and associated
    software. A photograph of the Multiprobe (foreground) and
    Orbiter in the manufacturer's facility (Hughes Aircraft
    Company, El Segundo, CA) is shown on the cover.
    II. Spacecraft Design and Operation
    The design and operation of the orbiter will be described
    first and the Multiprobe second. The Multiprobe description
    and operation will then be separated into the Bus, Large
    Probe, and Small Probe segments.
    A. Orbiter
    The general configuration of the Orbiter after launch by an
    Atlas SLV-3D/Centaur D-1 AR is shown in Fig. 1. The weight
    of the spacecraft and 12 scientific instruments immediately
    after separation from the launch vehicle was 553 kg (1220
    lbs) which included 32 kg (70 lbs) of hydrazine for
    trajectory correction maneuvers and spin axis orientation
    and 179 kg (398 lbs) of orbit insertion motor solid
    propellant and inserts. Immediately after separation from
    the Centaur launch vehicle, the spacecraft was automatically
    spun up to approximately 6.5 rpm and after establishing
    satisfactory ground communications with the Deep Space
    Network (DSN), commands to deploy the magnetometer boom and
    to orient the spacecraft spin axis perpendicular to the
    ecliptic were transmitted. The nominal spin rate was
    increased to 15 rpm with the spin axis (+Z axis) pointed in
    a northerly direction and this attitude was maintained in
    the cruise phase of the mission except during short periods
    for trajectory corrections. Communications were normally
    maintained through the despun high-gain antenna to maximize
    the data rates.
    Two days before reaching Venus, the spacecraft was
    configured for orbit insertion. Communications were
    transferred to the omni antennas and the spacecraft
    including the high- gain antenna spun up to 52 rpm to
    provide acceptable gyroscopic stiffness during motor burn.
    Since the motor burn had to occur while the spacecraft was
    being occulted by Venus, commands were loaded and
    subsequently executed from the on-board stored command logic
    without real-time communication. Orbit insertion was
    achieved with an orbit inclination of 105 degrees with
    respect to the equator and a nominal orbit period of 24 h.
    After reacquisition of the spacecraft from occultation, a
    series of maneuvers were performed to point the spacecraft
    spin axis (=Z axis) perpendicular to the ecliptic in a
    southerly direction, despin the high-gain antenna, and slow
    the spacecraft spin rate to about 5 rpm, the preferred rate
    for scientific data. Nominal orbital operations were then
    begun which include orbit periapsis and period adjustments
    and spacecraft attitude adjustments.
    The Orbiter spacecraft consists of the following subsystems
    and functions: Mechanical function (including the Spacecraft
    Structure), Thermal Function (accomplished by the
    Structure/Harness Subsystem), Controls Subsystem, Propulsion
    Subsystem, Data Handling subsystem, Command Subsystem,
    Communications Subsystem, and Power Subsystem.
    1) Mechanical: The mechanical features of the spacecraft can
    be described by six basic assemblies, as seen in Fig. 1. The
    despun antenna assembly, the bearing and power transfer
    assembly (BAPTA), the BAPTA support structure, equipment
    shelf, substrate (solar array), orbit insertion motor (OIM)
    and its case, and thrust tube. The shape and equipment
    layout conform to the basic mechanical requirements of a
    spin-stabilized vehicle. The solar cells on the cylindrical
    solar panel, antenna orientations, and thrust vector
    orientations provide efficient power, communications, and
    maneuverability while the Orbiter is spinning in its cruise
    and orbit attitudes.
    2) Thermal: The thermal design is based on isolating the
    equipment from the external solar extremes experienced
    during the mission. (Solar intensity increases by a factor
    1.98 from Earth to Venus). Commandable heaters are provided
    to maintain the orbit insertion motor and safe and arm
    device within their specified temperature ranges, to prevent
    possible freezing and hydrazine monopropellant , and to make
    up heat balance should there occur an inadvertent trip of
    nonessential spacecraft loads. Fifteen thermostatically
    controlled thermal louvers are mounted on the aft side of
    the equipment shelf beneath units having high dissipations.
    3) Controls: The controls subsystem provides the sensing
    logic and actuators to accomplish the following
    stabilization, control, and reference functions:
    a) spin axis attitude determination (via use of slit
    field-of-view type sun sensors and star sensors and star
    sensors), science roll reference signals generation, and
    spin period measurements;
    b) control of thrusters for spin axis attitude maneuvers ,
    spin speed control, and spacecraft velocity maneuvers;
    c) high-gain antenna azimuth despin control and elevation
    positioning to a desired earth line-of- sight pointing ;
    additionally, antenna slew control for open-loop tracking of
    the Earth line-of- sight;
    d) magnetometer sensor deployment;
    e) nutational damping, via use of a partially filled tube of
    liquid Freon E3.
    4) Propulsion: The propulsion subsystem provides the
    hydrazine monopropellant storage, pressurization,
    distribution lines, isolation valves, filtering, and
    thruster assemblies used to accomplish Orbiter maneuvers
    throughout the mission.
    5) Data Handling: The data handling subsystem conditions and
    integrates into command- selectable (choice of thirteen
    fixed and one programmable) formats, all analog and digital
    telemetry data (248 assigned channels) originating in the
    subsystems and science instruments. The selected format of
    the all-digitized data modulates a 16 384-Hz subcarrier at a
    command- selectable (choice of thirteen rates between 8 and
    4096 bps) bit rate. The resulting information is routed to
    the communications subsystem for modulation of the downlink
    S-band carrier. The data handling subsystem includes a data
    memory, consisting of two data storage units (DSU) that is
    intended primarily for use during any occultation. Data are
    stored or read out at the commanded bit rate. Each DSU has a
    capacity of 524 288 bits (equivalent to 1024 telemetry minor
    frames).
    6) Command: The command subsystem decodes all commands
    received via the communications subsystem at the fixed rate
    of 4 bps, and either stores the command for later execution
    or routes the command in real time to the addressed
    destination. Each of the 381 assigned commands is either
    completely decoded (discrete-type command) by the command
    subsystem and the execution command generated, or is
    partially decoded (quantitative-type command) by the command
    subsystem and the command is routed to the addressed
    destination for final decoding.
    7) Communications: The communications subsystem provides
    radiation reception and transmission capabilities for the
    command and telemetry information. The uplink command
    capability is maintained by modulating an S-band carrier of
    approximately 2.115 GHz. The downlink telemetry modulates an
    S-band carrier approximately 2.295 GHz. There are two
    redundant reception channels; each includes a
    hemispherically omnidirectional antenna (aft or forward)
    that spatially supplements the other to produce total
    spatial coverage. Optionally by command, the forward antenna
    is replaceable by a high-gain antenna or a high-gain back-up
    antenna.
    The S-band downlink is assignable by command to any one of
    the aft or forward omnidirectional antennas, or to the
    high-gain or high-gain back-up (directional) antennas. Its
    frequency is a multiple of the uplink frequency; or in the
    absence of an uplink signal, it is a multiple of a crystal
    oscillator located in the receiver. The downlink may also be
    transmitted via any one of, or some pairs of, four 10-W
    power amplifiers.
    There is an additional transmitter in the X-band range (the
    frequency is 11/3 of the S-band downlink frequency) that is
    for use in occultation measurements. The transmission is
    unmodulated through the high-gain antenna only.
    8) Power: The power subsystem provides semiregulated 28 V 10
    percent to all spacecraft loads (including science
    instruments). The primary source of power is the main solar
    array. When the solar panel output cannot provide adequate
    power for all spacecraft loads (at low sun angles and during
    eclipses), the two nickel/cadmium batteries (each rated at
    7.5 Ah full capacity) come on line automatically through the
    discharge regulators. Battery energy is replenished through
    a small boost charge array. The power interface unit
    provides power switching for the propulsion heaters and OIM
    heaters. It also contains fuses for these circuits and the
    science instruments input power lines.
    Power is distributed on four separate power buses. If a
    spacecraft over-current condition or under-voltage on either
    battery occurs, loads are removed to protect the spacecraft
    from potential catastrophic failure by tripping off buses in
    the following sequence: science, switched loads, and
    transmitter. This leaves only those loads that are
    absolutely essential to spacecraft survival in a
    continuously powered ON mode. The RF transmitters and
    exciters are on the transmitter bus. Controls and data
    handling units are on the switched loads bus. Scientific
    instruments are on the science bus. Command units, OIM and
    propulsion heaters, power conditioning units, and spacecraft
    receivers are on the essential bus. Excitation for the pyro
    bus is derived from a battery tap located 16 cells (of a
    total of 24) from the ground reference level. The bus
    voltage is limited to 30.0 V by seven shunt limiters that
    dissipate all excess solar panel capacity in load resistors
    mounted on the solar panel substrate and equipment shelves.

  Platform Descriptions
  =====================

    Platform MAGNETOMETER BOOM
    --------------------------

      An 4.8 meter long boom (188.9 inches) that was unfurled and
      extended automatically after launch. The magnetometer boom
      is located 240 degrees from the X-axis of the spacecraft
      coordinate system, measured in towards the Y-axis (in the
      spin direction) of the spin plane (XY). The total distance
      from the end of the boom to the orbiter spin axis is 5.94
      meters (234.0 inches)

  Reference
  =========
    Journal:          IEEE TRANSACTIONS ON GEOSCIENCE AND REMOTE SENSING
    Publication Date: 1980
    Reference Key ID: NOTHWANG1980

    Authors
    -------
      GEORGE J. NOTHWANG

    Citation
    --------

      Nothwang, G.T., `Pioneer Venus Spacecraft Design and
      Operation', IEEE Transactions on Geoscience and Remote
      Sensing, January 1980, Vol GE-18, No. 1, p5-10

   
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