Instrument Host Overview    ========================      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.      Extracted from:      NOTHWANG, George J.,'Pioneer Venus Spacecraft Design and      Operation', IEEE Transactions on Geoscience and Remote Sensing,      vol.  GE-18, No.  1, January 1980    Platform Descriptions    =====================      Also from [NOTHWANG1980]:      ORBITER      -------      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 Subystem, Data Handling Subsystem,      Command Subsystem, Communications Subsystem, and Power Subsystem.        Mechanical        ----------        The mechanical features of the spacecraft can be described by six        basic assemblies. 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.        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).        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 of 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 of 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.        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), science	     roll reference signals generation, and spin period	     measurement;          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.        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.        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 16384-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).        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.        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 of 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.        Power        -----        The power subsystem provides semiregulated 28V +/- 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        overcurrent 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.