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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 _________________________________________________________________