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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
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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
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GEORGE J. NOTHWANG
Citation
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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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