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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.
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