PDS_VERSION_ID = PDS3
LABEL_REVISION_NOTE = "R. SIMPSON, 1997-11-18;
S. SLAVNEY, 1998-09-10;
R. SIMPSON, 1999-01-20;
R. SIMPSON, 1999-04-30;
M. CAPLINGER, 2000-03-30;
R. SIMPSON, 2007-07-18"
RECORD_TYPE = FIXED_LENGTH
RECORD_BYTES = 72
OBJECT = INSTRUMENT_HOST
INSTRUMENT_HOST_ID = "MGS"
OBJECT = INSTRUMENT_HOST_INFORMATION
INSTRUMENT_HOST_NAME = "MARS GLOBAL SURVEYOR"
INSTRUMENT_HOST_TYPE = "SPACECRAFT"
INSTRUMENT_HOST_DESC = "
Instrument Host Overview
========================
For most Mars Global Surveyor experiments, data were collected
by instruments on the spacecraft. Those data were then relayed
via the telemetry system to stations of the NASA Deep Space
Network (DSN) on the ground. Radio Science experiments (such
as radio occultations) required that DSN hardware also
participate in data acquisition. The following sections
provide an overview first of the spacecraft and then of the
DSN ground system as both supported Mars Global Surveyor
science activities.
Instrument Host Overview - Spacecraft
=====================================
The Mars Global Surveyor (MGS) spacecraft was built by
Lockheed Martin Astronautics (LMA). The spacecraft structure
included four subassemblies: the equipment module, the
propulsion module, the solar array support structure, and the
high-gain antenna (HGA) support structure.
The equipment module housed the avionics packages and science
instruments. Its dimensions were 1.221 x 1.221 x 0.762 meters
in X, Y, and Z, respectively. With the exception of the
Magnetometer, all of the science instruments were bolted to
the nadir equipment deck, mounted above the equipment module on
the +Z panel. The Mars Relay antenna was the tallest
instrument rising 1.115 meters above the nadir equipment deck.
Inside, two identical computers orchestrated almost all of
the spacecraft's flight activities. Although only one of the
two units controlled Surveyor at any one time, identical
software ran concurrently in the backup unit in case of an
emergency. Each computer consisted of a Marconi 1750A
microprocessor, 128 Kbytes of RAM for storage, and 20 Kbytes
of ROM that contained code to run basic survival routines in
the event that the computers experienced a reset.
Additional storage for science and spacecraft health data
was provided by two solid-state recorders with a combined
capacity of 375 megabytes. Mars Global Surveyor was NASA's
first planetary spacecraft to use RAM exclusively (instead of
a tape recorder) for mass data storage. This technological
improvement reduced operational complexity and cost.
The equipment module also housed three 'reaction wheels'
mounted at right angles to each other. By transferring angular
momentum to and from the rapidly spinning reaction wheels, MGS
flight computers could control the spacecraft attitude to high
precision. A fourth reaction wheel, mounted in a direction
skewed to the other three, provided redundancy and backup.
Sun sensors were placed at several locations about the
spacecraft. They provided basic information on spacecraft
attitude -- namely, a rough vector toward the Sun. Their
primary use was during attitude reinitialization after a
spacecraft anomaly.
The Inertial Measurement Unit (IMU) contained gyroscopes
and accelerometers to measure angular rates and linear
accelerations. Angular rate measurements were used to
determine yaw attitude during the Mapping Phase. The IMU
also provided inertial attitude control, as might be
required during maneuvers.
The Mars Horizon Sensor Assembly (MHSA) determined the horizon
as seen from the spacecraft; from this, an empirical nadir
could be derived for pointing the science instruments. The
MHSA was mounted to the +Z panel of the equipment module, next
to the science instruments.
The Celestial Sensor Assembly (CSA) complemented the IMU by
providing attitude data based on determination of positions
of known stars. It was used during the Cruise Phase and Orbit
Insertion Phase for both attitude determination and control.
It was also used when precise attitude knowledge was required
during the Mapping Phase. The CSA was mounted to the +Z panel
of the equipment module, next to the science instruments.
The propulsion module contained the propellant tanks, main
engines, propulsion feed system and attitude control
thrusters. It was a rectangular box 1.063 meters on a side
and was bolted to the equipment module on the latter's -Z
panel. The propulsion module also served as the adaptor to
the launch vehicle.
Propulsion was provided by a dual mode bi-propellant system
using nitrogen tetroxide (NTO) and hydrazine. This dual mode
differed from conventional bi-propellant systems in that the
hydrazine was used by both the main engine and the attitude
control thrusters, rather than having separate hydrazine
tanks for each. The main engine was the only one that used
the bi-propellant system. The main engine maximum thrust
was 659 N. It was used for major maneuvers including large
trajectory corrections during Cruise, Mars orbit injection
(MOI), and transfer to the Mapping orbit (TMO).
Four rocket engine modules (REM), each containing three 4.45 N
thrusters, were provided. Each REM contained two aft-facing
thrusters and one roll control thruster. Four of the eight
aft-facing thrusters were used for the smaller trajectory
corrections during Cruise and for Orbit Trim Maneuvers (OTM)
during Mapping; they could also be used for attitude control
during main engine burns. Two sets of four thrusters were on
redundant strings so that one string could be isolated in the
event of a failure. Four thrusters were provided for
attitude control. In addition to their role during maneuvers,
the 4.45 N thrusters were also used for momentum management.
MGS carried about 385 kg of propellant; nearly 75 percent of
that was used during MOI.
Two solar arrays, each 3.53 meters long by 1.85 meters wide
provided power. Each array was mounted close to the top of
the propulsion module on the +Y and -Y panels and near the
interface between the propulsion and equipment modules.
Including the adaptor that held the array to the propulsion
module, the tip of each array was designed to stand 4.270
meters from the side of the spacecraft. During initial
deployment, the -Y solar array yoke was damaged leaving its
exact position and orientation in some doubt (and leading to
several changes in mission design). Rectangular, metal
'drag flaps' were mounted to the end of each array; these
flaps increased the total surface area of the structure and
added another 0.813 meters to the overall dimensions.
Between each array and flap was mounted a magnetometer sensor.
Each array consisted of two panels, an inner and outer panel,
of gallium arsenide and silicon solar cells,
respectively. During mapping operations at Mars, the amount
of power produced by the arrays varied from a high of 980
Watts at perihelion to a low of 660 Watts at aphelion.
While in orbit around Mars, the solar arrays provided
power as MGS flew over the day side of the planet. When
the spacecraft passed over the night side, energy flowed
from two nickel-hydrogen (NiH2) batteries, each with a
capacity of about 20 Amp-hours. Eclipses lasted from 36 to
41 minutes per orbit; depth of battery discharge was limited
to 27% except during emergencies.
The high-gain antenna structure was also bolted to the
outside of the propulsion module. When fully deployed, the
1.5-meter diameter antenna sat at the end of a 2-meter boom
which was mounted to the +X panel of the propulsion module.
Two rotating joints (gimbals) held the antenna to the boom
and allowed the antenna to track and point at Earth while
the science instruments observed Mars.
One of the two main functions of the HGA was to receive
command sequences sent by the flight operations
team on Earth. During command periods, data flowed to MGS
at rates in multiples of two from 7.8125 bits per second
(emergency rate) to 500 bits per second (750 commands per
minute); the nominal rate was 125 bits per second.
The other main function of the HGA was to send data back
to Earth. All transmissions from MGS utilized an X-band
radio link near 8.4 gigahertz. The transmitted power was
about 25 watts. Data rates as high as 85333 bits per second
were used.
The spacecraft was also equipped with four low-gain antennas
(LGA), two for transmit and two for receive. The LGAs were
used in Inner Cruise, during special events such as maneuvers,
during aerobraking, and for emergency communications following
a spacecraft anomaly.
The primary transmitting low-gain antenna (LGT1) was mounted
on the traveling wave tube amplifier (TWTA) enclosure, which
was mounted on the rim of the HGA reflector; its boresight
was aligned with the HGA boresight, which was in the +X
direction until HGA deployment. The backup (LGT2) was also
mounted on the TWTA enclosure. LGT2 boresight was aligned
at a cant angle approximately 160 degrees away from the
shared boresights of the HGA and LGT1. This angle was chosen
to minimize the consequences of a gimbal failure once
articulation commenced after deployment of the HGA boom in
mapping orbit. LGT2 was not used prior to HGA deployment
because its orientation and proximity to the nadir payload
deck would lead to irradiation of the payload instruments
while the HGA was in its stowed position. One receiving LGA
(LGR) was mounted on the -X panel of the equipment module;
the other was on the +X side of the propulsion module.
The spacecraft was equipped with an experimental Ka-band
downlink radio system. The transmitter converted the X-band
signal to 32 Ghz and amplified it to about 0.5 watts; the
Ka-band output was radiated through the HGA.
The spacecraft +Z axis vector was normal to the nadir equipment
deck; the main engine was aimed in the -Z direction. The -X
axis vector was in the direction of the velocity vector during
nominal Mapping (e.g., May 1999). +X was in the direction of
the HGA boresight during Cruise, and the HGA boom was mounted to
the +X panel of the propulsion module. The +Y axis completed an
orthogonal rectangular coordinate system. The +/-Y axes defined
generally the deployment directions of the solar panels. The
solar cells themselves were on the -Z sides of the panels.
There were three levels of anomaly response in the spacecraft
flight software. The first, emergency mode, was entered in
response to a command-loss timeout. Entry into emergency mode
reconfigured the telecom subsystem to its lowest data rate
settings to enhance the chances of successful contact from
Earth. After a programmable period of time in emergency mode,
the spacecraft transitioned to contingency mode.
Contingency mode was entered by four paths: failure to regain
contact with Earth while in emergency mode, power-related faults
such as gimbal faults and low battery state of charge, loss of
inertial reference, and explicit ground command. Contingency
mode set telecom rates to their minimum values, turned off
non-essential power loads (including the payload), disabled
stored sequences not explicitly specified as enabled for this
mode, and changed the spacecraft attitude to sun-coning to
optimize power and communications.
Safe mode was the deepest level of anomaly response. It could
be entered by three paths: failures of key spacecraft components
that could not be corrected by normal fault protection, power-on
reset of both Spacecraft Control Processors (SCPs), or explicit
ground command. The response to safe mode entry was similar to
that of contingency mode. Safe mode program code for the SCP
was executed from Programmable Read-Only-Memory (PROM).
For more information on the spacecraft and mission see
[JPLD-12088].
Instrument Host Overview - DSN
==============================
The Deep Space Network is a telecommunications facility managed
by the Jet Propulsion Laboratory of the California Institute of
Technology for the U.S. National Aeronautics and Space
Administration (NASA).
The primary function of the DSN is to provide two-way
communications between the Earth and spacecraft exploring the
solar system. To carry out this function it is equipped with
high-power transmitters, low-noise amplifiers and receivers,
and appropriate monitoring and control systems.
The DSN consists of three complexes situated at approximately
equally spaced longitudinal intervals around the globe at
Goldstone (near Barstow, California), Robledo (near Madrid,
Spain), and Tidbinbilla (near Canberra, Australia). Two of
the complexes are located in the northern hemisphere while the
third is in the southern hemisphere.
Each complex includes several antennas, defined by their
diameters, construction, or operational characteristics:
70-m diameter, standard 34-m diameter, high-efficiency 34-m
diameter (HEF), and 34-m beam waveguide (BWG).
For more information see [ASMAR&RENZETTI1993]. "
END_OBJECT = INSTRUMENT_HOST_INFORMATION
OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
REFERENCE_KEY_ID = "ASMAR&RENZETTI1993"
END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
REFERENCE_KEY_ID = "JPLD-12088"
END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
END_OBJECT = INSTRUMENT_HOST
END
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