| INSTRUMENT_HOST_DESC |
Instrument Host Overview
========================
For most Galileo Orbiter 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 also required
that DSN hardware participate in data acquisition on the
ground. The following sections provide an overview, first
of the Orbiter and then of the DSN ground system as both
supported Galileo Orbiter science activities.
Instrument Host Overview - Spacecraft
=====================================
Launched 1989-10-18 by the Space Shuttle Atlantis, Galileo
was the first spacecraft to use a dual-spin attitude
stabilization system. The rotor (or spun section) turned at
approximately three revolutions per minute while the stator
(or despun section) maintained a fixed orientation in space.
This design accommodated the different requirements of remote
sensing instruments (mounted on the stator) and fields and
particles instruments (mounted on the rotor); spacecraft
engineering subsystems were also mounted on the rotor. The
rotor and stator were connected by a spin bearing assembly,
which conducted power via slip rings and data signals via
rotary transformers.
There were eleven subsystems and nine scientific instruments
on the orbiter. The spacecraft power source was a pair of
radioisotope thermoelectric generators. Propulsion was
provided by a bipropellant system of twelve 10-newton
thrusters and one 400 newton engine. The command and data
subsubsystem consisted of multiple microprocessors and a
high-speed data bus. The telecommunications subsystem was
designed to transmit data to Earth at rates ranging from
10 bps to a maximum of 134 kilobits per second at S-band
and X-band frequencies. The rotor had one 4.8 meter high-gain
antenna and two low-gain antennas, but the high-gain antenna
never deployed properly so data were returned from Jupiter at
rates far below the design maxima using the low-gain antennas.
The stator contained a radio relay antenna operating at L band
for receiving data from the atmospheric probe, which is
described elsewhere.
Science instruments fell into two general categories. Remote
sensing instruments included:
PPR Photopolarimeter Radiometer
NIMS Near-Infrared Mapping Spectrometer
SSI Solid State Imaging Camera
UVS/EUV Ultraviolet Spectrometer/Extreme Ultraviolet
Spectrometer
Instruments primarily designed for 'in situ' measurements
included:
EPD Energetic Particles Detector
DDS Dust Detector Subsystem
PLS Plasma detector
PWS Plasma Wave Subsystem
MAG Magnetometer
The Heavy Ion Counter (HIC) is an engineering subsystem which
was added to the spacecraft to monitor high energy ions, but
it is also being used to collect science data.
The two Radio Science (RSS) experiments, Celestial Mechanics
and Propagation, were conducted using equipment on both the
Orbiter and on the ground.
The mass of the Orbiter at launch was 2223 kg, of which 925 kg
was usable propellant. The Orbiter payload mass was 118 kg.
Orbiter height was 6.15 m.
Overall project management for Galileo was provided by the
California Institute of Technology's Jet Propulsion Laboratory
in Pasadena, California, which also built the orbiter. Ames
Research Center in Mountain View, California, was responsible
for the development of the probe, which was supplied by Hughes
Aircraft Company and the General Electric Company. The Federal
Republic of Germany provided the orbiter's main propulsion
system, one complete scientific instrument one the orbiter
(DDS), another on the probe (HAD), and major elements of others.
For more information see [YEATESETAL1985; DAMARIOETAL1992]
Platform Descriptions
---------------------
The Rotor was the spinning section of the Galileo Orbiter
and represented most of the spacecraft mass; it carried the
high-gain communications antenna, the propulsion module,
flight computers, and most support systems. Two booms were
attached to the Rotor; each was unfurled and extended
automatically after launch. The science boom extended to a
distance of three meters from the spacecraft centerline;
to it were mounted the EPD, DDS, HIC, and PLS instruments.
The magnetometer boom extended outward eleven meters from
the centerline and was attached to the science boom. It
carried the PWS antenna and two MAG sensors, one at the
midpoint of the boom and the other at its outboard end.
The EUV spectrometer was mounted on the Rotor bus. For
more information see [YEATESETAL1985; DAMARIOETAL1992]
The Stator was the despun section of the Orbiter. It was
turned via an electric motor opposite to the rotation of the
Rotor, so that it maintained a stable orientation in space.
Attached to the Stator was a moveable scan platform which
contained the remote sensing instruments: PPR, NIMS, SSI,
and UVS. The Probe and the Probe relay antenna were also
attached to the Stator. For more information see
[YEATESETAL1985; DAMARIOETAL1992].
The Rotor and Stator were connected by a spin bearing
assembly (SBA), which conducted power via slip rings and
data signals via rotary transformers.
Telecommunications Subsystem
----------------------------
The Telecommunications Subsystem was located in the Rotor
section of the Orbiter. It included elements for receiving
uplink command signals and for transmitting downlink
telemetry. The uplink portion of the system received radio
signals with command data at 2115 MHz and demodulated,
detected, and routed those to the Command and Data System
(CDS). The downlink portion received telemetry data from
the CDS and was designed to modulate S-band and X-band
carriers at 2295 and 8415 MHz, respectively, at data rates
as high as 134.4 kilobits per second (kbps).
A 4.8 meter umbrella-like high-gain antenna (HGA) and two
low-gain antennas (LGAs) were mounted on the Rotor. The
LGAs operated only at S-band. One was mounted on a boom
and was included primarily to improve Galileo's
telecommunications during the flight to Venus (while
the heat-sensitive HGA remained furled). The other LGA was
mounted at the top of the HGA. The Stator contained a radio
relay antenna operating at L-band for receiving Probe data
during its atmospheric entry.
On 1991-04-11 the HGA was commanded to unfurl; but telemetry
showed that the motors had stalled with the ribs only partly
deployed. Months of tests and simulations followed, but
without further progress in opening the antenna. Engineers
deduced that the problem most likely resulted from sticking
of a few antenna ribs, caused by friction between their
standoff pins and sockets. The excess friction resulted from
etching of surfaces after dry lubricant, bonded to the standoff
pins during manufacture, was shaken loose during pre-launch
transport.
The mission was conducted using the LGA mounted on top of the
HGA (the boom-mounted LGA was stowed after its service en
route to Venus had been completed). Without adaptations, the
LGA data transmission rate at Jupiter would have been limited
to only 8-16 bits per second (bps), compared to the HGA's
134.4 kbps. Onboard software changes, coupled with hardware
and software changes at Earth-based receiving stations,
increased the data rate from Jupiter by as much as 10 times,
to 160 bps.
'Lossless' data compression allows data to be recovered
exactly, once they have been received on the ground. 'Lossy'
data compression allows controlled corruption of the data
through mathematical approximations but with significant
increases in transmission rate. Lossy compression was used
with Galileo Orbiter imaging and plasma wave data to reduce
volumes to as little as 1/80th of their original volumes.
On the ground S-band communications capabilities were upgraded
at the Canberra DSN tracking station (because Jupiter was at
southern declinations during most of the Galileo tour,
Canberra received more data from the Orbiter than the other
DSN stations). 'Block V' receivers were installed at all
stations; these could operate without need for a residual
carrier, meaning all of the spacecraft radiated power could be
assigned to carry its modulation. Early in the tour, arraying
of 34-m antennas with the 70-m antenna at each site was
implemented; arraying of pairs of 70-m antennas and arraying
with the 64-m CSIRO antenna at Parkes (Australia) were also
used to increase data rates.
The TCS as designed would have provided a dual channel
downlink. The high-rate channel would have provided a
convolutionally coded, pulse-code modulated microwave channel,
while a low-rate channel data was uncoded. Downlink
transmission of telemetry data would have been possible at
S-band and/or X-band over a wide range of selectable data
rates, including 134 and 115.2 kbps at Jupiter.
Approximately 160 W (33 percent of total available) was
provided for the combined S-band and X-band communications
function. Dual power level, traveling wave tube amplifier
transmitters were to provide maximum S-band cruise data return
and high-rate X-band data return from Jupiter while
simultaneously satisfying dual-frequency tracking and
radio science requirements.
Several other features were incorporated in the
telecommunications area, mainly to enhance radio science and
navigation. A noncoherent tracking mode was available which
permitted the Orbiter to be commanded while the downlink
frequency source was controlled by an auxiliary oscillator or
an ultrastable oscillator -- providing short-term frequency
stability of better than 5 parts in 10^12. A differential
downlink-only ranging mode was also available using one
S-band and three X-band sine wave tones modulated onto the
downlinks to enhance navigational accuracy. A single X-band
to S-band down-converter receiver was available for receiving
X-band uplink signals to enhance radio science and the search
for gravity waves. These X-band capabilities were never used,
however, because X-band was only available through the high
gain antenna. The capability existed to completely remove all
telemetry modulation from the downlink carriers, thus
maximizing atmospheric penetration depth during Earth
occultations.
Propulsion Subsystem
--------------------
The Galileo Retropropulsion Module (RPM system), located on
the Rotor platform of the Orbiter, was supplied by the Federal
Republic of Germany. It was based on earlier bipropellant
Symphonie designs.
The Propulsion Subsystem provided all directed impulse for
attitude control, trajectory correction, and Jupiter orbit
insertion. The propulsion functions consisted of spin rate
control, fine turning to point the HGA to Earth, and
orientation of the spacecraft for propulsive or science
maneuvers.
The RPM included four propellant tanks (two fuel tanks
containing monomethylhydrazine and two oxidizer tanks
containing nitrogen tetroxide), two helium pressurant tanks,
twelve 10-N thrusters (six each mounted on separate
cantilevered booms), one 400-N engine, and necessary isolation
and control elements. At launch, the system was fully loaded
with 932 kg of usable propellant and weighed about 1145 kg.
Four of the 10-N thrusters were mounted in a direction to
provide a functional backup for the 400-N engine. The
thrusters were mechanized on two separate branches providing
redundancy for spin control, HGA pointing, and trajectory
correction. The 400-N engine was used three times -- all
subsequent to Probe separation.
Control of propellant to the 10-N thrusters and the 400-N
engine was accomplished by opening and closing fuel and
oxidizer solenoid latch valves via electrical signals from
the attitude control system propulsion drive electronics.
The propulsion drive electronics also provided the control
signals for opening and closing the thruster and
400-N engine valves.
Command, Telemetry, and Data Handling Subsystem
-----------------------------------------------
Primary command, control, and data handling was performed
by the actively redundant Command and Data Subsystem (CDS).
Its major functions included receiving and processing
real-time commands from Earth and forwarding them to
appropriate spacecraft subsystems, executing sequences of
stored commands (either as part of a normal preplanned
flight activity or in response to the actuation of various
fault recovery routines), controlling and selecting data
modes, and collecting and formatting science and engineering
data for downlink transmission. The CDS architecture used
multiple microprocessors and a high-speed data bus for both
internal and user communication.
A majority of the CDS electronics were located on the Orbiter
Rotor platform in proximity to the data storage, science, and
telecommunications equipment. CDS Stator elements were
limited to those necessary to support the Probe and relay
radio hardware equipment, the remote sensing instruments
mounted on the scan platform, the launch vehicle, and sequence
operations. Six 1802 microprocessors, memory units, and the
data bus comprised the 'heart' of the CDS. Four of the
microprocessors (two high-level modules and two low-level
modules) and four memory units contained a total of 144000
words of random access memory (RAM) and were located on the
Rotor platform along with supporting electronics. The
low-level modules of the remaining two microprocessors, each
with 16K RAM, were located on the Stator platform. The data
bus comprised three dedicated busses. The bus interface was
used by all data systems -- that is, Orbiter science, the
attitude and articulation control subsystem, and relay radio
hardware receivers.
Interfacing between Rotor and Stator portions of the CDS was
accomplished via slip rings and rotary transformers mounted
on the spin bearing assembly. Efficient and effective
communication among data systems was accomplished using a
specifically defined protocol structure and real-time
interrupt time slicing. The protocol addressing schemes
provided for either a relatively simple bus adapter that
relied on direct memory access by the user's processor or a
more complex bus adapter with direct memory access capability
independent of the processor.
Attitude and Articulation Control Subsystem
-------------------------------------------
The Attitude and Articulation Control Subsystem (AACS) was
responsible for maintaining spin rate of the spacecraft;
orienting the spin vector; controlling propulsion isolation
valves, heaters, 10-N thruster firing, and 400-N engine
firing; and controlling the science platform containing the
remote sensing instruments on the Stator platform.
Design of the AACS was profoundly influenced by science
requirements and the various spacecraft operational
configurations that had to be accommodated. Configurations
included the basic cruise dual spin configuration (Orbiter
with Probe), dual spin without the Probe (for orbital
operations) and 'all spin' configurations with and without the
Probe for trajectory corrections at spin rates from 3 to 10
rpm.
The AACS incorporated many functional elements to meet the
demanding performance, lifetime, and reliability requirements
of the mission. The majority of the AACS functional elements
were block redundant and located on the Rotor platform.
Stator elements included those necessary for controlling the
pointing and slewing of the scan platform, pointing the relay
antenna, and interfacing with the Rotor section electronics.
The central element of the AACS was the attitude control
electronics (ACE) package that controlled the AACS
configuration; monitored its health; performed executive,
telemetry, command, and processing functions; provided spin
position data to other subsystems; and provided AACS fault
recovery. The 'heart' of the ACE was a high-speed 2900
ATAC-16 processor and memory containing 31K words of 16-bit
RAM and 1K words of 16-bit read-only memory (ROM).
ROM storage was used only for those functions required
to safeguard the science instruments, switch to the
low-gain antenna, and Sun point the Orbiter to permit
ground commanding. Activation of the ROM sequences
occurred only when a loss of RAM was detected.
The ACE also contained electronics necessary to interface with
AACS peripheral elements in the Rotor section, the Stator
electronics, and the CDS. Interfacing between Rotor and
Stator AACS elements was accomplished via rotary transformers
located on the Spin Bearing Assembly (SBA).
Other major AACS functional elements included:
- a radiation hardened star scanner employing photomultiplier
tubes for star field identification during in-flight attitude
determination
- linear actuators for raising or lowering the RTG booms to
reduce wobble and maintain stability
- acquisition sensors for attitude determination, spin rate
sensing during launch, and Sun acquisition
- propulsion drive electronics to control the RPM latch valve,
thrusters, and 400-N engine valves
- a spin bearing assembly to provide the mechanical and
electrical interface between Rotor and Stator sections of
the Orbiter as well as to provide despun orientation
- gyros mounted on the Stator scan platform to control platform
articulation and stabilization.
- accelerometers mounted on the Stator platform diametrically
opposite to each other and aligned parallel to the Orbiter
spin axis to measure velocity changes during propulsive burns
- a scan actuator subassembly to provide scan platform cone
actuation and positioning information.
After launch vehicle separation and RPM pressurization, the
spacecraft assumed the 'all-spin' configuration. This was
used frequently during the mission and for all propulsive
maneuvers to provide stabilization. In all-spin configuration
for 10-N thruster burns, the entire Orbiter would spin at
roughly 3 rpm; for 400-N engine burns, the Orbiter would
spin at 10 rpm. This configuration was also used during
science calibration target observations by the remote sensing
science instruments.
For most of the mission, the AACS operated in the cruise mode,
in which the Orbiter operated in the dual-spin configuration
with the Rotor platform inertially fixed. Major AACS
functions performed in this mode were wobble control, high-gain
antenna pointing, attitude determination, and spin rate control.
The final AACS mode was the inertial mode. Transition to this
mode was from the cruise mode with gyros active. While in this
mode the AACS performed functions such as closed-loop commanded
turns using the RPM thrusters, accurate pointing and slewing of
the scan platform, and closed-loop control for wobble angle
compensation.
Electric Power Subsystem
------------------------
Electrical power was provided to Galileo's equipment by two
radioisotope thermoelectric generators. Heat produced by
natural radioactive decay of plutonium 238 dioxide was
converted to electricity (570 watts at launch, 485 watts at
the end of the mission) to operate the Orbiter equipment for
its eight-year baseline mission. This was the same type of
power source used by the two Voyager spacecraft missions to
the outer planets, the Pioneer Jupiter spacecraft, and the
twin Viking Mars landers.
Spacecraft Coordinate Systems
-----------------------------
The Rotor coordinate system consisted of three mutually
perpendicular axes: Xr, Yr, and Zr. The Zr axis was
nominally parallel to the spin bearing assembly (SBA) axis
and passed through the center of the Rotor with +Zr directed
opposite to the HGA boresight direction. +Yr was normal to
Zr and was directed toward the science boom. +Xr was normal
to both Yr and Zr and formed a right-handed system. The
angular momentum vector for the spinning spacecraft was in
the +Zr direction.
\ / HGA
\ /
\ /\ /
------------
| ROTOR |-------------------\ Science and MAG
| |-------------------/ Boom
------------
SBA |
| ---> +Yr
+Zr
The Stator coordinate system consisted of three mutually
perpendicular axes: Xs, Ys, and Zs. The Zs axis was
nominally parallel to the SBA axis and passed through the
center of the Stator with +Zs directed opposite to the HGA
boresight direction (+Zs was parallel to +Zr). +Ys was normal
to Zs and was directed opposite to the scan platform direction.
+Xs was normal to both Ys and Zs and formed a right-handed
system.
SBA |
------------
| STATOR |-------------------\ Scan
| |-------------------/ Platform
------------
|
+Ys <--- |
+Zs
-Zr,-Zs
|
| /
| __(o)-._
| _.--_/\/' -
....- _/\/'
__---__ _/\/'
'-_/|\_-` _/\/'
__|]]_ _(o)'
__---- /|||\----__ _/\/' +Yr,-Ys
_--\ __----------__ /--_ _/\/' /
/ _--\ __|___ /--_ \/\/' /
\-/ __-\- | /-- \/\/' /
`\--/--___\-|-/___-\-///' /
,_`-`---| |___| |__/\/' /
,--/---===_/||\ -`---(o) /
,/--/ ,-, ,--('||))|---|)\|\
,/--/ |]]=\== \_|/ |___]-)\|\,--
/--/: '-' `__-------_=]= \|[[[
[=[=/! : [_-------_\== \[[[
' //_-- --_[=-- [-_ ---------- +Xr, -Xs
-Xr,+Xs ------- ---`\ /[_]' \/_\_
/'|`\[|`\_ //' [ ]=
`-[-'[]_] - [___]=]
---
/ |
/ |
/ |
/ |
-Yr,+Ys |
+Zr,+Zs
Figure - Perspective view of Galileo Orbiter spacecraft (Should
be viewed in a mono-spaced font such as Courier)
The scan platform coordinate system consisted of three mutually
perpendicular axes: L, M, and N. The platform had a primary
mounting plane which was established by three mounting points
on the platform. Two reference pins (Pin 1 and Pin 2) were
installed on the primary mounting plane to establish platform
alignment. The origin of the coordinate system was at the
intersection of the center line of Pins 1 and 2 and the primary
mounting plane. The coordinate axis L, defining look direction,
was parallel to the SSI instrument and passed through the center
line of Pins 1 and 2. Coordinate axis M was in the primary
mounting plane, perpendicular to L, and passing through the
origin. Axis N was mutually perpendicular to both L and M such
that L = M x N. Individual instruments were assigned
subscripted Li, Mi, Ni coordinate systems such that an
instrument pointing vector was specified by direction cosines
of its coordinate axes Li, Mi, Ni with respect to the platform
coordinates L, M, N.
Spacecraft Safing Summary
-------------------------
Throughout the mission there have been a number of occasions
when the spacecraft detected a fault condition onboard and
configured itself to a safe state. At that time, all onboard
sequences are cancelled, and a number of science instruments
are powered off. The following table lists the time of these
'safing' events, which stored sequence was aborted, and the
reason that the spacecraft entered its fault protection
routines. The times of the events have been extracted from
different sources. Some times are known exactly and others
have uncertainties of up to 5 minutes. The most uncertain
times are indicated with an *.
Date SCET (UTC) SEQ Cause of safing
1990-01-15 90-015/22:52* EV-5 star scanner calibration
1991-03-26 91-085/13:31:18 VE-14 B-string CDS bus reset
1991-05-03 91-123/05:26 n/a A-string CDS bus reset
1991-07-20 91-201/02:09:00 n/a A_string CDS bus reset
1993-06-10 93-161/16:53:05 EJ-1 A-string CDS bus reset
1993-06-17 93-168/18:22:04 n/a A-string CDS bus reset
1993-07-10 93-191/20:16:58 EJ-2 A-string CDS bus reset
1993-07-12 93-193/01:37* n/a A-string CDS bus reset
1993-08-11 93-223/22:04:40 EJ-2' A-string CDS bus reset
1993-09-24 93-267/14:14:54 EJ-3 A-string CDS bus reset
1994-09-14 94-257/03:10:51 EJ-7B DMSMRO memory failure
1994-09-16 94-259/16:38* n/a CAP privileged error
1995-02-04 95-035/17:44:39 n/a Phase 1 In-Flight Load-planned
1996-01-05 96-005/21:51:12 J0C-A SITURN cmd constr. violation
1996-05-18 96-139/01:26* n/a Phase 2 In-Flight Load-planned
1996-08-24 96-237/15:30:32 G01-C timing overrun from DACs
1997-12-22 97-356/16:52* E12BHG AACS Anomaly
1998-05-28 98-148/20:21:26 E14BGD Safing during OTM-47
1998-07-20 98-201/17:35:46 E16AKE Despun BUS POR
1998-11-22 98-326/05:24:13 E18AFE Simultaneous 2 string CDS bus reset
two resets: 98-326/05:24:13.102 and
98-327/01:29*
1998-12-09 98-343/17:05:10 E18BFE Sequence stopped by B18T24 RBS
1999-02-01 99-032/05:41:33 E19AHC SUNACQ Failure
1999-10-10 99-283/09:17:06 I24AGE B-String CDS bus reset
1999-11-26 99-330/22:00:02 I25ADF B-String code error in box 5
start ADD
2000-02-24 00-055/12:00:13 I27ADC A&B string CDS bus reset
2002-01-17 02-017/13:41:09 I33AFE A-string CDS bus reset (parity err)
2002-02-16 02-047/20:51:00 I33BED A-string CDS bus reset
2002-10-02 02-275/03:41:22 I33EDE Commanding Error
2002-11-05 02-309/06:35:36 A34AHG Radiation Failure
The most common cause of spacecraft safing was from a CDS despun
bus reset of either the A-string or B-string. It has been
determined by analysis that there has been current leakage
somewhere in the spacecraft power bus, and that the resulting
bus imbalances are most likely caused by brush debris forming
high-resistance leakage paths across the brush armatures in the
spin bearing assembly. These paths are formed and then
'blown open' before the resistance becomes low enough to permit
significant current flow. In some cases the brush was 'lifted'
briefing while debris paths were causing power to 'touch' the
brush and this tripped a reset signal in the CDS. Onboard fault
protection 'safes' the spacecraft when the reset trips
[ONEIL1991]. No damage has occurred on the spacecraft as a
result of these trips, but the spacecraft operations are
disrupted until the onboard sequences and spacecraft state can
be restored from the ground. In April of 1999 a change was made
to the CDS flight software that allows it to detect and
autonomously recover from despun bus resets. With this new
software enabled, the CDS strings do not 'go down', 'safing'
does not execute and the onboard sequences continue.
On September 13, 1994 a memory cell in the CDS failed during the
playback of Shoemaker-Levy 9 recorded data and resulted in
spacecraft safing to be entered twice. After 12 days the
spacecraft was reconfigured back to normal operations. The
failed memory cell was located in a bulk storage (DBUM-1A)
module of the CDS, and was only used during tape recorder/memory
readout playbacks and other short term storage of data
[ONEIL1995].
Following the successful insertion into Jupiter orbit in
December 1995, a spacecraft turn was attempted on January 5,
1996. The spacecraft was in a non-standard configuration
following the JOI maneuver which resulted in an incompatibility
between the turn design and the spacecraft state. The
spacecraft entered safing, but was recovered shortly afterwards.
On August 24, 1996 the spacecraft went into safing due to a
timing overrun condition in the CDS, ending any further data
return from the G1 encounter. The timing overrun was traced
to the transmission of 4 Delayed Action Commands which stressed
the limits of the CDS running the new Phase 2 flight software.
By September 1, the spacecraft had been returned to normal
operations and the G2 encounter sequence began on schedule
[ONEIL1996].
Twice during the Prime Mission, during the loading of new
flight software for Phase 1 and Phase 2, the spacecraft was
purposely commanded to trigger the safing response in order to
put all subsystems in a known state prior to the load.
On May 28, 1998 the spacecraft entered safing for the first
time in the Galileo Europa Mission. Safing occurred during
the maneuver, OTM-47, inbound to the Europa 15 encounter. The
spacecraft executed the majority of the maneuver before a
sequence timing error created an AACS command constraint
violation which caused the spacecraft to abort the on-board
sequence and safe itself. The Science Virtual Machine was
recovered on 98-149, and a mini-sequence was uplinked to
turn on the science instruments and match the spacecraft
states to the E15A sequence.
On February 1, 1999, four hours after completing the close
approach science recordings, the spacecraft entered safing
during a sun acquisition turn designed to move the spacecraft
from the science data taking attitude back to the nominal
earth pointed attitude. It appears that the cause of the sun
acquisition halt was the result of a failure of the two
acquisition sensors to provide the complete overlap they
were design for.
On October 10, 1999 the spacecraft entered safing when high
radiation on approach to the Io 24 encounter caused an error
in the CDS B-string memory. The hardware error causing the
safing was a memory read error in the CDS B string High
Level Module - the 'executive controller' for the CDS B
string. Because the error was detected by the CDS bus
controller (and not the microprocessor), this is likely to
be an error in memory used for data buffers. Within 18 hours
of safing the I24 sequence was regenerated, loaded onboard,
and the 75% of the I24 encounter data was acquired.
During the extended mission five of the anomalies were caused by
CDS bus resets that were nominally handled with software changes
implemented previously.I33EDE where the spacecraft entered safing
on October 2, 2002 was due to commanding error on the ground
during fault protection changes (ISA 11007).
In A34A an anomaly occurred on November 5, 2002 at 06:19 UTC, when
the spacecraft flew within 160 km of the surface of Amalthea. The
speed of the spacecraft relative to Amalthea was approximately
18.4 kilometers per second (41,000 miles per hour), taking less
than 15 seconds to pass by. Approximately 17 minutes after
closest approach, the intensity of the radiation caused a failure
in computer circuitry that handles timing of the events on the
spacecraft. This caused the computer to switch to the CDS B-string
and go into safe mode. There were also several additional faults
which triggered repeated requests to place the spacecraft in safe
mode.
Instrument Host Overview - DSN
==============================
Galileo Radio Science investigations utilized instrumentation
with elements both on the spacecraft and at the NASA Deep Space
Network (DSN). Much of this was shared equipment, being used
for routine telecommunications as well as for Radio Science.
The Deep Space Network was a telecommunications facility
managed by the Jet Propulsion Laboratory of the California
Institute of Technology for the U.S. National Aeronautics and
Space Administration.
The primary function of the DSN was to provide two-way
communications between the Earth and spacecraft exploring the
solar system. To carry out this function the DSN was equipped
with high-power transmitters, low-noise amplifiers and
receivers, and appropriate monitoring and control systems.
The DSN consisted 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 were located in the northern hemisphere while
the third was in the southern hemisphere.
The network comprised four subnets, each of which included
one antenna at each complex. The four subnets were defined
according to the properties of their respective antennas: 70-m
diameter, standard 34-m diameter, high-efficiency 34-m diameter,
and 26-m diameter.
These DSN complexes, in conjunction with telecommunications
subsystems onboard planetary spacecraft, constituted the major
elements of instrumentation for radio science investigations.
For more information see [ASMAR&RENZETTI1993].
|
.png)
.png)
