PDS_VERSION_ID = PDS3 RECORD_TYPE = STREAM LABEL_REVISION_NOTE = "original author/date unknown, suspect D. Simpson ~1993; Carol Polanskey, Oct 1998 - added info on S/C safings; Carol Polanskey, Oct 1999 - added GEM S/C safings; Dick Simpson, Jan 2000 - formatted for 72-byte lines; omitted internal references; Steve Joy, Dec 2003 - Updated s/c safings table with final info from Laura Bernard." OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = "GO" OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "GALILEO ORBITER" INSTRUMENT_HOST_TYPE = "SPACECRAFT" 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]." 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