PDS_VERSION_ID = PDS3 RECORD_TYPE = STREAM LABEL_REVISION_NOTE = "NULL" OBJECT = INSTRUMENT INSTRUMENT_HOST_ID = "CO" INSTRUMENT_ID = "MAG" OBJECT = INSTRUMENT_INFORMATION INSTRUMENT_NAME = "DUAL TECHNIQUE MAGNETOMETER" INSTRUMENT_TYPE = "MAGNETOMETER" INSTRUMENT_DESC = " ABSTRACT ======== The dual technique magnetometer system onboard the Cassini orbiter is described. This instrument consists of vector helium and fluxgate magnetometers with the capability to operate the helium device in a scalar mode. This special mode is used near the planet in order to determine with very high accuracy the interior field of the planet. The four-year orbiting Cassini mission will lead to a detailed understanding of the Saturn/Titan system. In addition to the prime scientific measurement of the planetary field, the instrument will also make measurements of the planetary magnetosphere, and the interactions of Saturn with the solar wind, of Titan with its environments, and of the icy satellites within the magnetosphere. INSTRUMENT DESCRIPTION ====================== The MAG instrument comprises a fluxgate magnetometer (FGM) and a vector helium magnetometer capable of operating in both vector and scalar mode (V/SHM). The instrument is intended to measure small changes in fields spanning four orders of magnitude with extremely high sensitivity. This goal is achieved in part by mounting the sensors on an 11-metre spacecraft boom; the V/SHM at the end of the boom, the FGM halfway along. The magnetometer boom distances the sensors from the magnetic field associated with the spacecraft and its subsystems, and especially from spacecraft-generated temporal field variations. Spacing the sensors at different distances along the boom allows the spacecraft fields to be better characterised and removed from the observations. However, mounting the sensors on a boom could result in their orientation with respect to the spacecraft axes changing from time to time, for example after spacecraft manoeuvres. A means of sensor-alignment determination has been provided by the Cassini project - the Science CAlibration Subsystem, SCAS. This system consists of two, perpendicular, coils rigidly mounted on the spacecraft body with a known alignment to the spacecraft axes. These coils produce well-defined magnetic fields on command which can be detected by the sensors and used to correct for any changes in sensor orientation. Both magnetometers are capable of measuring the magnetic-field vector at rates from 0 Hz up to 10 Hz (VHM) or at least 30 Hz (FGM). The VHM optimises low-frequency vector measurements in weak fields. The FGM is best suited to high-frequency measurements and can operate over an extremely wide dynamic range, from very weak fields up to Earth's strong field. The twin-sensor configuration contributes to overall instrument reliability; if one sensor fails, field measurements can be made with the other sensor, with sufficient performance to achieve many of the major objectives of the investigation. Reliability has been further increased by the provision of redundant instrument power supplies and data processing units, and by careful selection of electronic components that can survive the radiation environments encountered during the long cruise phase of the mission and in the Saturnian system. The VHM provides the stability needed to maintain calibrations, obtained in the solar wind, whilst Cassini is inside the Saturnian magnetosphere for long periods during the four-year tour. An implicit feature of scalar or resonance magnetometers are null zones which arise if the ambient field falls outside a cone of 45 degrees half angle with respect to the optical axis of the magnetometer. These null zones result in the signal being dramatically weakened, causing the absolute accuracy of the instrument to suffer. When Cassini is inside 4 RS a requirement has been placed on the mission to avoid spacecraft orientations which cause the planetary field to lie within the null zones of the SHM. Other features of the instrument that have been driven by the characteristics of the mission and by the design of the spacecraft are to be found in the data processing unit (DPU). The DPU contains a bus interface unit (BIU), provided by the Cassini project for interfacing to the onboard data handling subsystem (CDS) bus. In line with the spacecraft design, the DPU is capable of handling Packet Telemetry and Telecommands, and features a flexible telemetry storage and generation scheme to support the multiple telemetry modes of the spacecraft. The Tour operations concept requires that the DPU is able to handle trigger commands which initiate multiple actions within the instrument (macro commanding). Further, in order to optimise the analysis of discrete events such as shock crossings, a snapshot capability has been implemented by which up to 16 Mbytes of data can be stored for later downlink at higher time resolution than normal. This capability can be initiated by command or triggered by pre-defined events. Magnetic-field information is also needed by other investigations on the spacecraft. To this end magnetic-field data are made available to onboard users every second. These onboard ancillary data are raw and uncalibrated vectors, the data source being selectable by command between the two sensors. In total, the instrument consists of the two boom-mounted sensors, subchassis #1 (an assembly containing electronics for the FGM, VHM and SHM, the heater-control electronics, the power supplies and power-management system) and subchassis #2 (an assembly containing the data processing unit). Both subchassis are mounted in bay 4 of the Orbiter upper equipment module (UEM). The instrument ground-support equipment was provided by KFKI and TUB. Table I lists the main characteristics of the instrument. Power and data-rate values vary according to instrument mode. The values given in the table are for the delivered flight model where power values include power drawn by the Cassini-provided BIU. TABLE I Main Instrument Characteristics MASS V/SHM Sensor 0.71 kg FGM Sensor 0.44 kg Subchassis#1 (Power Supplies, Sensor Electronics) 5.15 kg Subchassis#2 (DPU) 2.52 kg Total 8.82 kg POWER Sleep Mode 7.50 W Vector/Vector Mode (FGM+VHM) 11.31 W Vector/Scalar Mode (FGM+SHM) 12.63 W NORMAL DOWNLINK DATA RATE FGM 32 Vectors per second VHM 2 Vectors per second SHM 1 Value per second Housekeeping 24 bits per second Total 2000 bits per second DYNAMIC RANGE, RESOLUTION FGM +/-40nT, 4.9pT +/-400nT, 48.8pT +/-10,000nT, 1.2nT +/-44,000nT, 5.4nT VHM +/-32nT, 3.9pT +/-256nT, 31.2pT SHM 256nT - 16384nT,36 pT THE FGM ------- The FGM sensor is mounted halfway along the magnetometer boom; its associated analog electronics form part of the electronics assembly on Subchassis#1. A cable of approximately 6.5-meter length runs along the boom between sensor and electronics. A high efficiency, tuned drive design of the electronics has been chosen to reduce power consumption and the effect of cable loading. The FGM is similar to the Imperial College instrument flown on Ulysses, and to many others flown on numerous missions. It is based on three single-axis ring-core fluxgate sensors mounted orthogonally on a machinable glass ceramic block. Ceramic is chosen for its low thermal expansion coefficient, minimising misalignments between sensors due to temperature changes. In each sensor, a drive coil is wound around a high-permeability ring-core which is completely enclosed in a sense winding. The drive coil is driven by a crystal-controlled 15.625 kHz square wave which is used to generate a magnetic field that drives the core into saturation twice per cycle. The three drive coils are connected in series to simplify the cabling and circuitry. The presence of an ambient magnetic-field component parallel to the axis of the sense coil causes the saturation of the core to become asymmetrical. This asymmetry induces a second harmonic of the drive frequency in the sense coil which is proportional to the magnitude of the magnetic-field component along that axis. The signal is processed through a narrow band amplifier tuned to the second harmonic of the drive frequency, which attenuates harmonics other than the second. The result is integrated, converted to a current and fed back to the sensor coil to null the ambient field. The integrated output voltage, amplified and corrected for scale factor and alignment errors, is proportional to the ambient field. The three analogue vector components are passed to the DPU for analogue to digital conversion and data processing. The noise performance of the FGM, measured on the ground at the analogue output of the electronics, is better than 5 pT/Hz at 1 Hz. The electronics can be checked in flight using an in-flight calibration (IFC) capability built into the electronics and controlled by command from the DPU. The IFC applies a fixed offset to each of the three vector outputs corresponding to a signal of approximately 10 nT. The frequency, number of on/off cycles, of the IFC is selectable by command. Changing the electronics feedback path and the output amplification allows the sensor to be operated in one of four different full scale magnetic field ranges, as listed in Table I. The largest range (+/-44,000 nT) was included mainly for ground testing in the Earth's field. Switching between ranges in normal operations is automatic, controlled by the DPU. If the magnitude of any of the FGM magnetic-field components exceeds an upper threshold for more than a specified number of samples, the DPU will switch the FGM to a higher range. Similarly, if all three component value magnitudes fall below a lower threshold for more than a specified number of samples, the DPU will switch the FGM to a lower range. All parameters are modifiable by command and autoranging can also be disabled and manual range changes commanded. A 1W heater has been provided to maintain the FGM within its operating temperature range of -30 to +50 degrees C. The specially designed, non-magnetic unit is mounted on the ceramic sensor block and has control electronics on Subchassis#1. Further thermal control is provided by an aluminised mylar-covered fibreglass case over the sensor block and by three, project-provided, radioactive heater units mounted at equal distances around the base of the sensor (these units provide a total of 3W). THE V/SHM --------- The V/SHM sensor is the flight-spare Ulysses vector-helium magnetometer sensor with an added small pair of coils nested inside the larger Helmholtz coils used in the vector mode. The sensor is mounted at the end of the 11-meter magnetometer boom. A set of cables running the length of the boom connects it to the VHM and SHM electronics on Subchassis#1. The VHM electronics box is also the Ulysses flight-spare unit with small modifications to change the sensor operating ranges and to compensate for the different boom cable lengths. A new electronics board has been added to Subchassis#1 containing the electronics to operate in the scalar mode. The operation of the magnetometer is based on field-dependent light absorption (the Zeeman effect) and optical pumping to sense the magnetic field. Helium in an absorption cell is excited by a radio-frequency (RF) discharge to maintain a population of metastable long-lived atoms. Infrared radiation (wavelength 1083 nm) from a helium lamp, also generated by RF excitation, passes through a circular polariser and the absorption cell to an infrared detector. The absorption (pumping efficiency) of the helium in the cell is dependent on the ambient magnetic-field direction. The optical pumping efficiency is proportional to cos^2(Theta) where Theta is the angle between the optical axis and the direction of the magnetic field. This directional dependence is utilised in the vector mode by applying low-frequency sweep fields rotating about the cell which allow the extraction of the three orthogonal ambient-field components. These fields are fed back using a set of triaxial Helmholtz coils mounted on the sensor housing around the cell. In the scalar mode, the directional dependence results in a 'field of view' restricted to a cone with half angle approximately 45 degrees, centred on the optical axis detector. Changing the VHM sweep fields allows the sensor to operate in different ranges. Two VHM ranges have been selected for Cassini (see Table I). As for the FGM, automatic ranging has been implemented in the DPU. The VHM electronics also have an internal autoranging capability (used for the Ulysses instrument). A single range has been implemented for the SHM. Injection of known currents into the Helmholtz-coil system provide an in-flight calibration (IFC) capability. The calibration fields apply an offset of approximately 1/8 of the full scale range to each vector component. A non-magnetic proportional heater using up to 2W is incorporated into the V/SHM sensor and is controlled from electronics built into the VHM electronics box on Subchassis#1. The operating temperature range of the sensor is -10 to +40 degrees C. In the scalar mode, a weak AC field at the Larmor frequency, which opposes the optical pumping, is applied to the cell. This field causes a reduction in the transmitted light detected by the IR detector. The Larmor frequency, which is proportional to the ambient magnetic field, is measured. In order to track the ambient field the applied field is frequency modulated so that the detector output contains a signal component harmonically related to the modulation frequency. The proportionality constant is the gyromagnetic ratio which for helium is 28.023561 Hz/nT. Detection and measurement of the Larmor frequency leads to a very accurate measurement of the ambient field magnitude. The result is passed as a 20-bit scalar word from the SHM electronics to the DPU. A more detailed description of the V/SHM may be found in Kellock et al. (1996). Smith et al. (2001) provides a detailed description of the SHM operation and observations from the Earth Swingby in August 1999. INSTRUMENT ELECTRONICS ====================== The instrument electronics are all mounted in the Upper Equipment Module in bay 4 and are split between two subchassis assemblies. Subchassis#1 contains the sensor electronics, the power supplies and power management system, as well as the heater control and instrument housekeeping electronics; its mass is given in Table I. On the underside of the subchassis are the power-management board, the FGM electronics board and the SHM electronics board. The power-management board contains a total of 14 non-latching power switches and cross-strapping circuitry for the two redundant secondary power supplies. The top side of the subchassis contains the VHM electronics box, two small boards to the left of the VHM box with latching relays to switch between VHM and SHM operation, two redundant secondary power supplies (PSU1 and PSU2), a dedicated BIU power supply (PSU0), and the FGM heater-control electronics and housekeeping circuitry. Proportional control electronics for the VHM heater are located within the VHM box. The power supplies and switches are Imperial College designs used on previous missions and feature built-in overcurrent trips. The basic power distribution scheme is described in Kellock et al. (1996). Power switches for the secondary voltage lines are controlled by the active processing unit and power switches for the power supplies and processing units themselves are controlled by discrete commanding from the spacecraft via the BIU and the Common Core (CC). Subchassis#2 contains the DPU, consisting of two redundant processor systems plus a small CC and the BIU. When power is first supplied from the spacecraft, only the BIU and the CC become active, powered from PSU0. The BIU allows data transfer to and from the spacecraft, the CC processes commands and data for power up of the secondary power supplies (PSU1 or PSU2) and the processors (PUA or PUB). Each processor system is based on an 80C86 processor with 4-MHz clock, 32-kByte PROM, 128-kByte Hi-Rel RAM and 16-MByte state-of-the-art commercial DRAM. The systems normally operate singly but can be operated in parallel. A high-accuracy 16-bit analogue to digital converter (ADC) is integrated into each processor system for sensor data collection. Two ADC clock speeds are available, 1 MHz and 2 MHz, the former being the default speed. Tantalum shielding has been used for the ADCs, DRAMs and Operational Amplifiers to reduce their susceptibility to radiation. The DPU boards are folded around the subchassis. The electronic components face the subchassis, because the 2.4 mm thick 16-layer boards provide additional radiation shielding. The Sensor Interface Board is on the right hand side. The flexible connection board goes through subchassis cutouts to the Processor Board on the other side. The JPL-provided BIU, plus its associated cabling, is located on the left side. To satisfy the demands of a deep-space mission with limited communications, the DPU has been designed with a large measure of autonomy and sophisticated data-handling functions. These functions include the following: telecommand handling, sensor autoranging and IFC, sensor data collection, sensor data processing, snapshot data handling, telemetry generation, error correction, fault detection and recovery, and onboard ancillary data generation. Some of the functions of the DPU are described below. The DPU is designed to handle both the packet telecommand standard adopted by Cassini, used for normal commanding, and discrete telecommands used when the processor is not active. A variety of command functions are supported and are discussed later. The DPU must be able to accept telecommands at all times, in all operational modes. Commands may be for immediate execution, or can contain relative or absolute timetags for delayed execution (relative timetags cause execution at a fixed time with respect to reception of the command, absolute timetags cause execution at specific spacecraft times). As noted earlier, a macro commanding capability has been implemented whereby sequences of instrument commands can be stored in the DPU and the sequence started by executing a macro command. Hamming single-bit error correction and double-bit error detection is provided for all memory devices except those in the BIU. Single Event Upsets (SEUs) can change the content of memory cells, cyclic access to every memory cell corrects single-bit and reduces the risk of double-bit errors. Memory scrubbing is initiated every 64 seconds in the Hi-Rel RAM and the 16-MByte, multi-snapshot, DRAM. It takes about 1 hour to scrub the complete RAM. Additional memory checks can be initiated by command: occurences of single and double-bit errors are monitored. The PROMS are checked separately and contain a pre-defined error pattern for detection. If a permanent RAM problem arises, the DPU can be commanded to run its software directly from PROM. In-flight tests have also been implemented for the ADC to check the noise and conversion and settling times on each analogue channel. Both processor systems contain four separate dual-level latch-up detectors, one for the processor, one for the multi-snapshot memory, and one each for the ADC +/-12V supply voltages. Detectors of similar design have been flown on the GEOTAIL, WIND and SOHO spacecraft. If a latch up is detected in the processor, it will be immediately switched off and on again and the instrument will be automatically reconfigured into its previous mode. Latch ups in either the memory or the ADC will cause it to be immediately switched off and back on again. There is also a hardware watchdog function in the DPU which will detect problems in the DPU program flow and which initiates a hardware reset followed by an automatic instrument reconfiguration." 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