The above objectives are accomplished on MMS with six sensors integrated into the FIELDS suite, as diagrammed in Fig. 6 (institutional contributions indicated), with the control and data flow managed by the Central Electronics Box (CEB). The CEB also contains controllers for the Analog Flux Gate (AFG), Digital Flux Gate (DFG), and the Electron Drift Instrument (EDI) as well as the Digital Signal Processor (DSP) for the high frequency instruments of Search Coil Magnetometer (SCM), Spin-plane Double-Probes (SDP), and the Axial Double-Probes (ADP). Within the CEB, the Central Data Processing Unit (CDPU) coordinates all the functioning of the FIELDS suite, including power switching through the Low Voltage Power Supply (LVPS). The CDPU, LVPS, and DSP are all fully cold redundant. The CEB determines the synchronous timing regimen for the entire suite, all command processing, and can deliver up to 4 Mbps of data to the MMS spacecraft Central Instrument Data Processor (CIDP). Each of these sensors is described more completely in following companion papers, but this communication attempts to describe how they are all integrated together into one highly calibrated and cross-calibrated instrument suite to measure electromagnetic fields more precisely and comprehensively than has ever been done in space before.
The sensors are arranged on each of the four MMS spacecraft according to Fig. 7.
Analog Flux-Gate and Digital Flux-Gate
The DC magnetometer measurements are provided by two flux-gate three-axis sensors, each at the end of 5-meter deployable booms, and associated electronics within the CEB. Provided by UCLA, each sensor consists of two magnetic cores, their housings and drive wire windings, 6 sense wire windings, 6 feedback wire windings, and two printed circuit boards mounted on an armature, which provides a framework for the components (see Russell et al. 2014, this issue, for a more complete discussion). The Analog Flux-Gate (AFG) has a somewhat different controller, provided also by UCLA, than the Digital Flux-Gate, which is provided by IWF, but both magnetometers evolved to basically the same digital feedback design, although the DFG is implemented on a specially designed ASIC. Each controller produces a fixed 128 samples/s data stream to the Central Data Processing Unit (CDPU) within the CEB, which implements digital filters to reduce the sample rate if necessary due to telemetry restrictions. Two output ranges are available for both AFG and DFG, of ∼500 nT magnitude for low range to ∼8200/10500 nT (AFG/DFG), for high range. The ranges are commanded by the CDPU using an algorithm with hysteresis based on the data from the magnetometer controllers. Because of the key role of the magnetic field for reconnection studies, AFG and DFG provide fully redundant 3D data streams that are used both on-board the spacecraft by other instruments and also for ground processing. Extensive calibration and cross-calibration of the magnetometers was undertaken at the Technical University Braunschweig. An extensive magnetic cleanliness program for the MMS satellite system was supervised and validated by the magnetometer team. Also, timing calibrations were performed to determine the phase and gain curves versus frequency, as shown in Figs. 16 and 17. These calibration data show that both the AFG and DFG have the capability to measure the DC and low frequency component of the vector magnetic field over the full range of each magnetometer with a timing accuracy of better than 0.1 ms.
In order to reach the science objectives, the AFG and DFG magnetometers are also calibrated on orbit. The calibration procedures include comparison of the AFG and DFG gains and offsets across range changes, Earth-field comparisons, cross-calibration with EDI, and inter-spacecraft calibration. More information can be found in Russell et al. (2014, this issue).
Search Coil Magnetometer
The SCM provides the three components of the magnetic fluctuations in the 1 Hz–6 kHz nominal frequency range, which is the imposed requirement. This range includes the hybrid wave and kinetic Alfvén wave frequency range as well as whistler mode waves (up to their cut off frequency equal to the electron gyrofrequency) and solitary waves. SCM consists of a triaxial induction search coil wound around a ferromagnetic core, mounted 4 meters out on the AFG boom, with the transfer function as measured for Flight Model 2 (FM2, Fig. 8). The noise equivalent magnetic induction (NEMI or sensitivity) of the search-coil antenna is less than or equal to 2 pT/sqrt (Hz) at 10 Hz, 0.3 pT/sqrt (Hz) at 100 Hz and 0.05 pT/sqrt (Hz) at 1 kHz. An in-flight calibration signal provided by DSP allows the verification of the SCM transfer function once per orbit.
The analog waveforms from a pre-amplifier are digitized and processed inside the DSP with a resolution at 1 kHz of 0.15 pT, and are telemetered as SCM 1, 2, and 3 as indicated in Table 1.
Spin-Plane Double Probe
The Spin-plane double probe instrument (SDP) measures the electric field in the spin plane by sensing the potential difference between four current-biased spherical titanium-nitride electrodes, each of diameter 8.0 cm at the end of 60-meter long wire booms. Together with the axial double probe instrument (ADP, described below), SDP measures the 3-D electric field with an accuracy of 0.5 mV/m over the frequency range from DC to 100 kHz. By means of a thin titanium wire, the spheres are held 1.75 m beyond the ends of a preamplifier which provides a low impedance, unity-gain signal of the sphere potential to electronics located at the base of each boom on the spacecraft, as seen in Fig. 7. The preamplifier outer casing is divided into an inner and outer “guard”, which can each be biased at ±10 V with respect to the sphere. Current biasing to the sphere is routed also through the preamplifier. The unity gain signal is used to drive the outer conductors around the primary signal wire, thereby reducing the effective capacitance of the long wires in the boom, up to a frequency of about 300 Hz and voltages from −80 to +100 volts with respect to the spacecraft ground. Electric field components are produced by dividing the potential difference between opposite pairs of spheres (e.g., derived from E12=0.0415∗(V1−V2)) by an effective antenna length. The factor of 0.0415 is the approximate electronic gain. The effective antenna length can be estimated by considering the field configuration of spheres just beyond the ends of long grounded wire booms immersed in a vacuum (Fahleson 1967). However, this length can vary slightly with plasma conditions and is determined in flight by comparison to known fields, such as those for co-rotation in the Earth’s inner plasmasphere or those measured by EDI, as described below. There are also AC coupled versions of Eij (E12, E34, and E56) with higher gain, called Eij_AC. The actual voltages of the spheres, V1–V6, and their AC coupled versions (V1_AC and V2_AC) are also telemetered. Vx (V1 to V6) is also shared with other instruments on board, such as ASPOC, that request this data.
The voltage of a sphere in a plasma floats to a value such that the total net current to the sphere is zero. Thus, error currents, such as asymmetrical photo- or secondary emission from either the sphere itself or the surrounding electrodes and spacecraft drives error voltages in the values of Eij. Every effort has been made on MMS to reduce this effect. The spacecraft spin axis is tilted with respect to the sun so that neither the preamp nor the spacecraft will shadow the sphere. The spacecraft itself was subjected to a rigorous electrostatic cleanliness and symmetry program. The sphere coating is manufactured to be as uniform as possible. UV reflectance tests were performed to ensure optimum matches of coatings for sphere pairs. Biasing the sphere at the minimum of dV/dI (lowest effective resistance) is very effective in reducing voltage offsets driven by error current effects. The photoelectron cloud and variation of the spacecraft potential are reduced by active spacecraft potential control of the ASPOC instrument (Torkar et al. 2014, this issue). In-flight comparisons of the resulting field with EDI also serve to identify and eliminate the remaining errors, as described below.
Although ASPOC reduces both the magnitude and variation of the spacecraft potential, the gun energy spectrum is not a delta function, and thus small variations (∼0.1 V) remain which are a function of the ambient electron flux to the spacecraft and the spheres. Analysis of these variations still allows an estimate of the local electron density, with assumptions about temperature, that are very useful in determining spatial variations of plasma conditions, predominantly the ambient density.
Axial Double Probe
The Axial Double Probe (ADP) instrument measures the electric field, DC to ∼100 kHz, along the spin axis of the MMS spacecraft with an accuracy of better than 1 mV/m. It uses the double probe technique by sensing the local plasma potential at two sensors separated by ∼29.2 m effective antenna length. The axial direction, which completes the vector electric field when combined with the SDP, has been the most challenging component of the DC electric field measurement. The physical antenna lengths are limited by mechanical difficulties, which include deployment of stiff booms while preserving spacecraft stability. The ADP baseline is nearly twice that of the Polar mission (∼16 m) creating the longest axial baseline ever attempted for a DC electric field measurement.
The ADP on each of the spacecraft consists of two identical, 12.67 m graphite coilable booms (made by ATK space systems). A guard ring, 30.9 cm in diameter and 2.6 cm high, encircles the mounting plate at the end of the coilable boom. A second, smaller boom is mounted on the out-board end of each coilable boom. These 2.25 m booms (Fig. 9) are folded onto the top and bottom of the spacecraft for launch. The smaller outer boom is deployed soon after launch, followed several weeks later (after SDP is fully extended) by the deployment of the 12.67 m coiled boom. When deployed, the outer booms are comprised of (going away from the spacecraft) a 90° base hinge, a 0.78 m long by 1 cm diameter tube, a 180° hinge, a 0.43 m long by 0.95 cm diameter tube, a 5.6 cm long by 2.1 cm diameter preamplifier, and a 1 m long by 0.64 cm diameter sensor.
As done with the SDP, a significant effort has been made to assure a constant photo- and secondary electron current to the ADP sensors. All elements of the 2.25 m booms are cylindrical after deployment, including the hinges. The stubs, preamplifiers and sensors are coated with graphite-epoxy (DAG 213) to assure consistent surface properties as the spacecraft rotates. Equally important is the symmetry between the top and bottom sensors. The ADP booms are mounted to be symmetric about the spacecraft electrostatic plane, which is dominated by the spin-plane wire booms. As such, the lower boom (underside of the spacecraft in Fig. 7) is recessed into the spacecraft by ∼0.1 m. The MMS spacecraft are to be oriented so that the ADP booms are within 5° of normal to the sun. A guard ring is placed at the base of the outer boom to shadow the top surface of the mounting plate. The guard ring is designed so that in the nominal attitude, the mounting plate is not exposed to the sun, assuring a nearly identical photo- and secondary electron environment near the two sensors.
As in the SDP, the ADP sensors are fed a bias current, often a significant fraction of the expected photoelectron current, to minimize the resistance between the plasma and the sensor. To further control the photoelectron environment, we control the surface potentials of (a) the guard ring, (b) the under side of the mounting plate, and (c) the hinges, stubs, and preamplifier housing. Each surface section can be set to ±10 V with respect to the sensor’s DC potential.
The ADP on MMS is expected to measure the DC electric field with an accuracy of ∼1 mV/m, a resolution of 0.026 mV/m, and a range of ∼ ±1 V/m in most of the plasma environments that MMS will encounter. Constant offsets between the booms will be removed by two methods: namely, minimizing E⋅B as measured by the double probe over long (>20 s) periods and comparison with EDI electric field measurements. The spectral power density has a dynamic range from 4×10−16 (V/m)2/Hz to 10−3 (V/m)2/Hz at 10 kHz.
Electron Drift Instrument
EDI determines the electric and magnetic fields quite differently from all the sensors above. It is basically a geometric measurement for the electric field and a timing measurement for the magnetic field. As seen in Fig. 10, two electron beams are emitted in nearly opposite directions from two Gun-Detector Units (GDU) on opposite sides of each spacecraft. Each beam drifts in the E×B direction and, if properly directed, returns to the spacecraft after nearly one or more gyroperiods. If the drift-step (d=drift velocity, v
, times gyroperiod) is of the order of the baseline separation of the two GDU’s, then the electric field is determined by triangulation as seen in Fig. 11. In this figure, the first GDU (G1) emits a beam in the direction V
1 that is detected by the opposite detector, D, and vice versa for G2. The drift step is the displacement of the intersection of the two beams (S) from the position of the detector. The actual geometry is slightly more complicated in that the detector for G1 is located at G2 and vice versa. The beams are pseudo-noise encoded so that the emitted electrons can be unambiguously detected in the presence of ambient electrons, and the time of flight of the beam can be determined. The difference of the time of flight of the two beams gives the magnitude of the drift step (which can be used in the “time-of-flight” mode when the drift step is large compared to the baseline) and the average of the two times gives the gyroperiod. From the gyroperiod, the magnitude of the magnetic field is determined, and from the directions of the successful beams, the direction of the magnetic field can be computed. The advantage of EDI over conventional electric and magnetic sensors is that the effects of the fields far from the spacecraft dominate the resulting vectors: electrostatic and noise magnetic fluctuations of the spacecraft have little effect when the gyroradius is of order kilometers, as is the case for MMS. But EDI also has a very slow time cadence for a full vector determination (of order 10 samples per sec) compared to AFG, DFG, SCM, SDP, and ADP. By combining multiple techniques, as described below, improved accuracy can be obtained with high time resolution. The requirements on FIELDS as a whole for an electric field accuracy of 0.5 mV/m and a magnetic field accuracy of 0.1 nT can thus be met. In order to detect the weak (∼100 nA) electron beams emitted by the guns, the detector of EDI has very large geometric factor (order 0.01 cm2 str) so that very fast sampling of ambient electrons is possible at a fixed energy and for a few directions. Thus, 0.5 or 1 keV electron fluxes can be determined in the “ambient” mode at 1024 samples per sec and very thin electron layers can be detected.
Central Electronics Box
The CEB directs all the activities of the FIELDS sensors and formats both housekeeping and science data for transmission to the CIDP and eventually to the downlink for ground processing. As diagrammed in Fig. 12, the operating system in CDPU (RTEMS) structures the flight software activities, and allows for command handling (CMD) by non-maskable interrupts (NMI) and memory error checking (EDAC). All magnetometer data comes from AFG and DFG as a continuous stream of 128 vectors per second. The CEB performs digital filtering on this data down to the commanded rate and also internal coordinate system rotation of the field components for use by FIELDS and other on-board instruments on MMS. The relevant coordinate systems are seen below (Figs. 13 and 14). The CEB directs the traffic of EDI and DSP data, and uses some of those data for internal calculations of the Trigger Data Numbers that are used in the ground algorithms to determined selection of BURST data. It also houses the Low Voltage Power Supply (LVPS) that controls all power distribution in FIELDS as well as the floating power supplies that drive both the SDP and ADP sensors.
Each of the FIELDS sensors can produce data over a large range of sampling rates, as seen in Table 1. In addition, all the DSP input channels (SCM, Vx, Eij, and Eij_AC) are used to produce spectral products over a time series of 1024 samples which are then averaged and sent down as a frequency spectrum (magnitude only). Ancillary data include in-flight calibration data for SDP, ADP, and SCM, and timing information in the housekeeping data. There are also specialized data products of the DSP that are short samples of all inputs at very high rates, but only small duty cycles. These include the High Speed Burst channels (HSB-E and HSB-B) as well as the output of a Solitary Wave Detector algorithm.
Depending on the science questions to be addressed, a very large number of possible telemetry modes can be constructed by ground command. These are limited only by the total bit rate allocation for FIELDS seen in Table 3. As described in a companion paper on BURST mode (Fuselier et al. 2014, this issue), the CEB executes two basic modes, Slow and Fast Survey, and produces continuous data products that are always telemetered to the ground. In addition, very high data rate “BURST” products are produced only in Fast Survey mode and sent to the CIDP for storage. Only interesting intervals of this data are selected for ground transmission because there is not enough overall telemetry to accommodate BURST mode data over the full orbit. At launch, there are two modes for the BURST data: one for dayside reconnection studies in Phase 1 and one from the magnetotail studies in Phase 2. These default sampling modes, which can be changed or replaced in flight, are listed in Table 2 (nominal duty cycles for HSB are indicated as percentages).
The overall resource utilization of FIELDS is given below. The masses do not include those for spacecraft harness or magnetometer booms.