Minimize MMS Observatory

MMS (Magnetospheric MultiScale) Constellation

MMS is a NASA solar-terrestrial probe constellation comprising four identically instrumented spacecraft that will use Earth's magnetosphere as a laboratory to study the microphysics of three fundamental plasma processes: magnetic reconnection, energetic particle acceleration, and turbulence. These processes occur in all astrophysical plasma systems but can be studied in situ only in our solar system and most efficiently only in Earth's magnetosphere, where they control the dynamics of the geospace environment and play an important role in the processes known as ”space weather.” 1) 2) 3) 4)

The fundamental plasma physics process of reconnection in the Earth's magnetosphere are to be monitored on temporal scales of milliseconds to seconds and on spacial scales of 10s to 100s of km. Three science objectives have been identified for the MMS mission. In priority order, these objectives are to:

• Determine the role played by electron inertial effects and turbulent dissipation in driving reconnection in the electron diffusion region

• Determine the rate of magnetic reconnection and the parameters that control it

• Determine the role played by ion inertial effects in the physics of reconnection.

The overall science objective is to investigate the physics of magnetic reconnection: 5)

- In the plasma universe magnetic fields connect and disconnect explosively transferring energy into the electrons and ions that make up the plasma.

- MMS will reveal for the first time the small-scale 3D structure and dynamics in the heart of reconnection regions occurring naturally in Earth’s geospace plasma environment.

- Only by measuring the behavior of the particles and fields in three-dimensions in the diffusion region can we learn fully how reconnection proceeds.

- The four MMS spacecraft will be placed “surgically” in a tetrahedral formation into the diffusion region.

Understanding the physics of reconnection:

- The most important goal of MMS is to conduct a definitive experiment to determine what causes magnetic field lines to reconnect in a collisionless plasma.

- Magnetic Reconnection is evident throughout the plasma universe

- Yet it occurs in the near vicinity of Earth. It can be studied internally by man-made probes using the Earth’s magnetosphere as a laboratory.

- The 4 spacecraft MMS flying in an adjustable pyramid-like formation through the core of these dynamic and impulsive energy conversion regions allows the 3D structure to be investigated.

- Through detailed measurements of the local plasmas and fields scientists will understand the physics that governs one of the most important drivers of space weather.

- ..... and learn how charged particles are energized on the sun and in astrophysical bodies throughout the universe.


Figure 1: Schematic view of the reconnection region (image credit: NASA)

The MMS mission employs four identically instrumented spinning spacecraft orbiting the Earth in a tetrahedral configuration to conduct definitive investigations of magnetic reconnection in key boundary regions of the Earth’s magnetosphere. The process of magnetic reconnection, which controls the flow of energy, mass, and momentum within and across plasma boundaries, occurs throughout the universe and is fundamental to our understanding of astrophysical and solar system plasmas. It is only in the Earth’s magnetosphere, however, that it is readily accessible for sustained study through the in situ measurement of plasma properties and of the electric and magnetic fields that govern the behavior of the plasmas.

Through high-resolution measurements made by each spacecraft, whose separations can be varied from tens of km to a few hundreds of kilometers, MMS will probe the crucial microscopic physics involved in these fundamental processes; determine the 3D geometry of the plasma, field, and current structures associated with them; and relate their micro-scale dimension to phenomena occurring on the mesoscale. By acquiring data simultaneously at multiple points in space, MMS will be able to differentiate between spatial variations and temporal evolution, thus removing the space-time ambiguity that has limited single-spacecraft studies of magnetospheric plasma processes.

The MMS mission is managed by the Heliophysics Division for NASA’s Science Mission Directorate at NASA/HQ. NASA/GSFC will build the four spacecraft and the intersatellite ranging and communication system. Southwest Research Institute (SwRI) of San Antonio, TX leads the science investigation and development of the instrument suite together with numerous partners, including the University of New Hampshire, NASA/GSFC, Johns Hopkins University /Applied Physics Laboratory (JHU/APL), LASP (Laboratory for Atmospheric and Space Physics) at the University of Colorado, Boulder, and international partners in Austria (IWF), Sweden (KTH and IRF-U), France (CETP) and Japan.

The technology requirements of the MMS mission call for:

• On-board propulsion

• Intersatellite communication

• Autonomous operation.


Figure 2: Artist's rendition of the MMS mission in Earth's magnetosphere (image credit: NASA, SwRI)

Project status:

• In May 2005, NASA selected the SwRI instrument suite team to work with GSFC in the MMS project for the mission formulation. In May 2009, the project passed the preliminary design review. In August 2010, the project completed its mission critical design review (CDR). The C/D Phase of the project started in 2011. 6) 7)

• On August 31, 2012, NASA's MMS mission passed the SIR (System Integration Review) which deems a mission ready to integrate instruments onto the spacecraft.8)

• In August 2013, the MMS mission is in Phase-D, Integration and Test.


The four spin-stabilized spacecraft are being developed and integrated at GSFC. Each satellite has an octagonal shape that is approximately 3.5 m wide and 1.2 m high. The satellites spin at 3 rpm during science operations. There are 8 deployable booms per satellite: four 60 m wire booms in the spin plane for electric field sensors, two 12.5 m booms in the axial plane for electric field sensors, and two 5 m booms in the spin plane for the magnetometers. 9) 10) 11)

The aluminum structure has a modular design to simplify I&T (Integration & Testing) consisting of a propulsion assembly, separation system/thrust tube, instrument deck, and spacecraft deck.

The ACS (Attitude Control Subsystem) keeps the spacecraft to within ±0.5º of the desired orientation during science operations and implements on-board closed loop maneuver control. The ACS sensor complement consists of a star tracker (ST) with four separate heads and two redundant DSS (Digital Sun Sensors), one being a cold back-up unit. The four ST heads output four independent quaternions at 4 Hz (that is, 16 independent attitude measurements/s). Once per spin period (~20 sec), the DSS will output a pulse indicating sun-crossing through the sensor FOV slit and a measurement of the sun elevation from the body X-Y-plane (the +Z-axis is the nominal spin axis). 12)

Star sensors and sun sensors provide attitude data, and accelerometers provide acceleration and ΔV data. Thrusters are used as actuators. A GPS receiver on board each spacecraft provides absolute position information. In addition, each spacecraft employs an IRAS (Inter-spacecraft Ranging and Alarm System) to determine its location relative to the other three spacecraft.

The propulsion subsystem is a mono-propellant blowdown system with 12 thrusters sized to achieve both small formation maintenance maneuvers and large apogee raise maneuvers. Approximately 360 kg of propellant will be contained in four titanium tanks per spacecraft.


Figure 3: Side view of a MMS spacecraft (image credit: NASA/GSFC)

EPS (Electrical Power System): The EPS is a 'direct energy transfer system' employing a battery dominated bus. Power to the spacecraft is supplied from 8 identical body-mounted solar array panels that are electrostatically and magnetically clean. The battery is sized to provide power during the 4 hour eclipses. - The thermal design is passive using thermostatically controlled heaters.

RF communications: Ground communications occur over a single S-band frequency for uplink to all four spacecraft and a single S-band frequency for downlink from all four spacecraft. Real-time coverage of all critical commands, including post launch separation and all maneuvers, will be accomplished through TDRSS (Tracking and Data Relay Satellite System) of NASA.

The spacecraft bus avionics performs command and telemetry processing, timing distribution, solar array regulation, battery charge management, and thruster control. Orbit determination is performed on-board using weak signal GPS processing. Each satellite including instruments, fuel, and margin has a mass of ~ 1,250 kg. The power budget at end of life with instruments and margin is approximately 318 W. The mission design life is 2 years.

Each MMS spacecraft is being developed using the standard GSFC protoflight testing approach. The first unit is being tested to qualification levels and the remaining units to acceptance levels. Design heritage from previous GSFC in-house development efforts are being used where possible, although each unit having heritage will go through the full design and test process. Commercially available components are being purchased competitively. Planned procurements include accelerometers, star cameras, sun sensors, batteries, solar arrays, transponders, thrusters, propellant tanks, and separation systems.

Spacecraft to instrument suite integration and constellation level testing will occur at GSFC. Each instrument suite will be integrated with the spacecraft bus to form the MMS observatory. Performance and environmental testing will be performed on each observatory. Additional testing will be performed in the stacked launch configuration.


Figure 4: Schematic view of the element arrangement within a MMS spacecraft (image credit: NASA)


Figure 5: Illustration of the MMS spacecraft (image credit: NASA)


Figure 6: Stacked launch configuration of the MMS spacecraft (image credit: NASA)


Figure 7: Photo of two MMS observatories (June 26, 2013) stacked up for shock testing to make sure they can withstand the launch environment (image credit: NASA)


Launch: A common launch of the four satellites is scheduled for the fall of 2014 on an Atlas-5-421 vehicle. The launch site is the Cape Canaveral Air Force Station, FL, USA. 13)


Orbits of the constellation:

MMS is designed to fly four identical spin-stabilized spacecraft in a tetrahedral formation in a set of highly elliptical orbits.

Because reconnection manifests itself in the Earth's magnetosphere at two locations with differing scale sizes and magnetic field orientations, a two-phase orbit strategy has been developed to test the universality of the mechanisms at work and to better understand how it controls planetary space weather. Phase 1 will probe reconnection sites at the mid-latitude dayside magnetopause, while Phase 2 focuses on reconnection sites that occur within the nightside magnetic neutral sheet. 14) 15)


Figure 8: Schematic view of Earth's magnetic field with the two regions of particular interest shown in red (image credit: SwRI)

The MMS mission aims to improve on the Cluster mission is several ways. Like Cluster, MMS will deploy a formation of four spacecraft moving in a close tetrahedral formation about a highly elliptical orbit with each of the spacecraft carrying a suite of instruments for in situ measurements of electric and magnetic fields and charged particle composition. The orbit selection and relative spacing are quite different with the mission intending to fly two distinct science phases each with multiple formation scale sizes.

Phase 1: day side of magnetic field 1.2 RE x 12 RE. MMS will probe reconnection sites at the mid-latitude dayside magnetopause (red region on the left in Figure 8). Here the interplanetary magnetic field (IMF) merges with the geomagnetic field, transferring mass, momentum, and energy to the magnetopause. The solar wind flow transports the merged IMF/geomagnetic field lines toward the nightside, causing a build up of magnetic flux in the magnetotail.

In Phase 1, the formation flies with a relative spacing ranging from 10-160 km in a 1.2 x 12 RE orbit. Primary science is taken when the tetrahedron is at distances greater than 9 RE from the Earth and is within 30º of the Earth-Sun line, on the sunward side.

Phase 2: night side of magnetic field 1.2 RE x 25 RE. The MSS constellation will investigate reconnection sites in the nightside magnetotail (red region on right side of Figure 8), where reconnection releases the magnetic energy stored in the tail in explosive events known as magnetospheric substorms and allows the magnetic flux stripped away from the dayside magnetopause by the solar wind/magnetosphere interaction to return to the dayside.

In Phase 2, the relative spacing is 30-400 km of the spacecraft occurs in a 1.2 x 25 RE orbit. Primary science is taken when the tetrahedron is at distances greater than 15 RE from the Earth and is within 30-40º of the Earth-Sun line on the night-ward side.

Orbits: HEO (Highly Elliptical Orbit) of 1.2 RE (perigee) x 12 RE (apogee), inclination = 28º. Note 1 RE = 6371 km. As the orbit evolves during Phase 1, the spacecraft will sample reconnection sites at different locations on the dayside magnetopause.

A transfer phase connects the two science phases by the execution of a series of maneuvers to raise each spacecraft's apogee while keeping the formation from drifting too far apart. The variable distances from Earth and inter-spacecraft spacing allows the MMS formation to act as a science instrument (Figure 9), with the inter-spacecraft distances being chosen to match the scale sizes of the physics occurring at different locations in the Earth's magnetosphere (Ref. 15).


Figure 9: Schematic of the MMS formation as a science instrument concept (image credit: NASA)

Execution of the basic mission design requires that both deterministic and random maneuvers be performed in operations. The deterministic maneuvers are used to change the overall orbital characteristics of the formation and the random maneuvers arise from the need to maintain the formation against the relative drift that builds up over time. The number, size, and direction of these maintenance maneuvers are unknown a priori and are very dependent on the realization of a number of error sources (natural and man-made) in the system.

Earth's J2 term in the geopotential causes differential evolution of the orbit states of each spacecraft and is the largest natural perturbation. Lunar gravitational affects also play a significant role in Phase 2, because of the large apogee, causing both relative drift and, on occasion, a secular lowering of the perigee altitude below acceptable limits. Taken as a whole, these natural perturbations are well understood and predictable, depending deterministically on the initial orbital state and its subsequent evolution. The formation lifetime, defined as the time between maintenance maneuvers, in the presence of these natural perturbations, is on the order of 40-70 days.

In contrast, the man-made perturbations due to knowledge and execution errors in the maneuver process generically dominate the natural perturbations. Taken together, this combination of knowledge and execution error causes a post-maneuver evolution that is different from the one desired. The situation is further complicated by the fact that each spacecraft experiences its own realizations of these errors with no correlation between them. The man-made error sources associated with maneuver operations generally lower the formation lifetime to a range of 5-20 days.

The variable nature in the number of and type of maneuvers coupled with the „cooperative effect? of the errors makes pre-launch planning and operational support of the MMS formation flying problem very difficult. Questions associated with the stability of the design, such as how much fuel to budget, how maneuvers will be performed, and when to schedule ground assets to support the maneuvers, cannot be answered with simple analytical models. It was to address this need that the MMS flight dynamics group developed the End-to-End (ETE) code. The ETE code is designed to perform Monte Carlo simulations of the entire MMS operations phase from launch until the end of Phase 2 science. It combines high-fidelity orbit propagation, realistic models of the propulsion system, and models for navigation knowledge and maneuver execution errors into an event-driven framework that allows it to produce different maneuver scenarios in response to the different errors it encounters. Statistical reduction of the resulting data is then used to allocate resources for the actual launch campaign (Ref. 15).


Figure 10: Illustration of orbits in Phase 1 and Phase 2 (image credit: NASA, SwRI)


Figure 11: An artist's concept of the four MMS spacecraft flying in formation through the space around Earth (image credit: NASA)



Each spacecraft will fly the GSFC-developed IRAS (Inter-spacecraft Ranging and Alarm System), which consists of the Navigator GPS receiver integrated with a crosslink transceiver and a high quality frequency reference [i.e. an ultra-stable oscillator (USO)]. The tracking loops in the Navigator receiver are tuned to acquire low strength GPS signals to increase the number of GPS Space Vehicles (SVs) that can be acquired at high altitudes. This receiver has been demonstrated to reduce the acquisition threshold below 25 dB Hz as compared with a threshold of 35 dBHz that is typical for GPS receivers designed for LEO (Low Earth Orbiting) satellites. Each IRAS will also acquire and transmit one-way crosslink range measurements from the other formation members at intervals of 4 minutes. The GPS pseudorange (PR) and crosslink range measurements and associated state vectors for each of the formation members will be provided as data via the intersatellite link. 16) 17) 18) 19)

To perform on-board orbit determination, the IRAS hosts the GEONS (GPS Enhanced Onboard Navigation System) flight software. GEONS is a flight software package developed by NASA to provide onboard orbit determination for a wide range of orbit types. GEONS is capable of using GPS measurements and intersatellite crosslink measurements to simultaneously estimate absolute and relative orbital states. GEONS employs an EKF (Extended Kalman Filter) augmented with physically representative models for gravity, atmospheric drag, solar radiation pressure, clock bias and drift to provide accurate state estimation and a realistic state error covariance.


Figure 12: Space/ground navigation configuration of MMS (image credit: NASA)


MMS operation concepts:

Each spacecraft deploys 8 booms: 4 SDP (Spin-plane Double Probe) instruments on wire booms in the spin-plane for measuring the electric fields; 2 magnetometer instruments, also on booms in the spin-plane, for measuring the magnetic field; and 2 ADP (Axial Double Probe) instruments on rigid booms parallel to the spin-axis for measuring electric fields.

The propulsion system for each spacecraft consists of 8 radial 18 N thrusters oriented parallel to the spin plane, with sets of 4 on opposite sides of the spacecraft, and 4 axial 4.5 N thrusters. Each spacecraft is also equipped with a suite of guidance, navigation, and control sensors including a digital sun sensor, a star camera, an accelerometer, and a Navigator GPS receiver with GEONS navigation software built-in. Onboard controllers process data from this sensor suite to actuate the thrusters to the desired accuracy in the orbit and attitude maneuvers while ensuring the stability and safety of the deployed booms. Figure 13 shows a diagram of the spacecraft showing the deployed booms and a schematic of the propulsion system.


Figure 13: MMS spacecraft with deployed booms and with a propulsion system (image credit: NASA)

Legend to Figure 13: (a) MMS spacecraft showing the 8 booms for the electric and magnetic fields deployed and (b) MMS spacecraft with propulsion system with the location of the 18 N radial thrusters shown in red and the location of the 4.5 N axial thruster shown in yellow (with symmetric placement of the other 6 thrusters on the opposite faces).

From the flight dynamics perspective, the MMS operations concept is best understood by dividing the driving requirements into two categories. The first category, called the baseline, holds all of the requirements related to how the formation as a whole moves with respect to the magnetosphere. The second category, called formation flying, holds all of the requirements related to the relative orbital evolution of the formation.




SMART (Solving Magnetospheric Acceleration, Reconnection, and Turbulence) is the name of the MMS science investigation program headed by James L. Burch of SWRI as PI (Principal Investigator) in cooperation with researchers from other institutions. 20) 21) 22) 23) 24) 25)

The SMART payload comprises three instrument groups: Hot Plasma, Energetic Particles, and Fields. In addition, the payload includes two ASPOC (Active Spacecraft Potential Control Devices) and a CIDP (Central Instrument Data Processor). The ASPOCs neutralize the electrical potential of the spacecraft, allowing measurement of low-energy ions and electrons by the plasma instruments and eliminating spurious electric fields that can contaminate double-probe measurements. The CIDP provides the interface between the instruments and the spacecraft C&DH subsystem. The ASPOCs are being developed at the IWF (Institut fuer Weltraumforschung) of the Austrian Academy of Sciences; the CIDP is being developed at Southwest Research Institute (SwRI).

Identical in situ instruments on each satellite measure:

• Hot plasma composition.

• Energetic particles

• Electric and magnetic fields


Figure 14: SMART instrument suite organization (image credit: NASA)

Legend to Figure 14: The organization was developed to give team PMs (Project Management) cost and schedule ownership.


Figure 15: Schematic view of an MSS spacecraft with the payload (image credit: SwRI) 26) 27)


Figure 16: MMS SMART instrument suite architecture (image credit: SwRI, Ref. #

Hot Plasma composition suite: (FPI, DIS, DES, HPCA)

FPI (Fast Plasma Instrument): The FPI consists of DIS (Dual Ion Sensors) and DES (Dual Electron Sensors) and measures 3D ion and electron flux distributions over the energy range ~10 eV to 30 keV with an energy resolution of 20%. Electrons will be measured with a time resolution of 30 ms, ions with a time resolution 150 ms. The FPI development is led by Co-I Tom E. Moore (GSFC). FPI partners: SwRI, JAXA/ISAS, CESR, MSFC, Meisei Electric Co., SPEI.

Each of the MMS spacecraft will carry four FPIs, so there are 16 DIS instruments total. They will be paired with 16 DES (Dual Electron Spectrometers) and 6 IDPUs (Instrument Data Processing Units) that are being built at Goddard to complete the full FPI.

HPCA (Hot Plasma Composition Analyzer): The HPCA employs a novel RF technique to measure minor ions such as oxygen and helium in regions of high flux. Energy range = ~10 eV to 30 keV; energy resolution = 20%; time resolution = 15 s. HPCA development is led by Co-I David T. Young (SwRI).

The HPCA is a toroidal top-hat electrostatic analyzer coupled to a TOF (Time-of-Flight) mass spectrometer. The instrument will take quick measurements at the mass resolution that can accurately separate and identify the minute amounts of hydrogen, helium, and oxygen ions from the magnetosphere. Furthermore, the instrument will also have several features essential to space travel: robustness, low-power requirements, and minimal mass.


Energetic Particles: (FEEPS, EIS)

FEEPS (Fly's Eye Energetic Particle Sensor): The two FEEPS devices will measure 3D energetic ion and electron flux distributions over the energy ranges ~25 keV to 500 keV (electrons) and ~45 keV to 500 keV (ions). Time resolution = 10 s. FEEPS development is led by Co-I J. Bernard Blake (The Aerospace Corporation).

EIS (Energetic Ion Spectrometer): EIS uses a time-of-flight/pulse-height sensor to provide ion composition measurements (protons vs. oxygen ions) and angular distributions over the energy range ~45 keV to 500 keV and with a temporal resolution of 30 s. EIS development is led by Co-I Barry H. Mauk (Applied Physics Laboratory), who also heads the Energetic Particle investigation as a whole.



Figure 17: MMS instrument suite components (image credit: NASA)


FIELDS sensor suite: (AFG, DFG, EDI, SPD, ADP, SCM, CEB)

The FIELDS investigation is an advanced suite of six sensors to measure critical electric and magnetic fields in and around reconnection regions. The investigation is led by Co-I Roy B. Torbert of UNH (University of New Hampshire). 28) 29)

AFG (Analog Fluxgate) and DFG (Digital Fluxgate) Magnetometers: The AFG and DFG sensors are provided by UCLA and the Technical University of Braunschweig, respectively. C. T. Russell (UCLA) has overall responsibility for fluxgate development. The two different kinds of magnetometers will provide redundant measurements of the magnetic field and current structure in the diffusion region.

EDI (Electron Drift Instrument): The EDI determines the electric and magnetic fields by measuring the drift of ~1 keV electrons emitted from the GDU (Gun Detector Unit). Each GDU sends (receives) a coded beam to (from) the other EDI-GDU. The EDI gun is being developed at the Institut fuer Weltraumforschung of the Austrian Academy of Sciences; EDI optics are being developed at the University of Iowa.

Double Probe (SDP/ADP): MMS requires two sets of double-probe instruments:

- The SPD (Spin-plane Double Probe ) consists of four 48 m wire booms with spherical sensors at the end. The SDP assembly is provided by KTH (Royal Institute of Technology) Stockholm and by IRF-U (Swedish Institute of Space Physics-Upsalla), Sweden (Ref. 25).

No of units, mass, power

4, 1.5 kg/unit, 330 mW/unit

Measurement range, resolution, accuracy

DC-100 kHz, 1 ms, 0.3 mV/m

Total length of boom

50 m

Wire diameter

1.4 mm

Preamplifier case

2.5 cm x 8 cm


8 cm

Hinge point

1 m

Table 1: Parameters of SDP


Figure 18: Illustration of the SDP components (image credit: KTH, IRF-U)

- The ADP (Axial Double Probe), developed at LASP of the University of Colorado and the University of New Hampshire, consists of two 10 m antennas deployed axially near the spacecraft spin axis. The SDP and ADP provide full 3D electric field measurements over a range from DC to 100 kHz with an accuracy of 1 mV/m.

SCM (Search Coil Magnetometer): The SCM will measure the 3-axis AC magnetic field up to 6 kHz and will be used together with the ADP and SDP to determine the contribution of plasma waves to the turbulent dissipation that occurs in the diffusion region. The SCM is being developed at CETP (Centre d'etude des Environnements Terrestre et Planetaires) Velizy/ Saint-Maur, France.

• CEB (Central Electronics Box): The CEB provides power, control and data processing for the Fields sensor suite. KTH (Royal Institute of Technology) provides the power supply. UNH provides the CEB and the software with contributions from the sensor team institutions: KTH, LASP, UCLA and IWF.


ASPOC (Active Spacecraft Potential Control Device):

The ASPOC instrument on MMS has technical and scientific heritage from ESA’s Cluster and joint ESA/CNSA (Chinese National Space Administration) Double Star space missions. On each MMS spacecraft are two ASPOCs, the generated indium ion beams (energy range ~ 4-12 keV, currents up to 70 µA, antiparallel direction) limit the spacecraft potential to several volts positive. An instrument consists of the electronics box with the digital-, low voltage-, and high voltage boards (stacked frames) and two cylindrical emitter modules, all units are mechanically connected.

Figure 19 shows the instrument with electrically equivalent dummies replacing the real emitter modules. Each emitter module contains two ion emitters connected to a common high voltage supply. The cable harness is routed outside the electronics box and during operation only one ion emitter is switched active. In the final configuration the MLI (Multi-Layer Insulation) of the spacecraft is connected to the ion emitter modules via a plate, i.e. except the emitters the instrument is inside the spacecraft envelope mounted on the instrument deck. 30)


Figure 19: Close-up view of the ASPOC EQM instrument (image credit: ESA, IWF Graz)

EMC strategy: During development of the instrument – in order to get an in-house electromagnetic emission baseline – routine EMC pre-compliance measurements and sniff testing where performed in the frequency range up to 1 GHz (the digital clock of the instrument is 12 MHz). The ASPOC instruments for MMS — 8 flight models (FMs), and the engineering-qualification model (EQM) — have to be verified in EMC and magnetic tests. The test matrix (major items: structural / mechanical, EMC and magnetics, thermal / vacuum, calibration and general verifications) is slightly different for the EQM, FM1 and FM2-8, the basic principle is that all tests are in flight-like conditions. - The ASPOC EQM instrument for the MMS mission fulfills the EMC requirements.


Figure 20: Block diagram with the major components electronics box (digital-, low voltage-, and high voltage boards) and the two emitter modules (image credit: ESA, IWF Graz)


Figure 21: Overview of the MMS instruments suite 31)



Ground segment:

The MMS ground system supports on-orbit operations of the MMS observatories, as well as the production, storage, management, and dissemination of MMS science data products. The MMS ground system consists of the following functional elements:

• MOC (Mission Operations Center), located at GSFC (Goddard Space Flight Center) in Greenbelt, MD. Responsible for spacecraft operations and telemetry capture.

• FDOA (Flight Dynamics Operations Area), located at GSFC. Responsible for orbit and attitude determination and control.

• SOC (Science Operations Center), located at LASP (Laboratory for Atmospheric and Space Physics) of the University of Colorado, Boulder, CO. Responsible for IS (Instrument Suite) operations, instrument data processing, archiving, and distribution.

• SMART ITF (Instrument Team Facilities). Instrument teams are responsible for data analysis and validation; instrument monitoring and special operations requests; software for producing Quicklook and Level-2 data products; Level-2 data processing; analysis tools for publicly available data products.

• EPO (Education and Public Outreach), located at Rice University, Houston, TX. Responsible for dissemination of educational materials to schools and the general public.


Figure 22: Overview of MMS ground data system responsibilities (image credit: SwRI, Ref. #

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15) Conrad Schiff, Edwin Dove, “Monte Carlo Simulations of the Formation Flying Dynamics for the Magnetospheric Multiscale (MMS) Mission,” Proceedings of the 22nd International Symposium on Space Flight Dynamics (ISSFD), Feb. 28 - March 4, 2011, Sao Jose dos Campos, SP, Brazil, URL:

16) Michael Volle, Taesul Lee, Anne Long, Cheryl Gramling, Russell Carpenter, “Maneuver Recovery Analysis for the Magnetospheric MultiScale Mission,” 2008, URL:,d.bGE&cad=rja

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28) “FIELDS Instrument Suite,” UNH, URL:

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30) H. Eichelberger, G. Prattes, G. Fremuth, F. Giner, H. Jeszenszky, Ch. Kürbisch, M. Leichtfried, K. Torkar, “EMC Measurement from the instrument ASPOC aboard Magnetospheric Multiscale (MMS) mission,” Proceedings of 2012 ESA Workshop on Aerospace EMC, Venice, Italy, May 21-23, 2012, SP-702


The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.