Minimize Swarm

Swarm (Geomagnetic LEO Constellation)

Swarm is a minisatellite constellation mission within the Earth Explorer Opportunity Program of ESA, proposed under the lead of DNSC (Danish National Space Center) of Copenhagen, Denmark (formerly DSRI). In January 2007, DNSC became DTU Space, an institute at the Technical University of Denmark. The Swarm mission will be the 4th mission in ESA's Earth Explorer Program, following GOCE, SMOS, and CryoSat-2.

The first mission to ever map the Earth's magnetic field vector at LEO was the NASA MagSat spacecraft (launch Oct. 30 1979). Due to the low perigee (perigee=350 km, apogee=551 km), MagSat remained in orbit for only seven and a half months until June 11, 1980. About 20 years later, the Danish Ørsted micro satellite (1999-), the German CHAMP (2000-), the Argentine SAC-C (2000-) have been designed specifically for mapping the LEO magnetic field. Common to these recent missions is the magnetometry package, which utilizes a vector field magnetometer co-mounted with a star tracker (2 in the case of CHAMP) on an optical bench. As the accuracy of the instrument package has constantly increased, as well as the modelling methods have been improved towards optimized signal decomposition, it has been realized that simultaneous data from several points in space is needed, if the ultimate modelling barrier, the spatial-temporal ambiguity, has to be broken.

The overall objective of the Swarm mission is to build on the Ørsted and CHAMP mission experiences and to provide the best ever survey of the geomagnetic field (multi-point measurements) and its temporal evolution, to gain new insights into the Earth system by improving our understanding of the Earth's interior and climate. 1) 2) 3) 4) 5) 6) 7) 8) 9) 10) 11) 12) 13) 14) 15) 16) 17) 18)

This will be done by a constellation of three satellites, two will fly at a lower altitude, measuring the East-West gradient of the magnetic field, and one satellite will fly at a higher altitude in a different local time sector. Other measurements will also be made to complement the magnetic field measurements. Together these multipoint measurements will allow the deduction of information on a series of solid-Earth processes responsible for the creation of the fields measured.

Background on the discovery of electromagnetism:

The history of magnetic discovery goes back to about 110 B.C., when the earliest magnetic compass was invented by the Chinese. They noticed hat if a “lodestone” (natural magnets of iron-rich ore) was suspended so it could turn freely, it would always point in the same direction, toward the magnetic poles. This directional pointing property of magnetic material was eventually introduced into the making of an early compass and used for maritime navigation . By the 13th century, the directive property of magnetism was widely recognized and used in navigation. The mariner’s magnetic compass is the first technological application of magnetism and, one of the oldest scientific instruments.

Until 1820, the only magnetism known was that of iron magnets and of lodestones. It was the Danish physicist Hans Christian Ørsted, professor at the University of Copenhagen, who, in 1820, was first to discover the relationship between the hitherto separate fields of electricity and magnetism. Ørsted showed that a compass needle was deflected when an electric current passed through a wire, before Faraday had formulated the physical law that carries his name: the magnetic field produced is proportional to the intensity of the current. Magnetostatics is the study of static magnetic fields, i.e. fields which do not vary with time.19) 20)

Magnetic and electric fields together form the two components of electromagnetism. Electromagnetic waves can move freely through space, and also through most materials at pretty much every frequency band (radio waves, microwaves, infrared, visible light, ultraviolet light, X-rays and gamma rays). Electromagnetic fields therefore combine electric and magnetic force fields that may be natural (the Earth's magnetic field) or man-made (low frequencies such as electric power transmission lines and cables, or higher frequencies such as radio waves (including cell phones) or television (Ref. 21).


Figure 1: The early history of electromagnetic discovery made by scientists throughout the centuries (image credit: ESA, Ref. 16)


Background on the Earth's magnetic field:

The Earth has its own magnetic field, which acts like a giant magnet. Geomagnetism is the name given to the study of this field, which can be roughly described as a centered dipole whose axis is offset from the Earth's axis of rotation by an angle of about 11.5º. This angle varies over time in response to movements in the Earth's core. The angle between the direction of the magnetic and geographic north poles, called the magnetic declination, varies at different points on the Earth's surface. The angle that the magnetic field vector makes with the horizontal plane at any point on the Earth's surface is called the magnetic inclination.

This centered dipole exhibits magnetic field lines that run between the north and south poles. These field lines convergent and lie vertical to the Earth's surface at two points known as the magnetic poles, which are currently located in Canada and Adélie Land. Compass needles align themselves with the magnetic north pole (which corresponds to the south pole of the 'magnet' at the Earth's core).

The Earth's magnetic field is a result of the dynamo effect generated by movements in the planet's core, and is fairly weak at around 0.5 gauss, i.e. 5 x 10-5 tesla (this is the value in Paris, for example). The magnetic north pole actually 'wanders' over the surface of the Earth, changing its location by up to tens of km every year. Despite its weakness, the Earth's dipolar field nevertheless screen the Earth from charged particles and protect all life on the planet from the harmful effects of cosmic radiation. In common with other planets in our solar system, the Earth is surrounded by a magnetosphere that shields its surface from solar wind, although this solar wind does manage to distort the Earth's magnetic field lines.

The Earth’s magnetic field shows deviations, called anomalies, from the idealized field of a centered bar magnet. These anomalies can be quite large, affecting areas on a regional scale. One example is the SAA (South Atlantic Anomaly), which affects the amount of cosmic radiation reaching the passengers and crew of any plane and spacecraft led to cross it (Ref. 21).


Figure 2: Artist's view of solar wind interacting with Earth's magnetic field (image credit: DTU Space)


Figure 3: Schematic view of the geomagnetic field, produced mainly by a self-sustaining dynamo in the outer fluid core (image credit: GFZ)


Figure 4: Map of the geomagnetic field strength at the surface of the Earth derived from the model produced using data from the Oersted satellite (image credit: LETI) 21)


Figure 5: Magnetic field contributions (image credit: ESA) 22)


The primary research topics to be addressed by the Swarm mission include: 23)

• Core dynamics, geodynamo processes, and coremantle interaction. - The goal is to improve the models of the core field dynamics by ensuring long-term space observations with an even better spatial and temporal resolution. Combining existing Ørsted, CHAMP and future Swarm observations will also more generally allow the investigation of all magnetohydrodynamic phenomena potentially affecting the core on sub-annual to decadal scales, down to wavelengths of about 2000 km. Of particular interest are those phenomena responsible for field changes that cannot be accounted for by core surface flow models.24)

• Lithospheric magnetization and its geological interpretation. - The increased resolution of the Swarm satellite constellation will allow, for the first time, the identification from satellite altitude of the oceanic magnetic stripes corresponding to periods of reversing magnetic polarity. Such a global mapping is important because the sparse data coverage in the southern oceans has been a severe limitation regarding our understanding of plate tectonics in the oceanic lithosphere. Another important implication of improved resolution of the lithospheric magnetic field is the possibility to derive global maps of heat flux.25) 26)

• 3-D electrical conductivity of the mantle. - Our knowledge of the physical and chemical properties of the mantle can be significantly improved if we know its electrical conductivity. Due to the sparse and inhomogeneous distribution of geomagnetic observatories, with only few in oceanic regions, a true global picture of mantle conductivity can only be obtained from space.

• Currents flowing in the magnetosphere and ionosphere. - Simultaneous measurements at different altitudes and local times, as foreseen with the Swarm mission, will allow better separation of internal and external sources, thereby improving geomagnetic field models. In addition to the benefit of internal field research, a better description of the external magnetic field contributions is of direct interest to the science community, in particular for space weather research and applications. The local time distribution of simultaneous data will foster the development of new methods of co-estimating the internal and external contributions.

The secondary research objectives include:

• Identification of the ocean circulation by its magnetic signature. - Moving sea-water produces a magnetic field, the signature of which contributes to the magnetic field at satellite altitude. Based on state-of-the-art ocean circulation and conductivity models it has been demonstrated that the expected field amplitudes are well within the resolution of the Swarm satellites. 27)

• Quantification of the magnetic forcing of the upper atmosphere. - The geomagnetic field exerts a direct control on the dynamics of the ionized and neutral particles in the upper atmosphere, which may even have some influence on the lower atmosphere. With the dedicated set of instruments, each of the Swarm satellites will be able to acquire high-resolution and simultaneous in-situ measurements of the interacting fields and particles, which are the key to understanding the system.

Historic background of Swarm: Ref. 14)

• The first Swarm proposal was made in 1998, prior to launch of the Ørsted mission.

• In early 2002, the Swarm mission was proposed to ESA by Eigil Friis-Christensen of DNSC (Copenhagen, Denmark), Hermann Lühr of GFZ (GeoForschungszentrum, Potsdam, Germany), and Gauthier Hulot of IPG (Institut de Physique du Globe, Paris, France) with support from scientists in seven European countries and the USA. In the meantime, the Swarm team comprises participation of 27 institutes on a global scale. The mission was selected for feasibility studies in 2002. The initial mission proposal considered a Swarm constellation of 4 spacecraft. 28)

• In May 2002 there were three mission candidates: ACE+, EGPM and Swarm; they were chosen for a feasibility study.

• At the end of two parallel feasibility studies, the Swarm mission was selected as the 5th mission in ESA's Earth Explorer Program in May 2004. Phase A was completed in Nov. 2005, resulting in a constellation of 3 spacecraft.

New Concept – Constellation to characterize external sources:

- The external contributions are highly influenced by solar activity and local time

- Simultaneous satellites in different orbital planes are necessary in order to overcome the time-space ambiguity in the measurements. The optimum constellation depends on the scientific objectives.

- But, measurements of high accuracy are not sufficient! A better understanding of the various sources is equally important, in particular when doing measurements with unprecedented precision, where new phenomena appear in the data. For this, additional and independent key information is needed: a) electric field, b) ionospheric conductivity.

• In 2006, the Swarm project was in Phase B, ending with the PDR (Preliminary Design Review) in the summer 2007.

The construction of the Swarm constellation commenced in November 2007 with the Phase C/D kick-off meeting. The Swarm project CDR (Critical Design Review) took place on Oct. 14, 2008 at ESA/ESTEC. 29)


Figure 6: Schematic view of Earth's magnetic field (image credit: ESA/ATG Medialab) 30)

Legend to Figure 6: The magnetic field and electric currents near Earth generate complex forces that have immeasurable impact on our everyday lives. Although we know that the magnetic field originates from several sources, exactly how it is generated and why it changes is not yet fully understood. ESA’s Swarm mission will help untangle the complexities of the field.



Space segment concept:

The Swarm mission architecture is driven by the requirement for separation of the various sources contributing to the Earth's magnetic field. Hence, the space segment concept employs a three-minisatellite constellation with the following characteristics:

- Three spacecraft in two different orbital planes, with two satellites in a plane of 84.7º inclination and with one satellite in a plane of 88º inclination

- The two satellites in the 87.4º inclination orbit will fly at a mean altitude of 450 km, their east-west separation will be 1-1.5º, and the maximum differential delay in orbit will be about 10 s.

- The satellite in the higher inclination orbit (88º) will fly at a mean altitude of 530 km.

- The spacecraft require some degree of active orbit maintenance to control the relative positions in the constellation (this is an element of formation flight to support flight operations). 31) 32)

In November 2005, ESA selected EADS Astrium GmbH, Friedrichshafen, Germany as the prime contractor for the Swarm spacecrafts. The Swarm consortium (main subcontractors) consists of: 33)

- EADS Astrium Ltd., UK (mechanical, thermal, AIV)

- GFZ Potsdam, Germany (end-to-end system simulator, calibration & validation)

- DTU Space, Copenhagen, Denmark [level 1b processor and instruments (VFM magnetometer and STR star tracker)]

The spacecraft design is governed by the following requirements:

1) Magnetic cleanliness: magnetometers on deployable boom, non-magnetic materials and caution during handling

2) Magnetic field vector attitude knowledge: ultra-stable connection between VFM (Vector Field Magnetometer) and STR (Star Tracker) assembly on the optical bench

3) Ballistic coefficient: small ram surface in flight direction to minimize air drag

4) Accelerometer proof-mass vs satellite CoG (Center of Gravity) location.


Figure 7: Artist's rendition of the Swarm constellation (image credit: ESA)

An important design measure is the accommodation of the magnetometer package at a distance from the main body/platform sufficient to minimize any magnetic disturbance. A boom ensures a magnetically 'clean' environment and provides very stable accommodation for the magnetometer package. Due to envelope constraints of the launcher fairing, the boom must be deployable. 34)

Optical bench: The vector magnetometer is mounted on an ultra-stable silicon carbide-carbon fiber compound structure (the SWARM optical bench). Both optical bench and scalar magnetometer are installed on a deployable conical tube of square cross section. The position tolerance of the optical bench to its tube interface has to be fixed within 0.2 mm. 35) 36)

The design driver of the composite tube assembly of Swarm is thermal stability. The main cause for observed thermal distortion is the non-uniformity of the cross-sections arising from the different adaptations of the filament winding process in order to manufacture the carbon fiber reinforced structure. The manufacture of the structure required use of thermally controlled high precision bonding jigs to join the composite tubes to the metallic fittings.

The scalar magnetometer and optical bench are fixed to a deployable large beam of square cross section, the SWARM (Carbon-fiber Tube Assembly (CTA) which fulfils the following main functions (Figure 8):

• Separate the sensitive instruments from the spacecraft to comply with the very high magnetic cleanliness requirements

• Provide a suitable stable structure for the fixation of instruments.

The chosen manufacturing technology for the SWARM tube was filament winding. The SWARM tube has a conical taper. Since the amount of fibers in a cross section is constant the tube had two main characteristics: the wall thickness increased linearly from the root to the tip and due to nature of the winding process the fiber angle became steeper at the tip than at the root. The overall effect is a variation of properties along the length of the tube.

The Swarm carbon-fiber tube assembly was subjected to various tests: Thermal distortion was measured by establishing a 65ºC gradient between the tip and the hinge and a 10ºC gradient between opposite sides of the CTA. The hole pattern of the optical bench was accurate to within 0.2 mm (Ref. 35).


Figure 8: The Swarm optical bench, carbon-fiber tube assembly (image credit: RUAG)


Figure 9: Configuration and performance requirements of a Swarm spacecraft (image credit: EADS Astrium)


Figure 10: Configuration details of a Swarm spacecraft (image credit: EADS Astrium)

The three identical Swarm minisatellites consist of the payload and the platform elements. The platform comprises the following subsystems: structure/mechanisms, power, RF communications, AOCS (Attitude and Orbit Control Subsystem), thermal control, and onboard data handling.

The AOCS design is based to a maximum extent on the CryoSat AOCS design of EADS Astrium. The gyro-less AOCS provides 3-axis stabilization with an Earth pointing attitude control in all modes. The requirements call for: 37)

- An attitude pointing control within a band of < 5º about all axis (roll, pitch, and yaw), the pointing stability is < 0.1º/s

- Provision of a sufficient torque capability for launcher tip-off rate damping and attitude acquisition

- Minimize acceleration and magnetic stray field disturbances to scientific instruments

- Provision of a high ΔV capability for orbit & attitude control maneuvers.


Figure 11: Functional architecture of AOCS (image credit: EADS Astrium)

The AOCS is tightly coupled with the propulsion subsystem. Actuation is provided by a cold gas propulsion subsystem, referred to as OCS (Orbit Control Subsystem), and magnetic torquers (used for ΔV maneuvers and to complement the magnetic torquers). The cold gas propulsion system is provided by AMPAC-ISP, UK. - Attitude sensing is provided by a star tracker assembly (3 star tracker heads), 3 magnetometers, and a CESS (Coarse Earth and Sun Sensor) assembly used in safe mode situations and in initial acquisition sequences, respectively (CESS is of CHAMP, GRACE, and TerraSAR-X heritage). A dual frequency GPS receiver (GPSR) is used to provide PPS (Precise Positioning Service) to the OBC and instruments for on-board datation.

Note: the star tracker (STR) assembly and optical bench are described below under a separate heading.

The nominal attitude has a nadir orientation. Rotation maneuvers of S/C about roll, pitch and yaw are used for instrument calibration and orbit Control. The safe mode is Earth-oriented. Pointing requirements are 2º about all axes, with limitations on use of actuators.

The Swarm rate damping design, in support of the critical spacecraft deployment phase, employs magnetic rate damping - magnetometers in combination with magnetic torquers and thrusters - to provide a significantly cheaper implementation than with the use of gyroscopes. From a control theory point-of-view, rate damping with magnetometers using 2-axis measurement is as “safe” as with gyroscopes using 3-axis measurement: Global asymptotical stability is achieved except for the case when the magnetic field does not change. This is only in near-equator orbits possible with perfect field symmetry which is in practice not realistic. The result is confirmed by the evaluation of the observability criterion where no loss of this property could be detected except for the mentioned case. Since SWARM is in a polar inclination orbit, the control concept is considered “clean”. 38)

Rate damping design: The RDM controller is a simple proportional controller on the S/C rate with reference rate zero. The S/C rate is computed by processing and derivation of the FGM measurements. The controller outputs the torque commands for the torquer and the thruster. A dead band for the thruster inhibits the thruster activation for low rates which can be covered by the torque rod.


Figure 12: Schematic of the Swarm RDM controller (image credit: EADS Astrium)


Each spacecraft features 2 propellant tanks, each with a capacity of 30 kg of N2. The thrusters provide thrust levels of 20 and 40 mN. The cold gas thruster system was developed and space qualified by Ampec-ISP, Cheltenham, UK consisting of 24 OCT (Orbit Control Trusters) and 48 ACT (Attitude Control Thrusters) for the Swarm constellation. The assembly and test of Ampac's SVT01 series of cold gas thrusters has included design modifications, full qualification and verification of suitability to operate with a new propellant. In 2010, a set of 72 units has been supplied and integrated into the constellation of three Swarm spacecraft. 39)

A GPS receiver provides the functions of timing and position determination. The spacecraft dry mass is about 370 kg.


Figure 13: Software architecture of AOCS (image credit: EADS Astrium)

EPS (Electrical Power Subsystem): The two body-mounted solar arrays and the varying orbits of the satellites require a MPPT (Maximum Power Point Tracking) system. Important requirements are related to the magnetic cleanliness of the satellites and result in following specific PCDU (Power Conditioning and Distribution Unit) design requirements: 40)

- Minimization of magnetic moment i.e. minimizing of magnetic materials and current loops

- Selection of switching frequencies outside the ‘forbidden’ frequency ranges

- Minimizing spacecraft surface charging by use of negative bus voltage concept (battery + is connected to spacecraft structure).

The PCU part of the PCDU covers all tasks to control the power flow in the unit from the different sources and performs the communication with the OBC (On Board Computer).
During eclipse and battery recharge mode, the bus voltage varies with the state of charge of the battery. In taper charge mode, the bus is controlled by the MEA (Main Error Amplifier) to a predefined (commandable) value.

The main power requirements for the PCDU are defined as follows:

- Solar array input: 0 to -125 V, max. 21 A (each of 2 panels)

- Maximum power per panel: 750 W

- Main bus voltage range -22 V to -34 V

- Maximum battery charge current 24 A

- Continuous discharge current 0 to 14 A.

Maximum discharge current/power up to 0.5 h: 20 A / 440 W.


Figure 14: Architecture of the PCDU (image credit: EADS Astrium)

Negative bus voltage concept: The Swarm satellite requires the positive line of the power system connected to structure. This implies that all bus protection functions have to be allocated in the ‘hot’ negative line. As all essential functions, ( i.e. bus voltage control) need to be independent from the auxiliary supplies, they have to be supplied by the negative bus voltage. Figure 15 shows a principle grounding/power supply diagram of the main functional blocks in the PCDU.

Power control concept: The PCDU uses a simple concept for control of the battery state of charge and the bus voltage:

- Whenever the bus voltage and the charge current are below the limits, the MPPTs are active

- When the either the bus voltage attains the ‘battery end-of-charge voltage or the battery attains the charge current limit, the MEA (Main Error Amplifier) supersedes the tracker operation.

A bus overvoltage detection logic has been implemented in the PCDU, which performs a rapid ramp-down of the solar regulator current by using hysteresis control.

The MEA is composed of 3 identical separated control stages and a majority voter. Each control stage has a dedicated set of sensors and receives the relevant set commands for the bus voltage via redundant internal control busses. The charge current limitation is implemented in a ‘cascade configuration’, using the output of the current error amplifier as a set signal for the voltage amplifier. This assures a low and constant bus impedance during all MEA control modes. The implementation of the regulation concept is given in Figure 16.


Figure 15: Grounding scheme of the Swarm PCDU (image credit: EADS Astrium)


Figure 16: Schematic of the bus control concept (image credit: EADS Astrium)

Spacecraft mass

Dry mass of ~369 kg
Cold gas propellant: 99 kg of CF4 (Freon)
Total mass = 468 kg

Spacecraft dimensions

Length: 9.1 m; width: 1.5 m (S/C body); height: 0.85 m; ram surface: ~0.7 m2

Boom length

5.1 m


- 3-axis stabilized; magnetometers; CESS; GPS; STR, magnetorquers; thrusters
- 3D position better than 20 m (3σ)
- 3D velocity better than 1 m/s (3σ)
- UTC time with respect to GPS system time
- Datation of PPS signal better than 0.5 µs (3σ)

AOCS sensors

AOCS actuators

- STR (Star Tracker) with 3 sensor heads
- 1 CESS (Coarse Earth & Sun Sensor) with 6 heads placed orthogonal on the S/C
and 3 Magnetometers (FGM) which can be used for rates up to 0.5º/s.
- 3 MTQ (Magnetic Torquer), each 10 Am2
- 24 Cold Gas Thruster (THR), of which 2 x 8 for attitude control in all 3 axes, each 20 mN
force and 2 x 4 for orbit control, placed in -x and +y direction, each 50 mN force
- MTQs and THRs are used by each control mode

AOCS control modes

- Rate damping: rates are measured by the FGMs, main actuation by THR
- Coarse pointing: power and thermal safe Earth pointing attitude using CESS
- Fine pointing: STR and GPSR are used for attitude and position knowledge
- Orbit Control: similar to FPM, additionally performing slews for instrument calibration
and for orbit change and maintenance which requires using orbit control thruster.

EPS (Electrical Power Subsystem)

Total power: 608 W nominal; solar cells: GaAs triple junction; solar panel positive grounding; a set of batteries: Li-ion with a capacity of 48 Ah

RF communications

S-band; downlink data rate: 6 Mbit/s; 4 kbit/s uplink, data volume: 1.8 Gbit/day; 1 dump/day to Kiruna ground station, data storage capability: 2 x 16 Gbit

Mission duration

3 months of commissioning followed by 4 years of nominal operations

Table 1: Overview of spacecraft parameters

RF communications: S-band for TT&C spacecraft monitoring services and for science data transmission.


Figure 17: Artist's rendition of the Swarm constellation in orbit (image credit: EADS Astrium)


Figure 18: Photo of the three Swarm satellites at the launch site (image credit: ESA)


Figure 19: The Swarm satellites separated by a few centimeters (image credit: ESA, M. Shafiq)

Legend to Figure 19: Attached to the tailor-made launch adapter, the three Swarm satellites sit just centimeters apart. This novel part of the rocket keeps the satellites upright within the fairing during launch and allows them to be injected simultaneously into orbit. 41)

DBA (Deployable Boom Assembly): The Swarm DBA, consisting of a 4.3m long CFRP tube and a hinge assembly, is designed to perform this function by deploying the CFRP tube plus the instruments mounted on it. mounted on a 4.3m long deployable boom. Deployment is initiated by releasing 3 HDRMs (Hold Down Release Mechanisms) , once released the boom oscillates back and forth on a pair of pivots, similar to a restaurant kitchen door hinge, for around 120 seconds before coming to rest on 3 kinematic mounts which are used to provide an accurate reference location in the deployed position. The motion of the boom is damped through a combination of friction, spring hysteresis and flexing of the 120+ cables crossing the hinge. Considerable development work and accurate numerical modelling of the hinge motion was required to predict performance across a wide temperature range and ensure that during the 1st overshoot the boom did not damage itself, the harness or the spacecraft. - Due to the magnetic cleanliness requirements of the spacecraft, no magnetic materials could be used in the design of the hardware. 42)


Launch: The Swarm constellation was launched on Nov. 22, 2013 (12:02:29 UTC) on a Rockot vehicle from the Plesetsk Cosmodrome, Russia. The launch was provided by Eurockot Launch Services. Some 91 minutes after liftoff, the Breeze-KM upper stage released the three satellites into a near-polar circular orbit at an altitude of 490 km.43) 44) 45) 46) 47)

The launch was planned for the fall of 2012, but due to the recent Breeze-M (Briz-M) failure the launch was postponed to permit proper investigations of the cause. In Nov. 2012, ESA is still expecting, from the Russian Ministry of Defence, the launch manifest for the year 2012/13 for Rockot launchers indicating the launch date for Swarm. 48) 49)

Note: Rockot, a converted SS-19 ballistic missile, has been grounded since February 1, 2011 when the Rockot vehicle with the Breeze-KM upper stage failed to place the Russian government’s GEO-IK2 geodesy satellite of 1400 kg (Kosmos 2470) into its intended orbit of 1000 km. However, in the meantime, the Rockot/Breeze-KM vehicle demonstrated its reliability by lifting 4 Russian spacecraft (Gonets-M No.3, Gonets-M No.4, Strela-3/Rodnik, and Yubileiny-2/MiR) successfully into orbit on July 28, 2012.

On April 9, 2010, ESA awarded a contract to Eurockot, for the launch of two of its Earth observation missions. The contract covers the launch of ESA's Swarm magnetic-field mission and a 'ticket' for one other mission, yet to be decided. Both will take place from the Plesetsk Cosmodrome in northern Russia using a Rockot/Breeze-KM launcher. Eurockot is based in Bremen, Germany and is a joint venture between Astrium and the Khrunichev Space Center, Moscow. 50) 51) 52) 53)

After release from a single launcher, a side-by-side flying lower pair of satellites at an initial altitude of 460 km and a single higher satellite at 530 km will form the Swarm constellation. The constellation deployment and maintenance require a total ΔV effort of about 100 m/s.

In LEOP (Launch and Early Orbit Phase), at least three ground stations will be involved. LEOP is expected to last 3 days for the full activation of the satellites, followed by an orbit acquisition phase of up to three months. In parallel with the orbit acquisition phase, the commissioning phase will start in order to check out all satellite subsystems and the payload. The commissioning phase is currently expected to last three months. After the commissioning phase the nominal mission phase of 4 year starts.



Orbits of the constellation:

Accurate determination and separation of the large-scale magnetospheric field, which is essential for better separation of core and lithospheric fields, and for induction studies, requires that the orbital planes of the spacecraft are separated by 3 to 9 hours in local time. For improving the resolution of lithospheric magnetization mapping, the satellites should fly at low altitudes - thus experiencing some drag, but commensurate with the goals of a multi-year mission lifetime. The three satellites are being flown in 3 orbital planes with 2 different near-polar inclinations to provide a mutual orbital drift over time (Figure 20 and 21).

• Two satellites (Swarm A+B) are in a similar plane of 87.4º inclination. The satellite pair of 87.4º inclination will fly at a mean altitude of 450 km, their east-west separation shall be 1-1.4º, and the maximal differential delay in orbit shall be about 10 s. The formation-flying aspects concern the satellite pair, a side-by-side formation, requiring some formation maintenance.

• One higher orbit satellite (Swarm C) in a circular orbit with 88º inclination at an initial altitude of 530 km. The right ascension of the ascending node is drifting somewhat slower than the two other satellites, thus building up a difference of 9 hours in local time after 4 years.





Orbital altitude

≤ 460 km (initial altitude of satellite pair)

≤ 530 km

Orbital inclination



ΔRAAN (Right Ascension of Ascending Node)

1.4º difference between A and B

~0-135º difference
wrt mean plane of A/B
(continuous drift)

Mean anomaly at epoch

Δt = 2-10 s difference between A and B


LTAN evolution (Figure 20, right-hand side)

24 hours of local time coverage every 7-10 months
9 hours of separation between lower pair and upper satellite at end of life

Table 2: Overview of the Swarm orbit configuration


Figure 20: Orbit altitude projection over mission time (left); Local time evolution of the S/C in two orbital planes (right), image credit: DTU Space


Figure 21: Equatorial projection of the Swarm orbit configuration over time (image credit: DTU Space) 54) 55)


Figure 22: Polar projection of the Swarm orbit configuration over time (image credit: ESA)


Figure 23: The Swarm tandem pair provides a stereo view (image credit: DTU Space)



Mission status:

• February 2014: Since the Swarm constellation was launched last November, engineers have been busy putting the satellites through their paces to make sure that the craft and instruments are working correctly. This commissioning phase is an essential part of the mission before it starts providing data to further our understanding of the complex and constantly changing magnetic field. 56)

- Some tricky maneuvers are now under way to steer the trio of Swarm satellites into their respective orbits so that they can start delivering the best-ever survey of Earth's magnetic field.

- Since the intensity of solar activity is currently lower than anticipated, the original plan of where to place the satellites at the beginning of science operations has been reviewed recently by the scientific community and experts in ESA. Low solar activity means the satellites experience lower atmospheric drag, as clearly demonstrated by ESA’s GOCE mission.

• On Nov. 26, 2013, a significant milestone for ESA’s magnetic field mission was reached, the Swarm satellites have completed the critical first phase of their new mission. 57)

• Immediately following separation, the three satellites started transmitting their first signals to Earth, marking the start of the critical ‘launch and early orbit phase’, known as LEOP (Launch and Early Orbit Phase). After separation, the project acquired signals from the first two Swarm satellites 91 minutes into the mission, followed by the third at the 95-minute mark — starting the LEOP phase. Around 23 hours GMT on Nov. 22, each of the three satellites deployed their 4 m-long boom carrying instruments essential to the mission’s scientific success. LEOP was formally declared as completed by Flight Operations Director Pier Paolo Emanuelli on Nov. 24, 2013 at 19:30 GMT (Ref. 57).

• First contact was established with the trio a few minutes after deployment through the Kiruna station in Sweden and the Svalbard station in Spitsbergen, Norway. All three satellites are controlled by ESA teams at ESOC (European Space Operation Center) in Darmstadt, Germany (Ref. 43).



Sensor complement: (VFM, ASM, CEFI, MAC-04 accelerometer, LRR)

High-precision and high-resolution measurements of the strength, direction and variation of the magnetic field, complemented by precise navigation, accelerometer and electric field measurements, will provide the necessary observations that are required to separate and model various sources of the geomagnetic field. 58) 59)

The observation concept is mainly determined by the following drivers set by the scientific payload: Each of the satellites carries an identical payload:

• High magnetic cleanliness required by the magnetometers (sub nT-range for VFM and ASM)

• EFI (Electrical Field Instrument) requires in-flight pointing with control accuracy of 5º

• ACC (Accelerometer) requires precise and stable accommodation in CoG (Center of Gravity).


Figure 24: Schematic overview of the Swarm sensor complement (image credit: GFZ Potsdam)

Magnetic Field (all values 2σ)

- In-Situ magnitude with a random error < 0.3 nT
- Stability of magnitude < 0.12 nT (2σ) over 3 months
- Vector components with a random error < 1 nT
- Stability of vector components better than 1 nT/year

Attitude knowledge

better than 0.1º

Satellite position

- POD (Precise Orbit Determination) < 10 cm (rms) – L2 products
- MOD (Medium Orbit Determination) < 1.5 m (rms) – L1b products

Air drag (all values 2σ)

Vector components with a random error < 5 x 10-8 ms-2

Electrical field (all values 2σ)

- Vector Components with a random error < 10 mV/m
- Stability < 1 mV/m over 1 month
- Ionosphere plasma density with an accuracy < 20% (ρ=109 -1011 m-3) – 5% (ρ >1011 m-3)
- Ion & electron temperature with an accuracy < 20% (ρ >1010 m-3)
- Ion Drift Velocity Vector with a random error < 200 m/s

Table 3: Key performance requirements


VFM (Vector Field Magnetometer):

VFM is the prime instrument of the Swarm mission developed at DTU Space. The objective is to measure the magnetic field vector, on the boom, together with the star tracker for precise attitude measurement. The boom mounted Swarm vector magnetometer instrument consist of a triple star sensor block and a CSC (Compact Spherical Coil) vector magnetometer sensor, mounted on a stable optical bench (Figure 8). Each satellite contains the optical bench with one CSC and three CHU (Camera Head Unit). 60) 61) 62)

The three star sensor units are arranged with the boresights 90º from each other so as to ensure that only one CHU may be affected by Sun or Moon intrusion at any given time. Hereby an attitude solution accurate in all three degrees of freedom can be delivered to the CSC throughout the entire mission. The CSC sensor and the triple star sensor block are mounted on either end of a highly stable mechanical structure.

The CSC vector sensor is supported by a zero CTE (Coefficient of Thermal Expansion) CFRP (Carbon Fiber Reinforced Polymer) adapter that on the one end matches the zero CTE CFRP tube, used to displace the CSC sensor from the star sensor heads (CHU), and on the other end matches the 32 ppm CTE CSC sensor, by means of a finger section. The rotational symmetry of this design ensures an excellent angular stability.

The other end of the CFRP tube is attached to a CSiC bracket holding the three CHUs. The CSiC exhibit a heat distribution capacity second to none, minimizing thermal biases of this section, from the inevitable thermal gradient induced when the sun happens to illuminate any of the three CHUs. Because the CSiC is weakly magnetic, this material can only be used at distances larger than 20 cm from the CSC sensor.

Each CHU is fitted with a straylight suppression system that is thermally decoupled from the optical bench. This separation minimizes thermal excursions from the time varying sun impingement over an orbit to less than a few degrees C. The straylight suppression system is mechanically mounted on an external thermal CFRP shroud, which also provides for thermal control of the entire optical bench. The material selection for all thermal protection has been performed to suppress soft or hard magnetic parts as well as parts that can generate magnetic fields under thermal gradients.

VFM instrument: The VFM (fluxgate type) is based on the fluxgate transducer using a ringcore with amorphous magnetic material, which has a very low noise (10-20 pT rms). It has an extremely high stability < 0.05 nT/year. VFM consists of a CSC (Compact Spherical Coil) sensor, non redundant, mounted on the deployable boom, an internally redundant data processing unit (DPU) and the connecting harness. The spherical coils that create a homogeneous vector field inside the sphere are mounted on an isotropic and extremely stable mechanical support. In feedback conditions the sensor is used as a nulling device and the coils define uniquely the magnetic axes of the sensor. The VFM exhibits high linearity (< 1ppm), a component accuracy of 0.5 nT and precision of 50 pT rms.

The operation of the fluxgate sensor is based on the extreme symmetry of the positive and negative magnetic saturation levels of the ferromagnetic sensor core material. Continuous probing of the core saturation levels by a high frequency excitation magnetization current enables the sensor to detect deviations from the zero field with only tens of pT noise and sub-nT long term stability.

The mounting of the VFM sensor is using a sliced adaptor ring. The optical bench ensures mechanical stability of the system. Three star trackers provide full accuracy attitude.

Instrument mass, power consumption

1 kg, 1 W

Dimension of sensor head (CSC)
Mass, power

82 mm Ø
280 g, ~ 250 mW

Dimension of DPU
Mass of DPU, power

100 x 100 x 60 mm
750 g, ~1 W

Data rate


Dynamic range

±65536.0 nT to 0.0625nT (21 bit)

Omnidirectional linearity

±0.0001% of full scale (±0.1nT in ±65536nT)

Intrinsic sensor noise

15 pTRMS in the band 0.01-10 Hz (6.6 pTRMS Hz-1/2 at 1 Hz)

Intrinsic electronics noise

50 pTRMS in the band 0.01-10 Hz (15 pTRMS Hz-1/2 at 1 Hz)

Sampling rate

50 Hz, linear phase filter, -3dB frequency 13.1 Hz

Temperature range

-20ºC to +40ºC (Operating performance)
-40ºC to +50ºC (Survival performance)

Thermal behavior
- Offset
- Scale factors
- Non-orthogonality angles

~0 nT/ºC (CSC), ~0.1 nT/ºC (electronics)
~10 ppm/ºC (CSC), ~2 ppm/ºC (electronics)
~0 arcsec/ºC (0.06, 0.07, 0.04)

Zero stability (thermal & long term)

< ± 0.5 nT

Absolute accuracy of Ørsted magnetometer parameters (relative to ASM & STR):

- Offset

< 0.2 nT (~120 dB)

- Scale factors

< 0.0005%

- Axes orthogonality

< 0.0006º (~2 arcsec)

- Axis alignment

< 0.0002º (~7 arcsec)

Ørsted magnetometer with 3 offsets, 3 scale factors & 3 angles for 6.5year:


< 0.5 nT

Table 4: Specification of the VFM instrument


Figure 25: The VFM flight model with redundant electronics unit or DPU (left) and CSC sensor (right), image credit: DTU Space

The µASC (micro Advanced Stellar Compass) of DTU Space provides the high accuracy, inertial attitude determination for the Swarm vector magnetometer. The microASC is a fully autonomous, internally hot/cold redundant star tracker, featuring up to four cameras. The microASC features a split DPU (Data Processing Unit) and CHU (Camera Head Unit) enabling the low power dissipation and very low magnetic disturbance CHU, to be placed close to many types of science instruments, including the CSC sensor (see µASC description below).

Inter-calibration: determining the internal angels between the CSC and the three CHUs: The optical bench provides a mechanically stable platform for the CSC and the three CHUs and will ideally fixate the internal angles between these. The prime objective of the inter-calibration is to establish the internal angles with the highest possible accuracy. The measurement frame of the CSC sensor is defined by the orientation of the compensation coils on the outside of the CSC sensor sphere. These coils form a nearly orthogonal triad, which has been thoroughly calibrated and orthogonalized prior to the mounting on the bench structure. Similarly the measurement frame of any of the CHUs is defined by the mechanical arrangement of the optics relative to the CCD sensor of the unit. Also this measurement frame has been established prior mounting the unit on the bench.

Despite the effort to minimize thermo-elastic deformations and the effort to make the platform as stiff and stable as possible, small residual variations exists. A secondary objective for the inter-calibration is therefore to assess the size of these residual errors, e.g. gravity release effects.

Finally, the mounting of the sensor units to the stiff optical bench will cause stresses to be built into the mounting interfaces. These stresses may cause minute changes to the internal calibration of the sensors. A third objective of the inter-calibration is therefore to verify the pre mounting calibration of the sensor units.


ASM (Absolute Scalar Magnetometer):

ASM is provided by CNES (French Space Agency) and CEA-LETI (French Atomic Energy Commission - Laboratoire d'Electronique de Technologie et d'Instrumentation), Grenoble, France. The objective is to measure magnetic field strength and to calibrate the VFM device to maintain the absolute accuracy during the multi-year mission. ASM is positioned at the very tip of the boom. The required main performance characteristics of the ASM are: absolute accuracy of < 0.3 nT (2σ), resolution < 0.1 nT within its full-scale range of 15000-65000 nT. 63) 64) 65) 66) 67) 68) 69) 70) 71)

Measurement concept: To overcome the limitations of the OVM (Overhauser Magnetometers) identified during the Oersted and Champ programs, a new magnetometer has been designed for the Swarm mission. The ASM pumped helium magnetometer relies on a low pressure helium vapor as the sensing medium (Figure 26), with the optical pumping process the counterpart of the dynamic nuclear polarization. 72) 73)

One important difference is however due to the fact that the optical pumping is a much more efficient polarization method, leading to an almost complete polarization. As a consequence, the signal amplitude does no longer depend on the magnetic field strength and a resolution of 1 pT/ (Hz)1/2 is now obtained over the complete measurement range.


Figure 26: Relevant helium energy levels involved in the ASM magnetometer (image credit: LETI)

As compared to most optically pumped magnetometers, the ASM operates with linearly polarized pumping light instead of circularly polarized light. The main reasons for that choice are the following:

- the strong interaction between the laser pumping beam and the helium atoms can in general affect their energy level and result in so-called light shifts whenever the pumping light wavelength is detuned from the helium transition center wavelength. Now using linearly polarized light suppresses this effect, thus significantly increasing the instrument’s accuracy.

- the key parameter governing the optical pumping angular dependence is then the direction of the laser polarization, whereas it is the propagation direction of the pumping beam that matters in circularly polarized light. Now when trying to design an isotropic instrument, i.e an instrument whose performances are independent of the sensor attitude, it is obviously easier to control the direction of the linear polarization than to rotate the whole sensor in order to align it properly with respect to the magnetic field direction. In our case the isotropy is thus simply achieved thanks to the use of an amagnetic piezoelectric motor which permanently controls the laser polarization and the RF magnetic field directions so that they are both perpendicular to the static magnetic field. The resulting magnetometer architecture is illustrated in Figure 27.


Figure 27: Isotropic helium magnetometer architecture (image credit: LRTI)

Contrary to the Overhauser solution based on a design trade-off between instrument’s resolution and omnidirectionality, the helium magnetometer is always operated in the optimal operational conditions thanks to this servo loop, but this is achieved at the expense of the use of a dedicated mechanism. As for the sensor anisotropy, resulting from a combination of induced and remanent contributions, a typical signature corresponding to the flight configuration is presented in Figure 28.

As for the environment susceptibility, the ASM significantly broader resonance line (close to 70 nT as compared to less than 7 nT for the OVM) reduces the impact of inhomogeneous magnetic fields on the magnetometer performances, while the principles of operation and architecture of the helium device makes it robust to low frequency radiated magnetic fields, thus making the EMC specifications much easier to meet in that respect.

Last but not least, the short metastable helium relaxation time (of the order of one millisecond) results in a much higher bandwidth for the helium magnetometer than was the case for the NMR sensors. While this feature is of no direct interest for the calibration of the vector instruments (the scalar data are sampled for that purpose at a fixed 1 Hz frequency), it opens new opportunities for the exploitation of the scalar instrument.


Figure 28: Residual anisotropy for the in orbit ASM configuration (image credit: LETI, Ref.72)


Instrument: The ASM magnetometer is based on the ESR (Electron Spin Resonance) principle and makes use of the Zeeman effect which splits the emission and absorption lines of atoms in an ambient magnetic field. The pattern and amount of splitting is a signature of the magnetic field strength. The optically pumped helium magnetometer uses a High Frequency (HF) discharge within a gas cell to excite 4He atoms from the ground state to the metastable state. This metastable level is split by the Earth magnetic field into 3 Zeeman sublevels. The separation of those sublevels is directly proportional to the ambient field strength and equals half the gyro frequency (eB/2m - where m is the electron mass).

Given the role of the laser in the ASM instrument, the following specifications have to be met:

• Wavelength stability with piezoelectric modulation piezoelectric modulation around the D0 transition : λ=1082.908 nm in air standard (std)

• RIN (Relative Intensity Noise) lower than -135 dB (Hz)1/2 at 1 kHz and 2 mW of output optical power

• Spectral linewidth lower than the D0 absorption width of 1.7 GHz

• Specific space environment and space design requirements.

The fiber laser consists of a pump laser diode, a WDM (Wavelength Division Multiplexer), an Yb doped FBG (Fiber Bragg Grating), an optical isolator and a splitter to allow a feedback control. In this architecture, the noise reduction loop acts on the pump diode current by detecting the low frequency fluctuations of the output optical power from the dedicated photodiode located in the 20 % splitter output. The corresponding block diagram is presented in Figure 29.


Figure 29: Block diagram of the laser architecture (image credit: CEA-LETI, CNES)

Athermal design of the LFA (Laser Fiber Assembly): The ASM laser has to be able to pump the 4He at 1082.908 nm (in air std) within the temperature range of [-5ºC; +50ºC] which is the specified qualification operating temperature range of the ASM electronics on the Swarm satellites. One of the main challenges to take up for the LFA conception was to reduce the thermal wavelength drift as much as possible in a passive way. We could have used, for example, a 1083 nm laser diode with a thermo-electric Peltier device but the laser consumption would have been highly increased.

The role of the piezoelectric actuator (made by the CEDRAT Group) is to modulate and to allow a fine tuning of the laser wavelength around the 4He D0 transition. A picture of the final LFA design is given in Figure 30.


Figure 30: Photo of the LFA (image credit: CEA-LETI, CNES)

Legend to Figure 30: The pre-stressed piezoelectric actuator on the right is glued on the Zerodur®; the LFA is then fixed in the ASM electronics box with the titanium bridge put upon the Zerodur®.

Another challenge to take up was the design of a suitable fixation system of the LFA allowing it to comply with the shocks and vibrations specifications. A solution has been developed using a titanium bridge, coned-disc springs (also known as Belleville washers) and elastomers (with suitable thickness and hardness, and a low outgassing characteristic). - The pieces of elastomer are located under the LFA and between the titanium bridge and the LFA in order to dampen any tri-axis vibrations or shocks. This fixation system has successfully passed the vibrations and shocks qualification tests. The final titanium fixation bridge is shown on the LFA of Figure 30.


Figure 31: RIN measured at 1 kHz with a SWARM flight model laser (image credit: CEA-LETI, CNES)

RIN performance: For both consumption and simplicity reasons, in the ASM instrument it has been chosen to detect the magnetic resonance signals coming from the 4He cell around 1 kHz, which corresponds to the 1 kHz modulation of the continuous part of the LA0 magnetic resonance signals between the Zeeman sub-levels. This signal is high enough to allow the measurement of the magnetic field with a low noise laser: the corresponding RIN measurement of the ASM laser is shown in Figure 31 (obtained from an electronic spectrum analyzer and an InGaAs photodiode). As shown, the low frequency feedback control loop presents the opportunity to reduce the RIN from about -105 dB (Hz)1/2 at 2 mW of output power around 1 kHz down to -140 dB (Hz)1/2. The low noise laser specifications are thus met.

Instrument mass, power consumption

3 kg, 5.3 W

Size of sensor head

40 mm x 60 mm

Size of DPU

200 mm x 150 mm x 100 mm

Dat rate

0.35 MByte/ day

Dynamic range

15000 - 65000 nT full scale

Absolute accuracy

< 0.3 nT (2σ)

Omni-directional response

< 0.1 nT angular dependence

Table 5: Specification of the ASM instrument

The instrument assembly consists of a DPU (Digital Processing Unit) and a separately installed sensor connected to the electronic box by a bundle of optical fibers and electrical cables (harness). A specific sensor bracket is designed to mechanically interface 2 identical sensors with the satellite boom (a cold redundancy has been chosen for Swarm, each sensor being connected to a dedicated DPU located within the satellite main body).


Figure 32: Photo of two ASM sensors and their sensor mount (image credit: CNES)


Figure 33: ASM DPU and ASM sensor functional architecture (image credit: CEA, LETI, CNES)


Figure 34: Illustration of the ASM DPU (image credit: CNES, CEA-LETI)

The ASM characteristics make the instrument ideally suited not only for the traditional role of scalar magnetometers as absolute references for the calibration of the on-board vector instruments, but also for extended operational capacities, such as higher frequency scalar measurements (of potential interest for magnetosphere studies for the low frequency part of the spectrum) or autonomous scalar / vector operations. Last but not least, the helium magnetometer can be operated in a zero field configuration with only very minor evolutions in the sensor overall design, thus extending its initial capabilities to new missions in planetary exploration (Ref. 72).

Electromagnetic cleanliness: Differential scalar measurements have been performed throughout the ASM development phases up to the satellite final ground magnetic tests. This allowed first to select the sensors parts in order to minimize their residual magnetic signature and second to evaluate the accuracy of the magnetic data that will be delivered by the ASMs for the three satellites of the Swarm constellation both at instrument and satellite level. This differential method has been adapted to the various rather challenging tests configurations met within the Swarm program. It allowed to demonstrate that the ASM performances meet the mission requirements, with measurement uncertainties below 25 pT. It finally contributed to improve the quality of the magnetic measurements that were carried out at satellite level and check the perturbation models that had been established. 74)


EFI (Electrical Field Instrument):

EFI, also referred to as CEFI (Canadian Electric Field Instrument), is provided by Canada (CSA funding, design by the University of Calgary, with ComDev Ltd. of Cambridge, Ontario as instrument manufacturer). The CEFI sensor is based on SII (Suprathermal Ion Imager)-a Canadian particle detector design that has already proven its capability - to gather precise measurements of ion winds. The goal of the CEFI instrument is to characterize the electric field about the Earth by measuring the plasma density, drift, and acceleration at high resolution; also for plasma density mapping in conjunction with GPS.. CEFI derives its heritage from the CPA (Cold Plasma Analyzer) instrument on Freja, the Nozomi TPA instrument and the CUSP, JOULE and GEODESIC sounding rocket missions. 75) 76) 77) 78) 79)

The plasma ion measurements are derived from energy-angle distributions that are generated by two orthogonal 2D electrostatic analyzers on each satellite. The ion bulk flow velocity and temperature are related to the distribution moments by transfer functions whose form are determined from simulations of the analyzers. The electric field is determined from measurements of the ion velocity and the magnetic field. 80)

The main sources of error come from uncertainties in the instrument transfer functions, the sensor-to-plasma potential difference, particle Poission noise, galactic cosmic ray event, and detector gain variations.

The CEFI instrument is comprised of three main parts: the SII (Suprathermal Ion ImagerI) sensors, the LP (Langmuir Probe) sensors, and the Electronics Assembly. The electronics assembly contains all of the electronics necessary to support power supply, sensor data acquisition, instrument control and communications with the spacecraft bus. The electronics assembly and SII sensors will be positioned on the ram face of each Swarm spacecraft along with the Langmuir probes positioned preferably on the ram and nadir faces of each spacecraft and connected to the electronics assembly with wire harnesses.


Figure 35: The EFI instrument measurement concept (image credit: University of Calgary)


Figure 36: Schematic of the CEFI device (image credit: University of Calgary)

The SII sensors are of CPA heritage; they are using a unique particle focusing scheme developed at the University of Calgary. Ions enter a narrow aperture slit and are then deflected by a pair of hemispherical grids that create a region having electric fields directed radially inward. Incoming low-energy positive ions are accelerated toward the center of the spherical system, whereas ions with larger kinetic energies travel farther toward the edge of the detector, creating an energy spectrum as a function of detector radius (Figure 37).

Particles arriving from out of the plane of Figure 37 land at different azimuths on the image plane. The resulting image from each SII sensor is a 2D cut through the ion distribution function, from which one can calculate ion density, drift velocity (2D), temperature, and higher-order moments. The two SII sensor head assemblies are oriented such that the aperture slits are oriented perpendicularly to each other, enabling 3D characterization of the ion distribution.

Range of operational conditions:

• Natural variability

- Ion densities (108-10 m-3)

- Ion and electron temperatures (0.1-0.5eV)

- Ionospheric plasma flow (~200 m/s)

• Active biasing of face plate

• Passive biasing of material components associated with different work functions and contact potentials.


Figure 37: Schematic view of the particle focusing system (image credit: University of Calgary)


Figure 38: Illustration of the CEFI sensor head assembly (image credit: University of Calgary)

When the charged particles strike the MCP Microchannel Plate) detector, the signal is amplified through secondary emission processes. The voltage applied across the MCP controls the gain of the device. In parts of the orbit where ion flux is high, the voltage applied to the front surface is reduced to limit the gain of the device and preserve its life. This gain adjustment is part of an automatic gain control realized through the use of a feedback loop using the CEFI instrument faceplate current as a control input. Where sufficient gain control cannot be achieved via the MCP voltage alone, an electrostatic shutter will gate the incoming ions with duty cycles ranging from 100% to well below 1%.

A Langmuir Probe assembly is part of the CEFI device to provide measurement of electron density, electron temperature and spacecraft potential. The LP design is based on hardware flown on the Cluster and Rosetta missions of ESA and was developed by the Swedish Institute of Space Physics. The overall height of the LP sensor is 10 mm.

The instrument electronics assembly includes the following electrical subsystems:

• Instrument controller

• Detector readout electronics

• HVPS (High Voltage Power Supply)

• LVPS (Low Voltage Power Supply). The primary function of the LVPS is to generate low voltages for other electronics within the Swarm CEFI instrument.

• Langmuir probe assemblies.


Figure 39: CEFI electrical block diagram (image credit: University of Calgary)


Figure 40: Engineering model of the EFI instrument with the two orthogonal sensor heads (image credit: COM DEV)


MAC-04 (Micro Accelerometer-04):

MAC-04 is an electrostatic accelerometer instrument assembly designed and developed at VZLU (Výzkumný a zkušební letecký ústav, a.s. - Aeronautical Research and Test Institute), Prague, Czech Republic. EADS Astrium (as prime contractor to ESA for the Swarm mission) awarded the MAC-04 contract to VZLU in March 2008. VZLU is the lead of a MAC-04 consortium, involving 14 Czech institutions and companies. 81)

The overall objective is to measure non-gravitational perturbing accelerations (time and spatial variability), such as air drag, winds, Earth albedo (reflected solar radiation from the cloud/snow coverage and the thermal radiation of the Earth), and solar radiation pressure on the spacecraft. In-situ air density measurements together with magnetic data can be used to obtain new insights on the geomagnetic forcing of the upper atmosphere.

Typical magnitudes of non-gravitational forces are presented in Table 6. For spacecraft altitudes below 800 km the atmospheric drag is the dominant acceleration; at higher altitudes, the direct solar radiation pressure and other radiative forces are increasing in dominance.

Acceleration origin

Acceleration magnitude (ms-2)

Atmospheric drag

10-4 to 10-9

DSRP (Differential Solar Radiation Pressure)

2.9 x 10-8


10-8 to 10-9

Reflected infrared radiation

4 x 10-9

Table 6: Expected values of non-gravitational accelerations in LEO

All these radiation effects have a common feature – namely a slow change of magnitude with time and position in orbit. However, for the solar radiative effects, the illumination of the spacecraft is a precondition for any acceleration detection in its orbital path.

The solar activity has a direct impact on atmospheric drag. The thermosphere usually expands and contracts in line with the sun's 11 year solar cycle. During solar maximum when solar activity increases, it causes the thermosphere to heat up — reaching temperatures of 1100ºC — and expand. The opposite happens during solar minimum. As solar activity increases, EUV (Extreme Ultraviolet) radiation heats our planet's gaseous envelope, causing it to swell and reach farther into space than normal. However, despite of the sun's possible rapid fluctuations, Earth's thermosphere response is a fairly slow process. These considerations led to the design performances of the MAC-04 accelerometer as shown in Table 7.

Linear acceleration range

± 10-4 m s-2

Angular acceleration range

± 3 x 10-8 rad s-2

Measurement bandwidth range

10-4 to 10-1 Hz

Resolution for linear acceleration better than

10-9 m s-2

Resolution for angular acceleration better than

10-7 rad s-2

Overall random error better than

5 x 10-10 m s-2

The accuracy of a component of the linear acceleration vector shall be better than 0.2% of the measured value

Table 7: Performance specification of the 3-axis MAC-04 accelerometer

Measurement principle:

The MAC-04 assembly is composed of a cubic proof mass which is free-floating in the cubic cavity. The center of the sensor will be accommodated in the spacecraft’s center of gravity. The proof mass is separated from external influences by the satellite structure and the construction of the micro-accelerometer.

Free motion of the proof mass is realized by virtue of the gravitational law. The cavity is rigidly connected to the satellite body. The gravitational as well as all perturbing forces acting on a satellite produce their acceleration contributions which are identical to the one of cavity. The difference between the acceleration of the cavity and the acceleration of the proof mass is the sum of all accelerations produced by the non-gravitational forces acting on a satellite.

A precise measurement of the proof mass position enables to properly detect its small relative displacements with respect to the satellite-fixed cavity. Applying a known electrostatic force, the instrument can compensate and measure the action of the non-gravitational forces. This proof mass position control is performed by a feedback loop of the servo control electronics.

A block diagram of the sensor control circuit is shown in Figure 41. Part of the block diagram placed in the dashed border represents the dynamics of the proof mass. The constants A1, A2 and A3 represent the shape and geometry of the electrodes and their distance to the proof mass.


Figure 41: Block diagram of proof mass position regulation system (image credit: VZLU)


Figure 42: Major elements of the MAC-04 (image credit: ESA)


Figure 43: The MAC-04 electronics stack (left) and the instrument (right), image credit: ESA


Figure 44: Photo of the mechanical part of the MAC04 accelerometer (image credit: VZLU)

MAC-04 science data output 1Hz rate:

- Linear acceleration: 3 vector components

- Angular acceleration: 3 vector components

- Linear positions: 3 vector components

- Angular positions: 3 vector components

- Temperatures: 8 sensor values.


Some background of Czech accelerometer instrument projects in spaceflight:

During the last two decades, the Czech Republic has developed several accelerometer instrument models within the framework of the non-gravitational accelerometer program. There is great interest among the theoreticians in satellite dynamics since our understanding of the non-gravitational effects is rather limited. In many cases, only rough phenomenological models are available. In situ measurements of the non-gravitational effects have thus great importance for checking the theoretical concepts. The MAC (Micro Accelerometer) program (formerly referred to as MACEK) is devoted to such measurements with an expected threshold up to 5 x 10-11 ms-2 .

Table 8 provides an overview of the various spaceborne missions whose payload included a MAC instrument. The obtained measurements are of different quality. The first two instrument models (flown on Resurs-F1 and on the STS-79 flight of the Space Shuttle) were mainly of in-flight technology verification nature. The MIMOSA project was the first small satellite dedicated to the research of the thermosphere. Unfortunately the failure of unlocking in one axis spoiled the mission goal.

The first MAC-04 instrument of VZLU, developed under in the frame of the project TEASER (Technological Experiment And Space Environment Resistance), was launched with the Russian Tatiana-2 mission in Sept. 2009. The project was funded Ministry of Industry and Trade of the Czech Republic. - However, after spacecraft separation from the upper stage of the launch vehicle, the satellite stabilization system did not work correctly (infrared Earth sensor failed). When all trouble shooting didn't bring any tangible results, the communication with the spacecraft was terminated after one month of “operations” with the spacecraft.

As of mid-2010, the development of the MAC-04 instrument for the Swarm mission is completed and instrument qualification for the mission is underway. Although the measurement principle is quite simple, past experiments showed high demands on the precise adjustment of the accelerometer in ground conditions. Moreover the complexity and the high precision adjustment of the sensor mechanical parts in combination with electronics properties (e.g. temperature dependence) represent a challenging task in the MAC-04 accelerometer verification.

Launch of mission, Instrument model ⇒





June 23, 1992

Resurs-F1 (proof of concept flight)




Sept. 16-26, 1996





June 30, 2003





Sept. 17, 2009










Table 8: Space missions with the MAC micro-accelerometer

Instrument model ⇒




Linear range (ms-2)

±4 x 10-4

±5 x 10-5

±2 x 10-4

Angular range (rad s-2)

±9 x 10-3

±9 x 10-3

± 9 x 10-3

Resolution for linear acceleration (ms-2)

3 x 10-10

2 x 10-10

2 x 10-10

Temperature resolution (ms-2 K-1)


3 x 10-9

3 x 10-9

Stability (ms-2 day-1)




Table 9: Basic as designed/build performances of different accelerometer models

The Czech Republic formally became ESA's 18th Member State on 12 November 2008. Already in 1996, Czech Republic signed the formal Framework Cooperation Agreement. In November 2003, however, the Czech Republic became an ESA European Cooperating State by signing the Plan for European Cooperating States Agreement (PECS Agreement) and entering the ESA PECS Program. 82)


LRR (Laser Retro Reflector):

The LRR instrument is being provided by GFZ Potsdam. The objective is spacecraft POD (Precise Orbit Determination) to cm-level accuracy by SLR (Satellite Laser Ranging).


Figure 45: Photo of the LRR unit (image credit: ESA)


Figure 46: Placement of instruments on the spacecraft (image credit: ESA) 83)

Internal Field Components

Research objective

Time range

Spatial range

Signal range

Signal at certain wavelength (wl)

(B = magnetic)

Core dynamics and geodynamo processes


300 km to global

< 65000 nT

2.35 nT @ 3000 km wl

B-field vector, attitude & position

3 months to decades

2500 km to global

±200 nT/year

0.025 nT/3 months @ 2850 km wl

Lithospheric magnetization

decades to static

300-3000 km

±25 nT

2.35 nT @ 3000 km wl
0.009 nT@ 360 km wl

B-field vector, attitude & position

3-D mantle conductivity

1.5 hrs to 11 years

300 km to global

± 200 nT

NA (modelled as conductivity)

B-field vector, attitude & position

Ocean circulation

12 hrs to 2 years

600-10,000 km

± 5 nT

0.5 nT@ 10000 km wl
0.01 nT @ 600 km wl

B-field vector, attitude & position

External Field Components

Research objective

Time range

Spatial range

Signal range

(B = magnetic)
E = electric)

Ionosphere-magnetosphere recurrent systems

0.1 s to 11 years

1 km to global

B-field:±1000 nT
E-field:±0.2 V/m

B-field, E-field, and ion drift velocity vectors, attitude and position

Ionosphere-magnetosphere recurrent systems

10 s to 3 months

10 km to global

Ion drift velocity: ± 4000 m/s

Magnetic forcing of the upper atmosphere

10 s to 2 years

20 km to global

Plasma density: 1 x 108 - 5 x 1013 m-3
Air drag: 1 x 10-5 m s-2

B-field and E-field vectors, ion and electron temperature and plasma density, acceleration, attitude and position

10 s to 3 months

200 km to global

Ion and electron temperature: 103-105 K

Table 10: Anticipated signals at 400 km (reference 7)


Star tracker assembly and OB (Optical Bench):

The STR (Star Tracker) assembly provides the attitude of the VFM. Both instruments are co-mounted in a common optical bench to ensure proper alignment for the determination of the highly accurate magnetic field components. The µASC (Micro Advanced Stellar Compass) of DTU Space is being used in the star tracker assembly. It features two fully cold/hot redundant DPU's. With full cross-strapping, each DPU can control one to four CHU's (Camera Head Unit). Mission specific baffles can be designed for optimum performance. The attitude is autonomously calculated based on all brighter stars in the FOV of the CHUs; µASC can provide 22 true solutions per second. The absolute accuracy is < 1 arcsec. The instrument mass is < 1.4 kg (3 x CHU, BFL's & DPU). The power is < 5.7 W (3 x CHU+DPU). The instrument can also support asteroid science - Near Earth Object (NEO) detection and planets triangulation.


Figure 47: Configuration of the µASC instrument for the Swarm mission (image credit: DTU Space)

OB (Optical Bench): The purpose of the OB is the transference of the attitude from the extremely precise star trackers to the magnetometer field components. The OB ensures a highly mechanical stable platform for the magnetometer and the star trackers. A exhaustive thermo mechanical design and analysis is carried out to determine and minimize any thermal gradient that could cause a shift in the relative attitude between the two systems.

The star trackers are very magnetically clean; the separation between the two instruments (STR and VFM) is about 40 cm to reduce magnetic perturbation from the STRs. The exploitation of the symmetric system has resulted in a cylindrical tube holding the VFM sensor, minimizing transversal thermal gradients. Emphasis has been on the matching of material parameters, the use of iso-static support interfaces, and a detailed analysis of the loads. The instruments are calibrated as stand alone, and once integrated in the OB, an intercalibration and system verification is carried out to determine the relative orientation between the VFM and STR and to verify that the stability is as required.


Figure 48: The Optical Bench with three star tracker cameras (yellow, only two of them are shown) and the magnetometer sensor (green), image credit: DTU Space


Figure 49: Alternate view of a Swarm spacecraft and instrument locations (image credit: ESA)



Ground segment:

The ground segment consists of the following elements: 84)

1) CDAE (Command and Data Acquisition Element), located at the Kiruna ground station, Kiruna, Sweden.

2) FOS (Flight Operations Segment), located at ESA/ESOC, Darmstadt, Germany. ESOC is in charge of monitoring and planning of satellite operations.

3) PDGS (Payload Data Ground Segment), located at ESA/ESRIN, Frascati, Italy. The main tasks of PDGS are the generation of data products from the science data, data archiving, and data distribution to the user community.


Figure 50: Overall architecture of the Swarm mission elements (image credit: ESA)



Figure 51: Ground segment architecture (image credit: ESA)


Figure 52: Swarm payload data ground segment (image credit: ESA)


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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.