Minimize Meteosat Third Generation

MTG (Meteosat Third Generation)

Overview    Spacecraft    Launch   Sensor Complement   Ground Segment   References

MTG is the next-generation European operational geostationary meteorological satellite system - a collaborative EUMETSAT/ESA program. Preparatory activities for the EUMETSAT MTG series started in late 2000 in cooperation with ESA, following the decision of the EUMETSAT Council to proceed with a Post-MSG User Consultation Process. The process is aimed at capturing the foreseeable needs of users of EUMETSAT's satellite data in the timeframe of 2015-2035. Typical development cycles of complex space systems are on the order of a decade or more. The current MSG (Meteosat Second Generation) system is expected to deliver observations and services until at least 2018.

MTG will continue services beyond this date and address future challenges in weather forecasting and other services for European citizens, such as improved air quality or UV-radiation warnings, as well as climate and atmospheric chemistry monitoring. 1) 2) 3) 4) 5) 6) 7) 8) 9)

In October 2008, an agreement was signed regulating the respective roles, responsibilities and financial commitments of the two organizations concerning future phases of the MTG program as well as the approval of the MTG payload complement:

• EUMETSAT will specify and consolidate the end-user requirements, the overall mission requirements, the space-to-ground requirements, and the ground segment requirements. In addition, EUMETSAT will be responsible for the overall mission and system engineering and ground segment design and development. Further along, it will fund the procurement of the recurring satellites, the launch services and launch and early orbit phases, and also execute commissioning and operations.

• ESA is responsible for the development and implementation of the space segment technologies and the first MTG twin satellites. ESA is funding the related cost, apart from 30% to be contributed by EUMETSAT. ESA will procure all recurring satellites as part of the EUMETSAT development and operations program.

The MTG series spacecraft and sensor complement are specified to be operated in-orbit for 20 years, compared to 15 years for MSG.

In March 2010, the MTG program took another step towards full approval with the EUMETSAT special Council, by accepting the MTG End-User Requirements Document which defines not only the deliverables to the user community, but also the duration of the operational service - at least 20 years for the imagery mission and at least 15 1/2 years for the sounding mission - the number of satellites, and the satellites in-orbit lifetime. 10)

• On February 25, 2011, the full MTG program entered into force. The consolidation of the full MTG program, as now approved, resulted from a long and intensive process through which the initial studies to define user needs for the post-MSG (Meteosat Second Generation) era led to the establishment of end user requirements for the new geostationary program, followed by the early definition phases of MTG until the end of 2007, when dedicated MTG activities started at EUMETSAT under the framework of the MTG Preparatory Program. 11)

- The SRR (System Requirements Review) for the space segment was completed in April 2011. This is being followed by preparation of the PDR (Preliminary Design Review) for late 2011.




User Consultation Process & Pre-Phase-A Studies (Phase 0)
- 2001-2003: Phase 1 - High-level user needs & priorities agreed, preparation of Pre-Phase-A studies
- 2004-2005: Phase 2 - system concept studies (Pre-Phase-A), evaluation/pre-selection of MTG missions


MTG (Phase-A): Feasibility studies of selected mission concepts. Approval process of MTG preparatory program. The Phase-A studies were successfully concluded in December 2008.


Phase-B: Detailed design activities under coordinated EUMETSAT and ESA Preparatory Programs, with approval processes for coordinated MTG development programs


Phases C/D: Development and testing of the MTG system elements

2018 onwards

Phase E: Nominal need date for MTG series. Launch, operations and utilization of MTG spacecraft

Table 1: MTG overall timeline



More than 50 leading experts in a variety of disciplines were involved, representing operational and research organizations from Europe, the United States and other international partners, as well as WMO (World Meteorological Organization). The definition of the requirements was driven by the EUMETSAT customers' long-term strategic objectives, namely the most business-critical improvements to meteorological and environmental services to be achieved in the 2015-2025 timeframe. The main customers include National Meteorological Services and other operational organizations from EUMETSAT Member States, the ECMWF (European Centre for Medium-Range Weather Forecasts) and EUMETNET. Identification of candidate observing techniques was also performed, together with the preliminary assessment of their capabilities and suitability to satisfy customer needs. 12) 13) 14)

Based on these assessments, five candidate observation missions were identified for MTG: 15)

- HRFI (High Resolution Fast Imagery) mission. For MTG, the most stringent requirement is the absolute geolocation knowledge error needed for the HRFI imaging mission: the Earth location of acquired samples needs to be determined with accuracy (knowledge) better than 250 m at SSP (Subsatellite Point) (i.e. 7 µrad for a sample distance of 0.5 km) with a 68.26 % confidence level over the HRFI coverage.

Further, the timeliness requirements are also very demanding. For the HRFI imagery level 1b data need to be delivered within 150 seconds of their acquisition. This sets constraints on the number of observations that can be accumulated and used in the image navigation processing. Therefore, efficient algorithms for the processing of image and on-board data are requested.

- FDHSI (Full Disk High Spectral Imagery) mission

- IRS (Infrared Sounding) mission. For the IRS sounder, the requirement is more relaxed, and amounts to 800 m at SSP (i.e. 22 µrad) and at 68.26% confidence over a LAC (Local Area Coverage), equivalent to a quarter of the Earth. However the sounder has a demanding pointing stability requirement of 300 m at SSP and at a 68.26% confidence level, over the dwell time of about 9 seconds.

- LI (Lightning Imagery) mission

- UVS (UV-VIS Sounding) mission.

Regarding the MTG missions, the satellite availability shall be at least 96% calculated on an annual basis for the duration of the satellite nominal operational life. From this outage, 1% is allocated to unscheduled outages (e.g. safe mode) and 3% to scheduled outages (e.g. station keeping maneuver and operations like satellite decontamination). Therefore disruptions due to orbit control maneuvers must be minimized, using appropriate station keeping strategies and satellite performances.

INR (Image Navigation and Registration) requirements: An INR system concept, generally applicable to all MTG instruments, but exemplified here for the FCI only, has been developed during the Phase A study. The concept has some affinity to the GOES-N INR system (Ref. 15).


Figure 1: INR and FD (Flight Dynamics) system architecture design (image credit: EUMETSAT)

The INR system shown in Figure 1 has been chosen for the simulation to prove the feasibility to MTG geolocation requirements. On the satellite, a combination of star tracker and gyroscope are the essential elements to determine the space bus attitude. The ST (Star Tracker) determines an absolute pointing reference while the gyroscope records relative changes in the accelerations. The star tracker uses an on-board star catalog, which is uploaded to the spacecraft (about every six months). Star tracker and gyro data are inputs to the Attitude Determination unit which calculates the necessary updates of the attitude. A control signal is sent to the reaction wheels (RW) to point the spacecraft towards the center of the Earth and correct for bus disturbances. The bandwidth of the control signal is typically significantly smaller than the bandwidth of the signals coming from the ST/gyro unit, which is providing typically gyroscope measurements of some tenth of Hertz.

The basic orbit and attitude estimation accuracy performance that can be obtained with standard tracking system and on-board attitude sensors is insufficient for meeting the image navigation requirements of MTG. The performance of this basic approach is therefore complemented by a simultaneous INR processing based on instrument observations. The derived performance estimates from simulations indicate that the combined approach is sufficient for fulfilling the MTG INR and operational requirements.


Figure 2: Orbit determination strategy with and without INR (image credit: EUMETSAT)



Space segment:

Following Phase-A studies and consolidation of MTG mission definition, a baseline for the EUMETSAT/ESA Phase-B activities and for the preparation of the EUMETSAT full MTG Program Proposal was agreed composed of the Flexible Combined Imager (FDHSI and HRFI missions), the IRS (Infrared Sounder) mission, the LI (Lightning Imager) mission and the accommodation of GMES Sentinel 4 instrument (UVS mission) to be provided by ESA.

The MTG satellite system concept is based on a twin configuration of 3-axis stabilized satellites:

MTG-I (MTG-Imaging mission satellite)

MTG-S (MTG- Sounding mission satellite)

The target date for the operational deployment of the MTG space segment element, ensuring continuity of the the imagery mission, is now the end of 2018. This requires a launch in early 2018 with a one year commissioning phase anticipated for the the new MTG system, including ground and space segments. Figure 3 shows the baseline MTG deployment strategy. This has a direct impact on the envisaged overall mission development and deployment schedule. High mission and satellite reliability and availability are required combined with the nominal and extended mission lifetime for both the MTG-I and MTG-S satellites. 16)

EUMETSAT, in coordination with ESA and industry, have established priorities vis-à-vis of the overall mission. The highest priority being for the Imaging missions (HRFI + FDHSI), which will provide continuity to the current MSG mission.

The EUMETSAT MTG program includes overall system activities, development of the ground segment, procurement of the recurrent satellites. The imagery mission will consist 4 MTG-I satellites, whereas the sounding mission will be fulfilled with 2 MTG-S satellites.

It is planned to use a common bus design for the MTG-I and MTG-S satellites with local adaptations as required per mission. It is expected that the following bus subsystems will be largely common for both satellite types:

- Structure (separate mechanical interfaces are foreseen for the two payload complements)

- Thermal (conventional thermal control design envisaged)

- Attitude and Orbit Control (similar design approach for both satellites)

- Propulsion (based on a chemical propulsion system)

- Electrical Power (being modular to adapt the power needs as required per satellite type)

- Command and Data Handling (similar design envisaged for both satellites with adaptation per satellite)

- Telemetry, Tracking and Command (heritage of existing conventional S-band systems).


Figure 3: MTG mission notional deployment scenario (image credit: EUMETSAT, ESA) 17)

Legend to Figure 3: The figure shows the anticipated overlap coverage between the MSG (Meteosat Second Generation) and MTG (Meteosat Third Generation).

Following on from MSG (Meteosat Second Generation), MTG is a cooperative venture between EUMETSAT and ESA, and will ensure continuity of high-resolution meteorological data to beyond 2037. The cooperation on meteorological missions between EUMETSAT and ESA is a success story that started with the first Meteosat satellite in 1977 and continues today with the MSG series and the polar-orbiting MetOp series.

In February 2010, ESA selected a consortium led by TAS (Thales Alenia Space) of France and Italy and OHB Technology of Germany with its subsidiary Kayser Threde (KT) to build Europe's next-generation meteorological satellites. The six MTG spacecraft - four imaging and two equipped with sounders - will operate from GEO (Geostationary Earth Orbit). However, the decision does not mean an immediate contract for the winning team. Instead, ESA will enter into final negotiations with TAS and OHB to settle open issues. 18) 19) 20)

On Feb. 24, 2012, ESA and TAS signed the MTG contract. Thales Alenia Space leads the industrial consortium that is building the MTG family. Along with being the prime contractor, TAS is responsible for the MTG-I imaging satellite, including the primary payload, the FCI (Flexible Combined Imager). Bremen-based OHB is responsible for the MTG-S satellites and the provision of the common six satellite platforms, supported by Astrium GmbH as the System Architect. The IRS (Infrared Sounder), to be flown on MTG-S, will be developed by Kayser Threde. 21)

Earth observation missions

- FCI: 10 minute FDC (Full Disk Coverage), 21 bands between 0.4 and 13.3 µm, very high INR performances
- IRS: 15 minute LAC (Local Area Coverage), 1900 spectral channels between 4.4 and 14.7 µm, high INR performances
- LI: lightning events detection with a False Alarm Rate < 350/s after ground processing


- 8.5 years in-orbit lifetime / 10.7 years for propellant
- 4% maximum outages over one year
- High reliability (0.75 at 8.5 years for FCI and IRS missions)
- High degree of operability

External interfaces

- UVN accommodation on MTG-S
- Compatibility with different launchers and injection orbit
- Ground segment interfaces


- 4 MTG-I and 2 MTG-S flight models
- Launch of first MTG-I in 2018, launch of first MTG-S in 2019
- Up to 10 years on-ground storage
- Maximized commonality between MTG-I and MTG-S

Table 2: Summary of key requirements for MTG missions 22)


Figure 4: Overview of the MTG industrial core team (image credit: ESA, EUMETSAT, Ref. 17)

MTG-I Spacecraft

• Significantly higher performances than MSG, similar to GOES R (of NASA/NOAA)

• Transition from spinner to 3 axis concept

- Solar entrance in instrument cavity

- Two axis scanning

• IR performance requires active cooling at 55 K

• Main technical challenges: attitude measurement & pointing stability, compensation of scan motion by reaction wheels, INR, LWIR
detectors, scan mechanism, active cryocooling, instrument thermal architecture (sun effect), micro vibrations & other attitude perturbations

MTG-S Spacecraft

• First Geo IR sounder of this class in the world (NOAA has stopped the GOES R sounding mission (HES)

• Many technical challenges solved on MTG-I

• IRS FTS concept derived from IASI

• Huge data rate (3 GByte/s at instrument output) / real time acquisition and processing

• Pointing stability over 10 s

• Higher resolution of imagery: 0.5 to 2 km instead of 1 to 3 km

• 16 more accurate spectral bands instead of 12

• Repeat cycle 10 minutes instead of 15 minutes

• Data rate around 60 Mbit/s instead of 3 Mbit/s

• More challenging INR requirements, combined with 3 axis stabilization introduction

Table 3: Overview of main technical challenges of the MTG series (Ref. 20)


Figure 5: Artist's rendition of the MTG-I series spacecraft (image credit: ESA)

Structure: A two-winged satellite configuration is selected, which minimizes the solar momentum build-up and therefore the reaction wheels off-loading frequency and its associated mission outage. A yaw flip is performed around each equinox, to ensure that the satellite –Y side is always protected from sun illumination, which is beneficial to the thermal control of the instruments and AOCS fine sensors.

AOCS (Attitude and Orbit Control Subsystem): The AOCS supports a "fine pointing mode" to provide high pointing stability. Attitude determination is based on high performances gyrometer and multi-head star trackers necessary to ensure an accurate and stable Earth pointing. To make the most of their performances, the AOCS fine sensors have been accommodated inside the platform towards its cold side. This configuration is indeed the only one which provides sufficient protection with respect to the varying sun illumination conditions, and minimizes therefore the attitude restitution errors at sensors level. Actuation is provided by 5 reaction wheels, enabling the avoidance of zero crossing and the related outage. The reaction wheels are off-loaded by the thrusters. The reaction wheels are also in charge of the compensation of the disturbances induced by the instruments mechanisms (FCI or IRS). Microvibrations are minimized by implementing isolating devices for the main disturbances which are the reaction wheels and the instruments cryocoolers.

EPS (Electrical Power Subsystem): The EPS architecture relies on a regulated 50 V bus with current limiter protections. Power conditioning and battery management are autonomous and based on dedicated hardware functions. A PCDU (Power Conditioning and Distribution Unit) provides all the satellite units with the necessary equipment regulated lines and thermal heater distribution.

Avionics: The MTG spacecraft are based on a central avionics architecture, with central OBSW (OnBoard Software) in charge of all satellite subsystems and instrument high level management. The SMU (Satellite Management Unit) is at the core of the MTG spacecraft, and provides the processing resources and the interfaces to all platform units and the bus management. It is also the master of spacecraft autonomy and FDIR (Failure Detection, Isolation and Recovery) functions.

The satellite OBSW is based on a modular and hierarchic architecture, providing standard "bus" interfaces to all applications. System management (modes transitions, FDIR recovery actions) is parametrized by a system database and based on On-Board Procedures or action sequences.

The on-board architecture relies on 3 main communication buses:

• 1 MIL-STD-1553 dedicated to platform units

• 1 MIL-STD-1553 for the command and control of instrumentation

• A high rate SpaceWire (SpW) network dedicated to payload data and high rate payload telemetry.

The architecture furthermore ensures the modularity and parallel development of instruments and platform modules, through the following payload related functions:

• The ICU (Instrument Control Units), which are in charge of the FCI and IRS (most complex instruments) management and data-processing. The ICU is part of each instrument.

• The PDD (Payload Data-Downlink), which is in charge of the payload data acquisition and high rate downlink in Ka-band.

• A Common MTG Packet User Standard (MTG tailoring), which is implemented in the platform and instrument software.


Figure 6: General overview of the MTG-I (top) and MTG-S spacecraft (bottom), image credit: TAS

MTG SpaceWire architecture:

The MTG satellites accommodate, respectively, the FCI imager, LI imager and the DCP digital transponder for the Imager S/C (or MTG-I), and the IRS and UVN sounders for the Sounder S/C (or MTG-S), over a payload SpaceWire network for mission data distribution and instrument's configuration with a total high rate telemetry of 295 Mbit/s (Imager S/C) and 557Mbit/s (Sounder S/C) after RS (Reed Solomon) concatenated encoding and encryption. 23) 24)

The payload data network is built around a DDU (Data Distribution Unit) that implements SpaceWire (SpW) routers for 3 instruments (FCI, LI, DCP or IRS1, IRS2, UVN) communication and one SMU (Satellite Management Unit) computer for INR auxiliary data collection and network management. The network supports full cross-strapping between each terminal (instrument's and SMU) and DDU leading to a total of 16 terminal ports based on to independent nominal and redundant DDU SpaceWire-10X routers.



Figure 7: On-board data handling architecture of the MTG-I spacecraft (image credit: TAS)


Figure 8: On-board data handling architecture of the MTG-S spacecraft (image credit: TAS)

The network architecture is identical for both imager and sounder configurations. The network is running at 200 Mbit/s on all links providing large margins. The large margin vs data distribution and the asynchronous behavior of the SpW link, allow to accommodate variable instrument's data flow according to their operational modes. For example the UVN instrument provides 40 Mbit/s in normal mode or 125 Mbit/s in commissioning mode.

The SpaceWire time code distribution is not used for payload synchronization due to the instrument heritage: a classical OBT (On-Board Time) associated to a PPS (Precise Positioning Service) pulse is broadcast through the payload Mil-Std-1553 command control bus.

Thanks to the implementation of SpW routers, the full-duplex capability of the SpW is used for command/control messages required to configure quickly the instruments without outage: large configuration tables are loaded from SMU mass memory into the instruments through SpW links; i.e. 135 Mbit of data are transferred between 2 consecutive image acquisitions.

All messages are formatted with ECSS (European Cooperation for Space Standards) PUS (Packet Utilization Standard) and distributed over the SpW network with the ECSS SpW CCSDS transfer protocol, using the user field for identifying the virtual channel for the high telemetry destination.


Figure 9: Schematic view of the DDU (Data Distribution Unit) within the on-board SpW configuration (image credit: TAS)

The MTG satellites are the first space mission in a TAS contract for implementing and using the complete SpW network capability with a full cross-strapping redundancy with SpW-10X routers and full duplex used for command/control configuration messages with large tables (Ref. 23).

RF communications: Use of S-band for TT&C functions. The S-band subsystem consists of redundant transponders and two hemispherical TT&C antennas. The payload data downlink uses the DDU (Data Distribution Unit), which gathers the data to be transmitted to ground, a transmitter and Ka-band antenna. The latter is folded at launch and then deployed once in orbit. It is mounted on a two-axis pointing mechanism so as to ensure accurate pointing of the spot beam towards the MTG ground station before and after the yaw flip, and when the satellite is repositioned on the geostationary arc.

House keeping telemetry is downloaded in real-time through both S-band and Ka-band links, whereas the mission telemetry is downlinked in Ka-band only.

Propulsion module: The propulsion module is accommodated on the platform central tube and shears webs. The MON (monopropellant) and MMH (Monomethylhydrazine) tanks are inside the tube. Further subsystems are: Helium tanks, the thrusters and LAE (Liquid Apogee Engine), the associated tubing, valves and pyros.

Astrium has developed a Unified Propulsion System (UPS), which is adapted for the MTG satellites and offers a complete pre-integrated drive system with 16 (sixteen) 10 N thrusters for orbit and attitude control, and one 400 N apogee engine, all fuelled from two propellant tanks (925 l) with hydrazine (MMH) and nitrogen tetroxide (NTO). After its release from the launch rocket, the 400 N apogee engine will propel the satellite from GTO to GEO . At this stage, the majority of the fuel, around 80% of the total, will have been expended. The remaining 20% of the fuel will serve the satellite's 16 thrusters to maintain its exact orbit path for the 13-year mission, and for any adjustments required. 25)


MTG-I series spacecraft

MTG-S series spacecraft

Spacecraft launch mass

3400 kg class

3600 kg class

Spacecraft power

2 kW, solar array of 10.7 m2

2 kW, solar array of 10.7 m2

Data rate

165 Mbit/s

260 Mbit/s

Spacecraft lifetime

In orbit lifetime 8.5 years, consumables for 10.7 years

In orbit lifetime 8.5 years, consumables for 10.7 years

Table 4: Some parameters of the MTG series spacecraft


Launch: A launch of the first MTG-I spacecraft is planned for early 2018 (Ref. 11).

Orbit: Geostationary orbit at an altitude of ~35,786 km, the nominal longitude of MTG-I-1 is 0º.



Sensor complement: (FCI, LI, IRS, UVN, GEOSAR, DCS)

The payload for MTG-I series consists of the FCI (Flexible Combined Imager), LI (Lightning Imager), DCS (Data Collection System), and search and rescue from GEO (GEOSAR). 26) 27) 28)

The payload for MTG-S series is comprised of two instruments: the IRS (Infrared Sounder), and UVN (Ultraviolet Visible Near-infrared) Sounder. The UVN sounder will be provided by ESA as part of the Copernicus program and is referred to as the "Sentinel-4 GEO component mission".Hence, a separate file, Sentinel-4, is provided on the eoPortal which describes only the UVN Sounder.


FCI (Flexible Combined Imager):

FCI is the follow-on instrument of SEVERI (Spinning Enhanced Visible and Infrared Imager) heritage flown on the MSG series missions of EUMETSAT which started with the launch of MSG-1 (Meteosat-8) on Aug. 28, 2002. Originally, there were two successor instruments defined for MTG - namely HRFI (High Resolution Fast Imagery) and FDHSI (Full Disk High Spectral Resolution Imagery) - whose requirements were eventually combined into one instrument named FCI (Flexible Combined Imager). This implied also some descoping actions of the requirements: 29) 30) 31) 32)

- Drop of the capability of targeted observations (selectable area 6º x 6º)

- Reduction of FCI channels to 16 - originally considered: 28

- Redefinition of FCI channels more in line with current MSG SEVIRI performances

- Relaxation of longest wavelength channels (13.3 µm) compatible with more mature detectors technologies - originally considered: up to 14.9 µm.

The key requirements for FCI spatial and temporal data resolutions remain unchanged, namely:

• Spatial resolution: 0.5 / 1.0 km (solar channels) and 1.0 / 2.0 km (thermal channels)

- For imagery on local scales (1/4 of full disk) the spatial resolution is 0.5 km with 4 channels at high spatial resolution 0.5 km (2 solar), and 1.0 km (2 thermal channels)

- For full disk imagery the spatial resolution is 1 km (8 solar channels) and 2 km (8 thermal channels)

• Temporal resolution: BRC (Basic Repeat Cycle) = 10 minutes for a full disk image; BRC = 2.5 minutes for local scale (Europe / North Atlantic) retrievals.

The relaxations adopted by the MMT (MTG Mission Team) result in a system:

- which is less risky

- and has a more efficient development - by still enabling to fulfil the user's needs.

The FCI instrument will outperform the SEVERI observations on cloud, aerosol, moisture and fire detection by adding new channels and by improving temporal-, spatial-, and radiometric resolution of the data.

Spectral band (µm)

Center wavelength (µm)

Spectral width (µm)

SSD (Spatial Sampling Distance), km

VIS 0.4




VIS 0.5




VIS 0.6



1.0, (0.5) #1

VIS 0.8




VIS 0.9




NIR 1.3




NIR 1.6




NIR 2.2



1.0, (0.5) #1

IR 3.8



2.0, #2 (1.0) #1

WV 6.3




WV 7.3




IR 8.7



2.0, (2.0) #2

IR 9.7 (O3)




TIR 10.5



2.0, (1.0) #1

TIR 12.3




TIR 13.3




Table 5: Overview of band specification/requirements for the FCI imager (Ref. 16)

The channels VIS 0.6, NIR 2.2, IR 3.8 and IR 10.5 are delivered in FDHSI spatial sampling and HRFI spatial sampling configurations. The spatial sampling and spectral requirements for the HRFI sampling configuration are indicated by #1. The fire application channels are marked with #2. All other channels are delivered in FDHSI sampling configuration. In total, up to 22 image colors and sampling configurations could be delivered by each image cycle, covering all the needs.

In the nominal imaging mode, the FDHSI covers the full Earth disk with a 10 minutes BRC (Baseline Repeat Cycle). RSS (Rapid Scan Services) are provided thanks to coverage equivalent to BRC/n referred to as LAC (Local Area Coverage). The LAC coverage can be variably placed anywhere over the Earth (this shall be taken into account for the scan mechanism qualification).

In particular for the FCI coverage (Figure 10), will be used for the RSS. The images will be delivered with the following considerations:

- Imagery data will now be delivered to the users at level 1c, e.g. after image rectification

- Co-registration error is specified at level 1b as knowledge; the end performances are to be determined after image rectification

- In-flight absolute and relative image geometric quality of FCI will be assessed after image rectification, so that high accuracy landmark processing can be fully applied

- The image quality process will be applied also for the IRS data, only for in-flight performance verification. However, the sounding data will be delivered to the users before rectification.


Figure 10: Example of RSS (Rapid Scanning Services) for the FCI (image credit; ESA, EUMETSAT)

Earth coverage and temporal registration:

A major characteristic of the instrument is its flexibility, allowing it to be used in several mission scenario. The nominal scenario is the FDC (Full Disk Coverage), which consists in a full Earth scanning in less than 10 minutes. Alternatively, it is possible to perform only a local area coverage (LAC scenario). For instance, by imaging only the North quarter of the Earth , the duration of the repeat cycle is then proportionally reduced to 2.5 minutes only.

The time needed to perform a complete East-West scan swath at the equator is lower than 10 seconds. This means that any two adjacent spatial samples in the image are very well temporally registered, even if they belong to two successive scan swaths.

Finally, any single point within the coverage zone can be acquired by all the 16 spectral channels in less than 0.3 seconds, thanks to a limited in-field separation of the spectral channels. This insures a high level of inter-channel temporal registration.


One of the challenging requirements of the MTG-I remains the ICRA (Inter-Channel Co-registration Accuracy). The contributors to the ICRA characterization include at instrument level:

- Detectors (spatial response uniformity of each pixel, pitch uniformity, dispersion and alignment)

- Optics (relay optics magnification, field distortion, alignment, stray light)

- Scan mechanism (actual scan rate and direction, scan rate stability)

- Integration (alignment)

- Satellite (launch stability, thermal stability, ageing, pointing stability in terms of drift and micro- vibrations, master-clock jitter).


FCI instrument: The FCI instrument is composed of the following elements Ref. 20):

• FCI-TA (Flexible Combined Imager-Telescope Assembly). TAS is prime contractor for the instrument, Kayser Threde is responsible for the procurement of the FCI-TA:

• The SSA (Spectral Separation & Detection Assembly)

• The FCI Electronics.


Figure 11: Overview of the main components of the FCI instrument (image credit: TAS)


Figure 12: Main components of the FCI-TA (image credit: Kayser Threde)

FCI optical design:

Table 6 gives the optical characteristics of the instrument, as a result of the missions needs.







Paraxial focal length (mm)






Entrance pupil diameter (mm)






IFOV: N/S x E/W (º)

0.40 x 0.40

0.40 x 0.40

0.40 x 0.74

0.40 x 0.74

0.40 x 0.74

Table 6: FCI optical paraxial characteristics

Scanning mirror: The telescope includes a double-gimballed flat mirror called M0 that scans the full Earth disk. The entrance pupil of the instrument is located on this mirror, in order to minimize its size. At nadir, the average incident angle of the optical beam on the mirror is equal to 30°. This architecture choice was the result of the need to minimize the polarization ratio induced by the reflective coatings at high incidence angles. Figure 13 shows a view of the optical bench assembly with the telescope and the scan mirror. The scan mirror is developed by REOSC, St. Pierre du Perray, France.



Figure 13: View of the optical bench assembly with the scan mirror (M0 in red) and the telescope mirrors (image credit: REOSC)

TMA telescope: After being reflected by the scan mirror, the light reaches a TMA (Three-Mirror Anastigmat) telescope. The characteristics of the mirrors are summarized in Table 7. The end-of-life WFE of the telescope, including the scan mirror M0, will be less than 80 nm rms over the whole FOV. The telescope has an off-axis angle of 1.4° and the mirror surfaces are aspheric in order to optimize image quality.






Radius of curvature (mm)





Conic constant





Optical clear aperture diameter (mm)





Table 7: Telescope mirror optical characteristics

The mirrors substrate will be Zerodur (except M0 in SiC), which exhibits a very low thermoelastic deformation. This is needed in order to avoid line of sight drift due to solar illumination. In order to reduce the mass, a light weighting will be applied on the rear face of the mirrors. All the mirrors will have a space qualified silver protected coating, which provides excellent transmission characteristics and good thermal behavior. The fixation on the optical bench will be insured thanks to Invar Mirror Fixation Devices (MFD) glued on the mirror side, as illustrated in Figure 14. These MFD will allow to filter out the deformation brought by the structure. The equipped mirrors will be developed by Thales-SESO.


Figure 14: Illustration of the M1 mirror with MFD (image credit: Thales-SESO)

The telescope exit pupil is materialized by a physical diaphragm of 40.2 mm diameter, which is the common instrument pupil for all spectral channels. Moreover, an intermediate image is formed after M2 mirror, allowing to insert a field stop. This field stop suppresses any stray-light path outside the field-of-view. After reflection on the mirror M3, a flat folding mirror at 45° named M4, is used to bend the optical beam before reaching the SSA. Figure 15 gives the optical layout of the telescope.


Figure 15: Optical layout of the FCI telescope, as seen from the top (image credit: Thales-SESO)

Baffles: Several types of baffles are implemented in the telescope structure in order to optimize the straylight rejection:

• An entrance solar baffle, whose aperture angle is limited by the need to scan the full Earth to a value about 10°. Its rejection angle is about 20°.

• An inner baffle which protects the entrance cavity from scan mirror M0 to M2, as illustrated in the Figure 16.


Figure 16: View of solar baffle (in green, including the vanes) and the telescope inner baffle (brown), image credit: Thales-SESO

Calibration mechanism: The FCI, being an absolute radiometer, it must be able to perform in-flight radiometric calibration both for solar channels (VIS and NIR) and thermal channels (IR1, IR2, IR3).

• The VIS and NIR calibration is performed by imaging the Sun through a MND (Metallic Neutral Density) inserted in the exit pupil of the instrument. Any spatial inhomogeneity of the density will then be completely averaged out in the focal plane. This component will be made out of fused silica, a material known for being insensitive to radiation, so that its transmittance will remain constant until the instrument end-of-life.

• The IR channel calibration is performed thanks to an embedded flat black-body inserted close to the intermediate image, in order to reduce its size and obtain the best thermal homogeneity. The black body is slightly out of focus to average out any spatial defect.

Both the black-body and MND are carried on a calibration wheel mounted on the optical bench. The mechanism wheel inserts those components in the optical path during calibration mode, as illustrated below. These calibration are performed without interruption the image acquisitions.


Figure 17: Illustration of the two calibration configuration of the mechanism (VNIR and IR), and blackbody breadboard (image credit: Thales-SESO)

SSA (Spectral Separation Assembly): The main function of SSA is to separate the optical focused beam coming from the telescope into 5 spectral groups.

• VIS : without changing the focal length of the telescope.

• NIR, IR1, IR2, and IR3 : by collimating the 4 output beams, and adapting the magnification ratio in order to achieve the required instrument focal length for each channel.

For that purpose, the SSA will be composed of the following elements, mounted in a very stable Titanium housing:

• Dichroic beam-splitters, whose function is to separate the spectral groups

• Folding mirrors in order to geometrically separate the 4 collimated NIR/IR optical beams at SSA output

• And lenses in order to collimate the beams. This collimation allows to relax the integration tolerances between the SSA and the cold optics.

The SSA will exhibit an end-of-life WFE in the range of 100 - 200 nm rms depending on the spectral group. The SSA will be mounted on the OBA optical bench, and is illustrated in Figure 18. The SSA is developed by REOSC.


Figure 18: Illustration of the SSA housing (image credit: REOSC)

Cold optics: After being spectrally separated and collimated, the NIR/IR optical beams reach the cold optics, which are located inside a common cryostat. When entering the cryostat, the beams first go through an anti-contamination window (ACW) that prevents external molecular contaminants from condensing on cold parts. In order to avoid vignetting, the exit pupil of the SSA is coincident with cold diaphragms on top of the cold optics housings. These cold diaphragms limit the thermal IR background flux received by the detectors. There are 4 cold optics housings, one for each spectral group, which are made out of Titanium alloy. The lenses are in classical IR materials such as mono-crystalline Germanium, fused silica, ZnS and ZnSe. The cold optics will be operating at a temperature about 80 K with their full performances. The cold optics are developed by Thales-SESO.

The integration principle will be quite straight-forward, thanks to tight manufacturing tolerance, and an accurate positioning by pinning both the cold optics and the detectors on the cold plate, in order to ensure a good registration.


Figure 19: CO-I housings inside the cryostat, with baffles and anti-contamination windows on top >(image credit: Thales-SESO)

Detector package: Finally, before reaching the detectors, the optical beam goes through spectral filters. There are two kinds of detectors :

• The VIS detector, operating at ambient, is mounted directly on the SSA housing

• The NIR/IR detectors, operating at 60 K for radiometric performances, are mounted inside the common cryostat.

Those filters use the state-of-the-art technologies in terms of dielectric coatings. Blocking coatings on the rear face of the filters will allow to achieve the stringent out-of-band rejection requirement of 1%. The NIR/IR filters have been manufactured by REOSC, and the VIS filters by Jena-Optronik.


Figure 20: View of the detectors integrated on the cold plate, with strip spectral filters on top (image credit: REOSC)

The FCI instrument accommodation parameters are : (Ref. 28)

• Volume: 1.57 m x 1.72 m x 2.2 m

• Mass: < 394 kg

• Power: < 495 W (max)

• Data rate: < 68 Mbit/s.


LI (Lightning Imager):

The overall objective of the LI system is to deliver on a continuous basis information on total lightning over the full disk (high timeliness, data quality homogeneity in time and space), allowing to extend "locally developed" algorithms for NWC (National Weather Center) severe weather warning to be applied over wider areas like Europe or the full Earth Disk. - LI has no heritage in Europe. Two USA LEO missions (LIS and OTD) have already flown and one, the GML (Global Lightning Mapper) of GOES-R, is scheduled for launch in November 2016. 33)

• The LI measurements of total lightning [IC (Intra Cloud)+CG (Cloud Ground)] are complementing the global measurements of CG lightning as provided by ground based systems and will improve the quality of information which is essential for air traffic routing and safety.

• The information on IC+CG will allow to assess the impact of climate change on thunderstorm activity by monitoring and long-term analyzing lightning characteristics – in cooperation with the two NOAA GLMs on GOES-R and GOES-S - a major part of the globe is covered by a long term committed GEO lightning (IC+GC) observing system.

• Providing IC+CG information on a global scale will be a prerequisite for studying and monitoring the physical and chemical processes in the atmosphere regarding NOx, playing a key role in the ozone conversion process and acid rain generation.

• Error characterized IC+CG information can be assimilated to improve very short range forecasts of severe convective events or used to verify/validate other satellite data based NWC algorithms to forecast time and location of initiation of lightning.

The LI system is conceived for the detection of lightning in Earth's atmosphere from a geostationary orbit, where the FOV (Field of View) of Earth's full disk extents to an angle of 17.5º. The basic LI observation concept is to cover this FOV with four lenses and four detectors, each one covering a square FOV of 8.7º x 8.7º, corresponding to 12.3º diagonal FOV, in order to reduce the incidence angle on the interferential filter placed on the pupil of each subsystem. An important advantage in splitting the interference filter in four parts is related to the possibility to reduce its dimensions and to ease its manufacturing. 34) 35) 36)

TAS (Thales Alenia Space) as the MTG prime contractor is responsible for the procurement of the LI (Lightning Imager) instrument developed and manufactured by Selex Galileo of Campi Bisenzio, Italy, a Finmeccanica company. 37) 38) 39)

FOV (Field of View)

16º Ø shifted northward or 84% of visible Earth disk, including all EUMETSAT member states

Spatial sampling

< 10 km @ latitude 45º and subsatellite longitude

Dynamic range of Earth background (Lbkg)

0 - 296.5 W/m2/µm/sr (night ÷ summer solstice at midday)

Optical pulse dynamic range (LLp)

6.7 - 670 mW/m2/sr

Optical pulse spectral range

777.4 ± 0.17 nm

Minimum optical pulse duration

0.6 ms

Optical pulse size

10 km - 100 km circular pulse diameter

Maximum number of optical pulses in the FOV

25 in 1 millisecond
800 in 1 second

IADP (Instrument Average detection probability)

90% for latitude 45º, 70% as average over the FOV, 40% over EUMETSAT member states (goal)

LI mass (total optical head and electronics box)

93 kg

LI optical head envelope

718 mm x 1200 mm x 1456 mm

Table 8: Main requirements of the LI instrument

The instrument works in a staring mode, detecting lightning events within its FOV of its 4 cameras. The lightning detection is achieved implementing the following functions:

• Earth image acquisition for continuous monitoring of the lightning's presence in the FOV; the image spectral bandwidth is reduced by an optical filter to the range of the lightning spectral pulse (Figure 21)

• Calculation of pixel by pixel adaptive background to cope with non-uniformities and low terms variations of the image (oceans, clouds, area in night conditions and areas with daylight conditions) and to reject at the same time noise effects and spurious events

• Removal of the background level from the overall pixel signal to obtain the net lightning illumination level

• Use of adaptive threshold; lower thresholds can be used in low noise dark areas of the scene, using higher thresholds only in highly illuminated areas (with corresponding higher shot noise)

• Pixels for which the difference between the pixel value and the estimated background signal exceeds the threshold are kept as DTs (Detected Transients)

• Collection of the DT video data and additional information for the ground processing with a dedicated processing electronics

• In flight processing of DTs to reduce the number of FT (False Transients) to a level compatible with the platform downlink data rate constraints (30 Mbit/s).

In addition LI is capable to acquire, process and transmit to the ground an Earth background image.


Figure 21: Optical emission from lightning (image credit: Selex Galileo)

Instrument overview: LI is composed of one LOH (LI Optical Head) and one electronics unit, the LME (LI Main Electronics). The LOH consists of four identical OCs (Optical Channels), each one including (Figure 22):

• a protective cover on the baffle aperture to prevent baffle and optics contamination during launch and prelaunch activities

• a baffle for stray light suppression and thermal load minimization

• a SRF (Solar Rejection Filter), to minimize both the background level and the thermal load inside the OC

• a NBF (Narrow Band Filter) to reduce the bandwidth in the range of the lightning spectral pulse (Figure 21)

• an optical system with F# 1.73, 110 mm entrance pupil diameter (determined by radiometry required to achieve the IADP performance) and 190 mm effective focal length [determined by the targeted GSD of 4.5 km at SSP (Sub Satellite Point) and the size of detector pixels]

• a CMOS detector with 1000 x 1170 pixels, 24 µm pitch, 1000 frame/s 40)

• a processing electronics device implementing the detection functions.


Figure 22: Illustration of the LOH (LI Optical Head), image credit: Selex Galileo

Each OC images a different portion of the visible Earth surface with the four line of sights tilted 4.75° from the SSP toward North, South, West and East in order to achieve the required coverage (Figure 23).


Figure 23: FOV of OCs (Optical Channels), image credit: Selex Galileo

The LME (LI Main Electronics) performs the overall payload functions, the interface to the platform, the configuration of the processing electronics, the data flow regulation, and finally compacts and packetizes the scientific data. The LME consists of the Power Units and the Single Board Computer. This first European processing board, implementing a "Power PC" technology (PPC7448 with 2300 DMIPS @ 1GHz), allows managing the Application Software that implements the algorithm of the microvibration filter (to reduce the number of FTs due to microvibrations) and the algorithm for data regulation implemented to prevent the instrument from saturation of internal interfaces and memory buffers.


Optical Design and Spectral Filters

The optical system of each Optical Channel is composed of 5 lenses and two spectral filters: a SRW (Solar Rejection Window) and a NBF (Narrow Band Filter). The SRW is the first optical element, placed close to the baffle. It is devoted to reflect the solar radiation and to work in synergy with NBF to obtain the required spectral filtering.

The NBF performs the spectral discrimination of the lightning pulse from the Earth background radiance. Since the spectral response of an interferential filter is a function of the AOI (and the filter equivalent refractive index), the NBF has to be placed where, in the optical system, the beam slant is smaller. The LOH multi-channel approach reduces the FOV of each optical system allowing an advantageous placing of the NBF in front of the lens so that the beams are parallel and the filter AOI (Angle of Incidence) coincides with the FOV.

The driver of the optical design are the narrow wavelength range and the 90% ensquared energy requirement on a 25°C thermal range, guaranteed by the OTC (Operational Thermal Control) system.


Figure 24: Schematic of the optical layout (image credit: Selex Galileo)

Effective focal length

190.8 mm

Entrance pupil diameter

110 mm

Entrance pupil position

on the NBF filter

FOV (Field of View)


Transparency (including filters)


Wavelength range

777.15 - 777.75 nm

Max vignetting


Thermal operating range

15 to 40ºC

Maximum distortion


Table 9: Summary of the LI optical system parameters


Optical Head thermo-mechanical design

The skeleton of the OH consists of a Structure made of composite material. In its lower part the structure is conceived as a baseplate with monolithic sandwich structure, supporting the four telescopes which compose the optical system of the LI instrument and interfacing the satellite Earth deck by means of three bipods, resulting in a quasi-isostatic mounting. This solution guarantees:

• Compromise between mass and stiffness-strength (low mass/stiffness-strength ratio)

• Thermo-elastic compatibility, matching the CTEs (Coefficient of Thermal Expansion) with the underneath satellite structure, aiming to the maximum reduction of mutually induced thermal stresses and then distortions.

In its upper part the structure hosts the aluminum plate where the four baffles are linked to; this includes also the SRW for thermal reasons.

In its bottom side the Structure brings the thermal system, composed by the radiators and heat pipes for the thermal control of the instrument, such to guarantee the proper heat rejection for internally generated heat, and the maintaining of the required temperature levels during mission phases. The entire LOH (LI Optical Head) is enclosed in a MLI (Multi-Layered Insulation) blanket, providing thermal shielding towards the external irradiative environment.

The optical barrels consist of cylindrically shaped housing structures in Titanium alloy, inside which the optical elements (lenses, NBFs) are mounted.

The FPAs (Focal Plane Arrays), containing the DUs, are mounted on the bottom of the barrels by means of an isolating system: this will also allow enhancing thermal stability of the optical system when temperature isothermally changes.

The FEE (Front End Electronics) is accommodated close to the FPA, for signal integrity reasons. The functional connection between FPA and FEE is provided by PCB Flex elements.

The LOH thermal design is devoted to the maintenance of the minimum temperatures for all items during the instrument Safe and Survival mode and to the fulfilment of the performance operative temperature range during the nominal Operational mode. The thermal design has to cope with the dissipation of four FPAs and four FEEs for a total of about 110W, in addition to orbital heat fluxes directly incoming in each baffle and SRW and external MLI. The heat transport toward the radiator is obtained by means of a cold plate and a redundant heat pipe.

Two cavities can be identified in the design, dedicated to thermal and thermo-elastic stabilization of the relevant hardware: one FPA cavity and one optical channels cavity.

In addition to passive elements, the thermal design includes active thermal hardware for OTC and STC. The OTC is devoted to the temperature stabilization of LOH internal items in order to meet the functional and pointing stability requirements and is based on 13 heater lines individually controlled by means of thermistors and drivers located inside LME PU (Power Unit). The STC (Survival Thermal Control) is composed by 8 heater lines individually controlled by means of thermistors and electronic thermostats located inside STC box in the LOH, directly powered by the platform when LI and OTC are switched OFF.

DU (Detector Unit)

The DU is a key element of the LI design as many of the system requirements are dependent on its characteristics and performance. It is a backside illuminated CMOS APS (Active Pixel Sensor) detector with the architecture of Figure 25 and the characteristics reported in Table 9. The backside illuminated technology improves the QE (Quantum Efficiency) and consequently the SNR required by the detection process.

The DU has internal column ADCs (Analog Digital Converters) allowing digitization of 5 rows in parallel, providing the necessary processing capacity to operate at 1 kHz frame rate. Digital data are then multiplexed internally to the serial LVDS (Low Voltage Differential Signaling) outputs running at 250 Mbit/s.

Format, pixel

1000 x1170, 24 µm

Digital outputs

60 @ 250 Mbit/s

ADC (Analog Digital Conversion) resolution

12 bit

Full well

450.000 e-

Global QE

> 0.7

MTF @ Nyquist

> 0.55

Frame rate

1 kHz

Noise in darkness

< 150 e-

PRNU (Photo Response Non-Uniformity)

< 3%

Dark current

< 66.000 e-/s

Nominal temperature


Table 10: Parameters of the DU


Figure 25: Detector architecture (image credit: Selex Galileo)


On-board Processing

The on-board processing detection architecture is composed by two three main blocks: RTPP (Real Time Pixel Processor), SDTF (Single DT Filter) stage, and MVF (Micro-Vibrations Filter) stage. The RTPP interfaces the digital outputs of the DU, reading the pixel outputs and identifying DTs (Detected Transients).


Figure 26: On-board processing scheme (image credit: Selex Galileo)

The lighting event elementary detection process mainly consists of the following tasks, based on time-domain filtering, performed on each pixel independently by the RTPP stage:

- reference background estimation for thresholding

- subtraction of estimated background from current pixel signal

- threshold calculation/optimization, based on estimated background noise

- thresholding of differential signal

- encoding of triggered event position within the DU array

- triggered event radiometric data collection.

The RTPP threshold selected for each pixel is adapted to the current estimated background level of the pixel itself. At each frame (typically 1ms frame rate), the radiometric data of the 3 x 3 window surrounding the pixels which are revealed as Detected Transients are collected and sent to the next stage of on-board processing.

The SDTF is a filter implemented in order to reduce the FT rate per frame, mainly generated by temporal noise affecting the background acquisition. It is based on the integration of the signal in the 3 x 3 pixels window surrounding the DT, where, in case of false detection of the central pixel, neighbor pixels do not bring information (uncorrelated noise).

In case a DT is confirmed by the SDTF, its relevant 3 x 3 window is then passed into the MVF stage, which is a filter designed in order to reduce the number of FTs due to microvibrations. The radiometric characteristics of the background of the 3 x 3 pixels surrounding the DT allows to implement a classification algorithm, which applies an adaptive threshold based on the estimated tuples of background radiance/gradients.

If the latter threshold is passed, the DT is confirmed for the pixel and the 3x3 window data is transmitted to Platform and then to ground, where additional filtering stages are applied in the L1b processor for further classifying its possible FT nature.

Both RTPP and SDTF stages are implemented in the FEE of each OC (Optical Channel).

Due to the high amount of pixels (each DU array is 1170 x 1000 pixels) and to the high frame rate, 60 independent RTPPs are implemented in 4 ASICs to work in parallel over each DU. Moreover, due to the high amount of data generated by the optimal threshold settings (needed to maximize the detection efficiency in a very noisy environment), four SDTF filters are implemented in parallel in an FPGA, managing the data coming from each ASIC. The implementation of processing in ASICs/FPGA, with relevant input/output buffers, internal RAM and high speed bus, allows to strongly reduce the bottlenecks in the FEE, which provides a very high data throughput capability.

The MVF and Data Regulation algorithms are implemented in the SW of the LME.



The LI does not have an in-flight calibration unit, thus the temporally continuous acquisitions performed by the instrument will be exploited, together with vicarious calibration techniques and dark acquisitions, in order to monitor the instrument through time, detect degradation of its components to some degree and to be able to perform in-flight calibration, within the required accuracy of 10% for both background and lightning pulse radiance samples.

A set of radiometric calibration parameters , KDPs (Key Data Parameters), has been identified and introduced in the on-ground L1b Radiometric Processor, in order to retrieve the spectral radiance of background samples and the radiance of the lightning pulse samples. All the KDPs, together with a set of characterization parameters, will be calibrated during on-ground phase for each Optical Channel. Only of a subset of them will be updated in-flight, according to Table 11.


LI mode


Dark acquisitions (offset and readout noise, dark signal, linearity)



Vicarious calibration



TOA reflectance database (cloud mask)



Overlapping zones (DT, radiance)



Long term pixel DT statistics



Long term pixel temporal average (pixel gain)



Long term spatial average (common gain)



Moon observation (MTF)



Table 11: In-flight calibration and characterization methods


IRS (Infrared Sounder):

The IRS (also referred to as GeoSounder) objectives are to provide "break through" measurements on the time evolution of horizontal and vertical water vapor structures ins the atmosphere – an unprecedented source of information available for the operational services in NWCs (National Weather Centers) and regional/global NWP (Numerical Weather Prediction). The IRS has no heritage instrument in previous Meteosat missions.

• The IRS (30 minute repeat cycle over Europe) will fill large spatial and temporal voids in the 12 hour time standard radiosonde observations and will allow time and space interpolation of moisture/temperature observations taken from the polar orbit.

• The IRS derived information on low tropospheric moisture and its changes in time is expected to lead to a better depiction of the hydrological cycle in models, potentially providing better precipitation forecasts.

• The IRS will provide information on vertically resolved atmospheric motion vectors with improved height assignment, which in particular is beneficial for the tropical areas having only a weak coupling between the dynamic and thermodynamic atmospheric fields.

• The IRS will provide information to identify pre-convective situations supporting NWC applications to forecast convective initiation.

• IRS will support forecasting pollution and monitoring of atmospheric minor constituents through its capability to provide estimates of diurnal variations of tropospheric contributions of atmospheric trace gases such as O3 and CO.


Frequency range (cm-1)

Main constituents

Resolving power (@ center frequency)

NEDT @ 280 K

GSD (Ground Sampling Distance), km











0.2 K




Surface, clouds


0.24 K






0.2 K




Source, clouds


0.3 K




H2O, N2O, CH4


0.2 K






0.2 K




CO2, N2O


0.2 K










Surface, clouds, N2O




Table 12: IRS atmospheric sounder requirements (Ref. 7)

The IRS operates at millimeter wavelengths which are capable to penetrate clouds and rain. Its design relies on a principle called interferometry, with separate signals from multiple antennas correlated together to produce an image of otherwise impossible sharpness. To further reduce the number of antennas needed, the device rotates at 1º/s, filling in further detail.

Geostationary infrared sounding missions offer good temporal coverage, however due to the large distance to the observed targets on Earth, the effect of diffraction is increased compared to sounding from LEO (Low Earth Orbit). Due to the wavelength dependence of diffraction, the spectral channels do not sample the same volume of air, as in general assumed by the retrieval algorithms for LEO infrared sounder data. This additional error introduced in the retrieval by diffraction limited instruments is in general referred to as ‘pseudo noise.' 41) 42)

Contrary to infrared sounders in LEO, the main objective of the IRS (in GEO) is not absolute temperature/humidity sounding. Rather, the prime objective of IRS is to support the dynamics via tracking of vertical water vapor structures. The user of IRS data is mostly interested in the information on vertical structures (temperature, humidity and wind at high horizontal, vertical and temporal resolution), which comes from the fidelity of the spectral information. This fidelity would be destroyed by excessive and spectrally uncorrelated random noise. In case of a spectrally high correlated noise (spectral bias) the information on vertical structure is not equally destroyed, i.e. it will have a lesser impact.

The IRS data will contribute both through assimilation into convective-scale, regional and global NWP (Numerical Weather Prediction) models and through nowcasting products. The data will be particularly important for observing the advection and convergence of low-level moisture associated with some types of severe weather in Europe. The main goal is to document the added value of water vapor observations derived from a hyperspectral infrared sounding instrument on a geostationary satellite for regional forecasting.

Preliminary IRS instrument accommodation parameters are (Ref. 28) :

• Volume: 1.44 m x 1.30 m x 1.25 m

• Mass: ~438 kg

• Power: ~858 W

• Data rate: ~167 Mbit/s

The state-of-the-art IRS instrument is being developed by Kayser Threde of Munich, an OHB company. The objective of IRS is to probe the atmosphere and provide information on the horizontal, vertical and temporal resolution of water vapor and temperature distributions. 43)

The IRS instrument is designed to detect, with a high radiometric accuracy, the signals emitted from gases in the atmosphere. A demanding spectral resolution for velocity determination is used to determine the wind profiles at various heights above ground together with a high spatial and temporal resolution.

The instrument will be able to scan the full Earth circle within 1 hour with a spatial on ground resolution of 4 km x 4 km from geostationary orbitin two spectral bands (MWIR: 1600 to 2175cm-1 and LWIR: 700 to 1210cm-1) with a spectral resolution of 0.6 cm-1. This resolution can be achieved by means of a high resolution telescope operating in the infrared spectral range with a scan mirror assembly allowing a step and stare of the line of sight. Radiometric, spectral and geometric requirements are met both in nominal and restricted operations conditions. 44)

Among the critical technologies and processes, the IRS detection chain shall offer outstanding characteristics in terms of radiometric performance like SNR (Signal to Noise Ratio), dynamic range and linearity. Selected detectors are HgCdTe 2D arrays, cooled at 55 K, hybridized on snapshot silicon read-out circuit at 160 x 160 format. The video electronics provide a resolution of 16 bits, and the whole detection chain (detectors and electronics) permits to reach SNR between 2 000 and 10 000 as requested by the application.

The IRS instrument is a sounding Fourier transform spectrometer, which allows recomposing atmospheric spectrum after infrared photons detection. Transmission spectrums will be used to support numerical weather prediction. The IRS comprises the following main subsystems:

• The scan assembly allows performing a complete scan of the earth disk in a fully programmable scan pattern including stares to cold space for fast and repeated recalibration of the full optical path.

• The unique design of the entrance cavity allows the sun to enter the instrument through an entrance baffle together with an internal baffle thermally decoupled from sensitive elements like mirrors and optical bench.

• The front-telescope reduces the entrance pupil, located at the scan mirror to the back-telescope, by a factor of 4, while an in-built imager allows, via a small band beam splitter, a direct and parallel detection of the investigated ground area in the spectral band between visible and infrared.

• The Michelson interferometer generates the interferograms at each position of the scan assembly by highly accurate motion of a corner cube. At each observation position 2 x 25,600 interferograms are recorded simultaneously in two spectral bands, by two detector arrays. The detectors are integrated and co-aligned in order to achieve a high inter-band coregistration.

• The back-telescope is an in-size reduced afocal front-telescope to adapt the interferometer to the cold optics. Efficient straylight reduction is achieved by means of an intermediate field stop located in the front telescope before the interferometer. The cooling of the detectors and optics behind the back telescope, down to 55 K, is realized by means of two active cryocoolers for redundancy.


Figure 27: Illustration of the IRS instrument (preliminary design of Kayser Threde)

Detection chain: The detection chain uses two MCT (Mercury Cadmium Telluride) staring arrays hybridized on silicon ROIC (Read-Out Integrated Circuit), linked to the Video Chain Unit, includes driver boards and ADCs (Analog to Digital Converters). The following items are considered as performance drivers at space instrument level: (Ref. 44)

- Spectral bands of interest: MWIR: 1600 to 2250 cm-1 (4.44- 6.25 µm) and LWIR: 680 to 1210 cm-1 (8.26- 14.70 µm).

- Full disk coverage in one hour

- On-ground resolution of 4 km x 4 km

- Radiometric measurement ranges between 180 K and 313 K (equivalent black-body temperature)

- Spectral radiometric noise (excluding spectral calibration) at 280 K blackbody: between 170 and 900 mK depending on the considered wave-number inside the band of interest.

Considering these performance drivers at instrument level, the following design choices and associated performance budget have been flown-down and proposed for the detection chain:

- Two 160 x 160 staring arrays at 90 µm pitch, cooled at 55 K

- 16 bit resolution video chain, operated at 2 Gbit/s

- Number of useful loads (or electrons) to be stored: 30 Mega-electrons in MWIR, 440 Mega-electrons in LWIR, in the useful bandwidth, considering 395 µs integration period

- SNR flux of 280 K at instrument input: 8800 for LWIR band

- Residual non-linearity after correction: 0.05%.

These figures are very challenging and cannot rely on existing space heritage in Europe. Hence, a representative detection chain breadboard has been developed and tested to secure the design choices and these high-end performance budgets.

IRS detection chain receives the optical signal from cold optical chain and delivers digitized information to the processing chain of the instrument. The detection chain main subassemblies are the IR detectors, developed and validated by Sofradir and the VCU (Video Chain Unit), developed and validated by Thales Alenia Space - Espana.

The MWIR and LWIR detectors are hybridized on dedicated CMOS silicon ROIC (Read Out Integrated Circuit). This ROIC works in snapshot integration mode, the 'Integration While Read' function is implemented to optimize the frame rate. It takes benefit of a high performance mixed CMOS technology, proposing 5 V as analog voltage, and the use of Metal-Inter-Metal capacitance as analog devices. Thanks to these technological parameters, the maximum handling capacitance requested by the application in LWIR band can be met. The pixel integration capacitances are 2.7 pF (MWIR) and 31.3 pF (LWIR) for ~2.3 V of useful voltage swing.

The focal plane topology is using a 160 x 160 pixel structure at 90 µm pitch, each pixel being composed of 9 sub-pixels of 30 µm pitch with independent integration capacitance. Each sub-pixel can be integrated and read-out independently (imager or high resolution mode) or can be read by averaging all or a part of the 9 sub-pixels (normal or interferogram mode). Acquisitions in high resolution mode at 480 x 480 format will be used by the system for mission purposes, but also to determine defective sub-pixels location. Each sub-pixel of each super-pixel can be deselected independently in order to increase the performance of the super-pixel and to minimize the number of defective super pixels. — The detector ROIC presents 16 analog video outputs. The array is divided into 16 sub-arrays of 160 lines per 10 columns, which leads to have 1600 super-pixels per outputs. Each output can be operated in parallel at 4 Mpixel/s.

The VCU contains two FEEs (Front End Electronics) located close to the detector cryostat and one VAE (Video Acquisition Electronics) device, located on the platform. The FEEs integrate the following functions: first stage of analog amplification, detector signal conversion from pseudo-differential to full differential, detector management items and very low noise detector analog voltage reference for detector biasing. Those biases are able to supply up to 50 mA peak current with a white noise about 15 µV rms. The VAE integrates the following functions: second stage amplification, the 16 bit – 4 MHz analog to digital conversion using VASP (Value Added Service Provider) ASIC. This VASP has been developed by TAS under ESA contract on HIVAC / VASP digitizer, confirming a better performance than the state of the art devices. Detector images are continuously acquired by the VCU with a fixed rate of 2.4 kHz, managing 32 active 16 bit acquisition chains in parallel for both MWIR and LWIR synchronized bands.


CCM (Corner Cube Mechanism)

The CCM of the IRS is a critical subsystem of the interferometer. The primary function of the CCM is to linearly displace one of two corner cubes in order to create an OPD (Optical Path Difference) between both arms of the interferometer. The CCM is based on a design developed for the IASI (Infrared Atmospheric Sounding Interferometers) on the MetOp satellites. As in the IASI MetOp project, Ruag Space Systems Nyon (formerly Mecanex) is responsible for the detailed mechanical design, manufacture and assembly of the mechanical elements of the CCM in close collaboration with CSEM (Centre Suisse d'Electronique et de Microtechnique), Neuchâtel, Switzerland. 45)

The CCM assembly is composed of:

• A linear scanning mechanism including in particular, a flexure guide translation stage, a voice-coil actuator and a LLD (Launch Locking Device) as shown in Figure 28.

• An electrical unit, the ICE (Interferometer Control Electronics), under Syderal-CH responsibility, which allows to control the CCM.

• A control law algorithm embedded in the Interferometer Control Electronics.

The high precision mechanism is designed to achieve low lateral shift deviations, very long life, limited exported forces to the instrument and a launch lock device that generates no shocks. Flexure technology is used throughout the mechanism to achieve the demanding requirements. The critical specification driving the project is the exported forces and speed stability which are minimized by optimizing the control law.


Figure 28: CAD model of Corner Cube Mechanism shown with mirror (image credit: CSEM)

The basic concept of the mechanism has not changed from the IASI application and relies heavily on the flexural structure technology developed by CSEM. In this way, high-accuracy linear guiding is possible without need for sliding surfaces or bearings. Compensated parallelogram flexures are used for the main mechanism constituted by two blade membranes machined in high-strength copper-beryllium alloy. An innovative aspect is given by the driving lever with flexural blades that constrain the motion of the parallelogram in order to dramatically decrease the remaining lateral error caused by the unavoidable manufacturing and assembly tolerances as well as operational dynamics.

The control law being developed by CSEM is based on a state space controller, with position speed and integral of position error as states. A feed-forward is added to improve system time response. Since the CCM speed is low, the control bandwidth is limited to 5.7 Hz with high order modes notched. The controller exhibits robustness with respect to electromechanical parameters.


UVN (Ultra-Violet Near-infrared) Sounder:

The UVN sounder on MTG-S is the GEO component of the joint Copernicus Sentinel-4 (GEO) and Sentinel-5 (LEO) concept for climate protocol monitoring and air quality applications expected to deliver data products on ozone, nitrogen dioxide, sulphur dioxide, formaldehyde, aerosol optical depth, and aerosol scattering height.

A separate file, Copernicus: Sentinel-4 (or simply Sentinel-4), is provided on the eoPortal which describes only the UVN Sounder, a hosted payload on the MTG-S spacecraft series.


GEOSAR (Geostationary Search & Rescue) service:

GEOSAR is part of the COSPAS-SARSAT international system; the objective is to provide distress alert and location information to appropriate rescue authorities for maritime, aviation and land users in distress.


DCS (Data Collection System):

The DCS mission involves, as a continuity of the MSG mission, the collection and transmission of observations and data from the ground segment consisting of surface, buoy, ship, balloon or airborne DCP (Data Collection Platforms).



Ground segment:

The ground segment contains the main ground elements necessary to support the mission. They are logically decomposed in facilities as follows: 46)

• GSTF (Ground Station Facilities)

• MOF (Mission Operations Facility)

• IDPF (Instrument Data Processing Facility)

• MPF (Multi-Programme Facilities)

• Infrastructure Facilities and Supporting Facilities

• L2PF (Level 2 Processing Facility), as part of the Application Ground Processing System

• SAF (Satellite Application Facilities).

All system operations will be conducted from the EUMETSAT HQs in Darmstadt, except for the LEOP (Launch and Early Orbit Phase) which will be performed by a LEOP service provider, to be selected. Following the spacecraft hand-over from LEOP, EUMETSAT will then execute the commissioning of the system before entering the routine operations phase. The MTG in-orbit constellation will grow progressively until it will reach its FOC (Full Operational Capability), which consists of 3 operational in orbit satellites (two MTG- I and one MTG-S).

The operational (prime) MTG-I satellite, nominally at 0º longitude, will be set to fulfill the requirements of the FDHSI (Full Disc High Spectral resolution Imagery) mission, with an FCI operating in Full Disk Scanning mode with 10 min repeat cycle. It will also be the nominal operational spacecraft for the other payloads it carries, the LI, the DCP and GEOSAR payloads. The second in-orbit operational MTG-I satellite, expected to be at 9.5º east longitude, will be performing the HRFI (High spatial Resolution Fast Imagery) mission needs, with an FCI operating in Rapid Scanning mode over ¼ of Full Disk with 2.5 min repeat cycle. It will also constitute the in-orbit hot back up for the prime MTG-I satellite (noting that initially an MSG satellite will be the in-orbit hot backup for MTG-I1. The second in-orbit MTG-I satellite may also operate the LI, DCP and GEOSAR payloads in case the corresponding payload becomes not available or for cross-calibration or to improve the quality of the combined ground processing (Ref. 28).


Figure 29: Overview of the MTG ground segment (image credit: EUMETSAT) 47)



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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 (

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