Meteosat Second Generation - ESA

The 

Satellite 

Development 

BR-153 

November 1999 

Meteosat 

Second 

Generation

Meteosat 

Second 

Generation 

BR-153 

November 1999 

The Satellite Development 

i

ESA BR-153 ISBN 92-9092-634-1 

Technical Coordinators: Bernard Weymiens & Rob Oremus 

MSG Project, ESA/ESTEC 

Published by: ESA Publications Division 

ESTEC, P.O. Box 299 

2200 AG Noordwijk 

The Netherlands 

Editor: Bruce Battrick 

Layout: Isabel Kenny 

Cover: Carel Haakman 

Copyright: © European Space Agency 1999 

Price: 50 DFl / 20 Euros 

i

CONTENTS 

Foreword 1 

1 Introduction 3 

1.1 Programme Outline 3 

1.2 History of the MSG Satellite Concept 4 

1.3 Mission Objectives 6 

2 Programmatics 11 

2.1 Organisation 11 

2.2 Overall Schedule 12 

3 Satellite Development 13 

3.1 Design & Development of the MSG Satellite 13 

3.2 AIT Programme 16 

3.3 Product Assurance 18 

3.4 Image-Quality Ground Support Equipment 20 

4 Payload 

4.1 The Spinning Enhanced Visible and 

Infra-Red Imager (SEVIRI) 23 

4.2 The Mission Communication Package (MCP) 33 

4.3 The Geostationary Earth Radiation Budget 

Experiment (GERB) 38 

4.4 The Search and Rescue (S&R) Mission 40 

5 Satellite Subsystems 43 

5.1 The Structure 43 

5.2 The Unified Propulsion System 45 

5.3 The Attitude and Orbit Control System 47 

5.4 The Electrical Power System 50 

5.5 Data Handling and Onboard Software 52 

iii

iv 

The authors wish to thank those companies and institutes that have provided illustrations and 

photographs for this Brochure, but for which a specific acknowledgement has not been possible. 

Contributors (in alphabetical order) : 

D. Aminou 

J. Azcarate 

C. Bassoua 

H. Bran 

R. Brandt 

A. Camacho 

F. Cavé 

G. Dieterle 

G. Dupré 

S. Fiorilli 

G. Ibler 

K. van ‘t Klooster 

N. Koppelmann 

D. Levins 

H.J. Luhmann 

N. McCrow 

K. McMullan 

H.L. Möller 

J-M. Nonnet 

R. Oremus 

A. Ottenbacher 

L. Ouwerkerk 

J-L. Parquet 

A. Ramusovic 

J. Schmid 

C. Schöser 

W. Schumann 

H. Stark 

I. Stojkovic 

S. Strijk 

W. Supper 

W. Veith 

P. Vogel 

B. Weymiens

Foreword 

Now, in November 1999, MSG-1, the 

Meteosat Second Generation development 

flight model, is about one year away from its 

scheduled launch. Its flight-readiness review 

is planned to take place in August 2000, 

with launch on an Ariane vehicle scheduled 

for the end of October 2000, from Kourou, 

French Guiana. 

We in the Project look forward to these 

events with confidence, secure in the 

knowledge that the flight-model spacecraft 

will deliver excellent performance, based on 

a development plan that includes: 

• the mechanical and thermal tests already 

successfully performed on a Structural 

and Thermal Model spacecraft 

• the electrical performance tests, some of 

which are still ongoing, on an Electrical 

Model spacecraft, and 

• last but not least, subsystem tests 

performed on Flight Model hardware and 

software that prove that the performance 

margins identified on earlier models are 

also available on the Flight Model. 

At this point, integration of the second of 

the three-spacecraft series has also begun, in 

time for its scheduled launch in 2002. 

This Brochure provides a comprehensive 

overview of the history of the MSG 

programme, the mission objectives, which 

are tailored to meet the ever evolving and 

ever more demanding needs of operational 

meteorology and climatology, and the 

design and development of the MSG 

spacecraft, the systems and subsystems of 

which incorporate many technical advances, 

and of their state-of-the-art payloads. 

G. Dieterle 

MSG Project Manager 

1

1 Introduction 

1.1 Programme Outline 

The primary objective of MSG is to ensure 

continuity of atmospheric observation from 

the geostationary orbit at 0.0 degrees longitude 

and inclination, as part of a worldwide, 

operational meteorological satellite system 

consisting of four polar-orbiting and five 

geostationary satellites (the World Weather 

Watch programme of the World Meteorological 

Organisation). 

The Meteosat Second Generation (MSG) 

satellites benefit from several major 

improvements with respect to the first 

generation in terms of performance: 

– 12 imaging channels instead of 3 

– an image every 15 minutes instead of 

every 30 minutes 

– improved spatial resolution, and 

MSG Facts and Figures 

– extra services such as a Search and 

Rescue Mission and an experimental 

Radiation Budget measurement 

instrument, along with much improved 

communications services. 

The MSG development programme is now 

about 1 year away from the first scheduled 

satellite launch. A satellite thermal and mechanical 

model was successfully tested already 

in 1998, an engineering model is currently 

undergoing final testing to demonstrate the 

electro-optical performance and, in parallel, 

the first flight unit (MSG-1) is being 

integrated and tested for an Ariane launch 

from Europe’s Guiana Space Centre in 

October 2000. Two more spacecraft, MSG-2 

and MSG-3, which are identical to MSG-1, 

are also being manufactured to be ready in 

2002 for launch and 2003 for storage. 

Purpose – To make an image of the Earth and its atmosphere every 

15 minutes in 12 spectral bands (2 visible, 1 high-resolution visible, 

7 infrared, 2 water vapour) 

– Dissemination of the image data and other meteorological 

information to data user stations 

Technical Features – Spin-stabilised spacecraft 

– Mass (at launch) about 2 ton 

– Diameter 3.2 m 

– Height 3.7 m 

– Lifetime 7 yr 

– Orbit geostationary 

– Orbit location in the equatorial plane and above 0˚ 

longitude 

– Launch vehicle compatible with Ariane-4 and Ariane-5 

– Launch date October 2000 (MSG-1) 

– Payload • Spinning Enhanced & Visible InfraRed 

Imager (SEVIRI) 

• Geostationary Earth Radiation Budget 

(GERB) Instrument 

• Search & Rescue (S & R) Transponder 

• Mission Communication Package (MCP) 

3

The MSG programme is a co-operative 

venture with Eumetsat, the European 

Organisation for the Exploitation of 

Meteorological Satellites, based in 

Darmstadt, Germany. For the first MSG 

satellite, Eumetsat is contributing about 

30% of the development cost of the ESA 

programme and is financing 100% of the 

two additional flight units, MSG-2 and 

MSG-3. In addition to having overall system 

responsibility with respect to end-user 

requirements (i.e. operational meteorology 

from geostationary orbit), Eumetsat is also 

developing the ground segment and 

procuring the three launchers, and will 

operate the system nominally from 2001 

until 2012. 

The MSG programme is based on the 

heritage of the first-generation Meteosats, 

which have now been operated for about 

22 years with 7 consecutive satellites in 

orbit. This allows the technological risk to be 

kept to a minimum. Moreover, costs are also 

being kept to a minimum thanks to the lowcost 

spinning-satellite design principle used 

and due to the economy of scale of a 

three-satellite procurement in combination 

with contracting rules with industry such as 

firm fixed pricing and incentives based on 

meeting schedule and on in-orbit 

performance. 

MSG is an ESA Optional Programme, which 

was started in 1994 and is funded by 

thirteen of the Agency’s Member States: 

Austria, Belgium, Denmark, Finland, France, 

Germany, Italy, the Netherlands, Norway, 

Spain, Sweden, Switzerland and the United 

Kingdom. 

4 

1.2 History of the MSG 

Satellite Concept 

The concept of the Meteosat Second 

Generation (MSG) satellites has been 

developed through a series of workshops 

organised by ESA with the European 

meteorological community, which started in 

Avignon, France, in June 1984. 

This first MSG workshop identified the major 

future requirements for space meteorology 

in Europe as follows: 

• geostationary satellites providing highfrequency 

observations 

• an imaging mission with higher 

resolution and more frequent 

observations than the first-generation 

Meteosats 

• an all-weather atmospheric-sounding 

mission. 

Based on the Avignon workshop, three 

expert reports on imagery, infra-red and 

millimetre-wave sounding and on data 

circulation were commissioned by ESA. 

The reports on imagery and sounding were 

presented to a second workshop with the 

European meteorological community in 

Ravenna, Italy, in November 1986. That 

workshop confirmed the basic requirements 

of the Avignon workshop and provided 

some updates and refinements. 

The data circulation report was reviewed at 

a workshop in Santiago de Compostela, 

Spain, in May 1987. This workshop 

recommended two important changes 

concerning the Data Circulation Mission 

(DCM) of the first-generation Meteosat

satellites: the processed image data must be 

available within 5 minutes of acquisition, as 

required for nowcasting applications, and 

the current analogue WEFAX service to 

secondary user stations must be replaced by 

a digital format. 

In 1986, a new European intergovernmental 

organisation called Eumetsat was set 

up in Europe to ‘establish, maintain, and 

operate a European system of operational, 

meteorological satellites’. Since then, ESA 

has been collaborating with Eumetsat on 

the definition of the MSG satellites. 

In 1987, ESA initiated several instrument 

concept studies, covering: 

• a visible and infra-red imager (VIRI) 

• an infra-red sounder (IRS) 

• a microwave sounder (MWS) 

• the data-circulation mission (DCM) 

• the proposed scientific instruments. 

Parallel studies of an 8-channel VIRI and of 

the infra-red sounder were performed by 

industry, and they demonstrated the basic 

feasibility of these instruments. The 

microwave sounder was studied via parallel 

contracts, which revealed major problems 

with respect to, for example, mass, diameter 

and sensitivity. 

In the same year, ESA also provided parallel 

contracts to study possible satellite 

configurations for MSG. As a result, a spinstabilised 

satellite configuration was 

excluded due to the presence of the 

microwave sounder. Dual-spin configurations 

were considered but rejected as the 

MWS and IRS instruments require very stable 

pointing of the platform, whilst the IRS 

instrument requires a very stable rotation of 

the drum, and these two requirements 

cannot be satisfied simultaneously. 

Accordingly, the only viable configurations 

for the multi-instrument satellites were 

three-axis-stabilised configurations. 

These results were presented at a workshop 

with Eumetsat and the meteorological 

community in Bath (UK) in May 1988. As a 

conclusion of this workshop, the overall 

mission philosophy was again endorsed, 

while some mass-driving requirements were 

reconsidered and eventually revised. 

However, a few months later further doubts 

were raised about the usefulness of the 

sounding mission, as proposed in Bath, and 

about the relationship between the 

sounding mission of MSG in a geostationary 

orbit and sounding missions from polarorbiting 

satellites. As a consequence, the 

mission requirements were again 

reconsidered, and further mission studies 

were called for. 

The essential point of the reconsideration of 

the mission requirements was that some 

sounding capability had to be retained. It 

was proposed to achieve this by adding 5 

additional narrow-band channels to the 

imager VIRI that had been defined in 

Avignon and in Bath, in order to obtain a 

pseudo-sounding capability. Consequently, 

the corresponding instrument was then 

named the ‘Enhanced VIRI’, or EVIRI. Thus, 

further mission-feasibility studies were 

requested by Eumetsat and initiated by ESA. 

On the basis of the results of these 

deliberations and a recommendation by 

5

ESA, the Eumetsat Council determined in 

June 1990 that: 

• MSG should be a spin-stabilised satellite 

• the spin-stabilised satellite should have a 

capability for air mass analysis as the 

essential part of the former sounding 

mission and a high-resolution visible 

channel. 

Following this decision, and the new 

requirement that the Spinning Enhanced 

Visible and Infra-Red Imager (SEVIRI) should 

also be capable of providing data for air 

mass analysis, ESA conducted an 

assessment study of the feasibility of 

accommodating the extra channels into 

SEVIRI. Originally, EVIRI had 8 channels, 

and the new SEVIRI requirements called for 

14 channels (1 high-resolution visible, 3 in 

the VNIR, and 10 in the IR). 

The assessment study concluded that the 

imager could be expanded to 

accommodate 12 channels in total 

(1 high-resolution visible, 3 in the VNIR, 

and 8 in the IR). The requirement for 

10 cooled IR channels was essentially not 

feasible, given the cost and schedule 

constraints. 

Finally, with Eumetsat endorsement, ESA 

initiated the development of a spinstabilised, 

geostationary satellite with a 

12-channel imager, called the ‘Meteosat 

Second Generation’ satellite. 

6 

1.3 Mission Objectives 

As the successor of the Meteosat firstgeneration 

programme, MSG is designed to 

support nowcasting, very short and short 

range forecasting, numerical weather 

forecasting and climate applications over 

Europe and Africa, with the following 

mission objectives: 

• multi-spectral imaging of the cloud 

systems, the Earth’s surface and radiance 

emitted by the atmosphere, with 

improved radiometric, spectral, spatial 

and temporal resolution compared to the 

first generation of Meteosats 

• extraction of meteorological and 

geophysical fields from the satellite image 

data for the support of general 

meteorological, climatological and 

environmental activities 

• data collection from Data Collection 

Platforms (DCPs) 

• dissemination of the satellite image data 

and meteorological information upon 

processing to the meteorological user 

community in a timely manner for the 

support of nowcasting and very-shortrange 

forecasting 

• support to secondary payloads of a 

scientific or pre-operational nature which 

are not directly relevant to the MSG 

programme (i.e. GERB and GEOSAR) 

• support to the primary mission (e.g. 

archiving of data generated by the MSG 

system).

The mission objectives were subsequently 

refined by Eumetsat, taking into 

account further evolutions in the needs 

of operational meteorology, and resulted in: 

• the provision of basic multi-spectral 

imagery, in order to monitor cloud 

systems and surface-pattern development 

in support of nowcasting and shortterm 

forecasting over Europe and 

Africa 

• the derivation of atmospheric motion 

vectors in support of numerical weather 

prediction on a global scale, and on a 

regional scale over Europe 

• the provision of high-resolution imagery 

to monitor significant weather evolution 

on a local scale (e.g. convection, fog, 

snow cover) 

• the air-mass analysis in order to monitor 

atmospheric instability processes in the 

lower troposphere by deriving vertical 

temperature and humidity gradients 

• the measurement of land and sea-surface 

temperatures and their diurnal variations 

for use in numerical models and in 

nowcasting. 

Imaging Mission 

To support the imaging mission objectives, 

a single imaging radiometer concept known 

as the Spinning Enhanced Visible and Infra- 

Red Imager (SEVIRI) has been selected. This 

concept allows the simultaneous operation 

of all the radiometer channels with the 

same sampling distance. Thus, it provides 

improved image accuracy and products like 

Imaging format 

Imaging cycle 

Channels 

Sampling Distance 

Pixel Size 

atmospheric motion vectors or surface 

temperature and also new types of 

information on atmospheric stability to the 

users. Moreover, as the channels selected 

for MSG are similar to those of the AVHRR 

instrument currently flown in polar orbits, 

the efficiency of the global system will be 

increased owing to the synergy of polar 

and geostationary data. 

2.25 km (Visible) 

4.5 km (IR + WV) 

1 km (HRV) 

3 km (others) 

2.25 km (Visible) 1.4 km (HRV) 

5 km (IR + WV) 4.8 km (others) 

Number of detectors 4 42 

Telescope diameter 400 mm 500 mm 

Scan principle Scanning telescope Scan mirror 

Transmission raw data rate 0.333 Mb/s 3.2 Mb/s 

Disseminated image 0.166 Mb/s 1 Mb/s 

Transmission burst mode 2.65 Mb/s Search & Rescue package 

The imaging mission corresponds to a 

continuous image-taking of the Earth in the 

12 spectral channels with a baseline repeat 

cycle of 15 minutes. The calibration of the 

infra-red cold-channel radiometric drift may 

be performed every 15 minutes, owing to 

the presence of an internal calibration unit 

involving a simple and robust flip-flop 

mechanism and a black body. The imager 

MOP MSG 

30 min 15 min 

Wavelength 

Visible 0.5 - 0.9 HRV 

VIS 0.6 

VIS 0.8 

IR 1.6 

Water vapour WV 6.4 WV 6.2 

WV 7.3 

IR 3.9 

IR window IR 11.5 

IR 8.7 

IR 10.8 

IR 12.0 

Pseudo Sounding 

IR 9.7 

IR 13.4 

DATA CIRCULATION MISSION 

7 

The mission evolution 

from First- to Second- 

Generation Meteosat

Earth imaging frames: 

full image area, HRV 

channel normal mode 

and alternative mode 

provides data from the full image area in all 

channels except for the high-resolution 

visible channel, where the scan mode may 

be varied via telecommand from the normal 

mode to an alternative mode. 

The six channels VIS 0.6, VIS 0.8, IR 1.6, IR 

3.9, IR 10.8 and IR 12.0 correspond to the 

six AVHRR-3 channels on-board the NOAA 

satellites, while the channels HRV, WV 6.2, 

IR 10.8 and IR 12.0 correspond to the 

Meteosat first-generation VIS, WV and IR 

channels. The following channel pairs are 

referred to as split-channel pairs, since they 

provide similar radiometric information and 

may therefore be used interchangeably: VIS 

0.6 & VIS 0.8, IR 1.6 & IR 3.9, WV 6.2 & WV 

7.3, and IR 10.8 & IR 12.0. 

The HRV channel will provide highresolution 

images in the visible spectrum, 

8 

which can be used to support nowcasting 

and very short-range forecasting 

applications. 

The two channels in the visible spectrum, VIS 

0.6 and VIS 0.8, will provide cloud and landsurface 

imagery during daytime. The chosen 

wavelengths allow the discrimination of 

different cloud types from the Earth’s surface, 

as well as the discrimination between 

vegetated and non-vegetated surfaces. 

These two channels also support the 

determination of the atmospheric aerosol 

content. 

The IR 1.6 channel can be used to 

distinguish low-level clouds from snow 

surfaces and supports the IR 3.9 and IR 8.7 

channels in the discrimination between ice 

and water clouds. Together with the VIS 0.6 

and VIS 0.8 channels, the IR 1.6 channel

The spectral characteristics of the SEVIRI channels 

Channel Absorption Band Channel Type Nom. Centre Spectral 

Wavelength Bandwidth 

(µm) (µm) 

HRV Visible High Resolution nom. 0.75 0.6 to 0.9 

VIS 0.6 VNIR Core Imager 0.635 0.56 to 0.71 

VIS 0.8 VNIR Core Imager 0.81 0.74 to 0.88 

IR 1.6 VNIR Core Imager 1.64 1.50 to 1.78 

IR 3.9 IR / Window Core Imager 3.92 3.48 to 4.36 

WV 6.2 Water Vapour Core Imager 6.25 5.35 to 7.15 

WV 7.3 Water Vapour Pseudo-Sounding 7.35 6.85 to 7.85 


IR 9.7 IR / Ozone Pseudo-Sounding 9.66 9.38 to 9.94 



IR 13.4 IR / Carbon Diox. Pseudo-Sounding 13.40 12.40 to 14.40 

may also support the determination of 

aerosol optical depth and soil moisture. 

The IR 3.9 channel can be utilised to detect 

fog and low-level clouds at night and to 

discriminate between water clouds and ice 

surfaces during daytime. Furthermore, the 

IR 3.9 channel may support the IR 10.8 and 

IR 12.0 channels in the determination of 

surface temperatures by estimating the 

tropospheric water-vapour absorption. 

The two channels in the water-vapour 

absorption band, WV 6.2 and WV 7.3, will 

provide the water-vapour distribution at two 

distinct layers in the troposphere. These two 

channels can also be used to derive 

atmospheric motion vectors in cloud-free 

areas and will support the IR 10.8 and IR 

12.0 channel in the height assignment of 

semi-transparent clouds. 

The IR 8.7 channel may also be utilised for 

cloud detection and can support the IR 1.6 

and IR 3.9 channels in the discrimination 

between ice clouds and Earth surfaces. 

Moreover, the IR 8.7 channel may also be 

applied together with the IR 10.8 and IR 

12.0 channel to determine the cloud phase. 

The SEVIRI channel, which covers the very 

strong fundamental vibration band of ozone 

at 9.66 µm, denoted as IR 9.7, will be 

utilised to determine the total ozone 

content of the atmosphere and may also be 

applied to monitor the altitude of the 

tropopause. 

The two channels in the atmospheric 

window, IR 10.8 and IR 12.0, will mainly be 

used together with the IR 3.9 channel in 

order to determine surface temperatures. 

The IR 13.4 channel covers one wing of the 

fundamental vibration band of carbon 

dioxide at 15 µm and will therefore mainly 

be utilised for atmospheric temperature 

sounding in support of air-mass instability 

estimation. 

Product Extraction Mission 

The product extraction mission will provide 

Level 2.0 meteorological, geophysical and 

oceanographical products from SEVIRI Level 

1.5 imagery. It will continue the product 

extraction mission of the current Meteosat 

system, and provide additional new 

products. MSG meteorological products will 

be delivered to the meteorological user 

community in near-real-time via the Global 

Telecommunication System (GTS) or via the 

satellite's High-Rate Image Transmission 

(HRIT) and Low-Rate Image Transmission 

(LRIT) schemes. 

9

MSG mission overview 

Data Collection and Relay Mission 

The data collection and relay mission will 

collect and relay environmental data from 

automated data-collection platforms via the 

satellite. The mission will be a follow-on to 

the current Meteosat Data Collection 

Mission, with some modifications as follows: 

• Increased number of international Data 

Collection Platform (DCP) channels 

• Increased number of regional channels 

• Data Collection Platform (DCP) 

retransmission in near-real-time via the 

LRIT link 

• Some of the regional channels will 

operate at a higher transmission rate. 

Dissemination Mission 

The dissemination mission will provide 

digital image data and meteorological 

products through two distinct transmission 

channels: 

• High-Rate Information Transmission 

(HRIT) transmits the full volume of 

processed image data in compressed 

form 

• Low-Rate Information Transmission (LRIT) 

transmits a reduced set of processed 

image data and other meteorological 

data. 

Both transmission schemes will use the 

same radio frequencies as the current 

10 

Meteosat system, but coding, modulation 

scheme, data rate and data formats will be 

different. Different levels of access to the 

high- and low-rate information transmission 

data will be provided to different groups of 

users through encryption. 

The Meteorological Data Distribution 

mission of the current Meteosat system will 

be integrated into the HRIT and LRIT 

missions of MSG. 

Geostationary Earth Radiation 

Budget (GERB) experiment 

The GERB payload is a scanning radiometer 

with two broadband channels, one 

covering the solar spectrum, the other 

covering the infrared spectrum. Data will be 

calibrated on board in order to support the 

retrieval of radiative fluxes of reflected solar 

radiation and emitted thermal radiation at 

the top of the atmosphere with an accuracy 

of 1%. 

Geostationary Search and Rescue 

(GEOSAR) relay 

The satellite will carry a small 

communications payload to relay distress 

signals from 406 MHz beacons to a central 

reception station in Europe, which will pass 

the signals on for the quick organisation of 

rescue activities. The geostationary relay 

allows a continuous monitoring of the 

Earth’s disc and immediate alerting.

2 PROGRAMMATICS 

2.1 Organisation 

The MSG system is developed and 

implemented under a co-operative effort 

by Eumetsat and ESA, with responsibilities 

shared as follows: 

ESA: 

• develops the MSG-1 prototype 

• acts, on behalf of Eumetsat, as 

procurement agent for: 

- MSG-2/3 satellites 

- interchangeable flight-spare equipment 

- Image Quality Ground Support 

Equipment (IQGSE) 

- the ‘Enhanced Suitcase’. 

For the development and follow-up of the 

production of the satellites, ESA has 

established the MSG Project team at its 

European Space Research and Technology 

Centre (ESTEC) in Noordwijk (NL). This team 

is part of the ESA Directorate of Application 

Programmes, within the Earth Observation 

Development Programmes Department. 

Eumetsat: 

• contributes one third of MSG-1 funding, 

and funds procurement of MSG-2/3 

• finalises and maintains the End User 

Requirements for the MSG mission 

• procures all launchers and the services 

for post-launch early operations 

11 

The MSG industrial 

consortium

• develops the ground segment 

• ensures consistency between the system 

segments (space, ground, launcher 

services segments) 

• operates the system (over at least 12 years). 

A project team in Eumetsat acts as the system 

architect and integrator. The development 

and integration of the overall ground 

segment is carried out by the Eumetsat 

team, with the development of the individual 

ground facilities subcontracted to industrial 

companies across Europe. Once integrated 

and fully tested, the MSG system will be 

routinely operated by the Eumetsat 

operations team. 

Industrial Consortium 

For the development, manufacturing, 

integration and testing of the MSG satellites, 

ESA placed a contract with a European 

industrial consortium, led by the French 

company Alcatel Space Industries (Cannes). 

The work has been subdivided over 105 

contracts, which were negotiated with 56 

different companies. 

The UK National Environmental Research 

Council (NERC), acting through the 

Rutherford Appleton Laboratory (RAL), is 

responsible for the provision of the scientific 

12 

payload. The GERB instrument is developed, 

based on funding from the United Kingdom, 

Belgium and Italy, as an ‘Announcement of 

Flight Opportunity’ instrument. This optical 

instrument, monitoring the Earth’s radiation, 

will make use of a small free volume and 

available resources on the spacecraft 

platform. 

Launcher 

Arianespace is providing the launch vehicle 

and all associated launch services. The 

launch will be performed nominally by an 

Ariane-5 vehicle, as part of a dual or triple 

launch. Compatibility with Ariane-4 (as part 

of a dual launch inside Spelda-10) is 

retained as a back-up. 

2.2 Overall Schedule 

The Phase-B activities were started in 

February 1994, during which detailed 

plans and requirements were established, 

necessary for precise definition of the main 

development, qualification and manufacturing 

activities. Phase-C/D started in 

July 1995 and will last until the Flight 

Acceptance Review in August 2000. It will 

cover the detailed design, development, 

qualification and manufacture of the satellite.

3 SATELLITE DEVELOPMENT 

3.1 Design & Development 

of the MSG Satellite 

Heritage 

In order to limit MSG development cost and 

risks, existing hardware/design heritage 

from the Meteosat first generation and 

other satellite programmes has been used 

to the maximum possible extent. This 

approach could be implemented 

successfully for several units within the 

classical support subsystems. 

Within the Electrical Power Subsystem (EPS), 

the Power Distribution Unit (PDU) is based 

on the Cluster/Soho design, and the 

solar-array cell implementation was also 

taken over from Cluster. The batteries are 

based on standard cells from SAFT (F). The 

Data Handling Subsystem is based on a 

standard design from Saab-Ericsson Space 

(S). In the Attitude and Orbit Control 

Subsystem (AOCS), all sensors (ESU, SSU 

and ACU) are off-the-shelf items, with only 

the control electronics having to be 

specially developed. Most of the Unified 

Propulsion Subsystem (UPS) elements are 

off-the-shelf items, and only the tanks and 

Gauging Sensor Unit (GSU) are new 

developments. Nearly all of the Mission 

Communication Package (MCP) units are 

based on the design heritage of the 

Meteosat first generation. The Search 

and Rescue Transponder is a new 

development, and the S-band 

Telemetry/Telecommand Transponder is a 

standard Alenia (I) design. 

The scientific payload (GERB) and the main 

imaging instrument (SEVIRI) are completely 

new developments. 

Model Philosophy 

For all support and MCP units, a model 

concept including a Structural and Thermal 

Model (STM), an Engineering Model (EM), a 

Qualification Model (QM) and Flight Models 

(FM1, 2 and 3) has been implemented. For 

SEVIRI, the EM/QM and FM1 models are 

replaced by an Engineering Qualification 

and a Proto-Flight Model (EQM and PFM). 

The STM units were manufactured 

exclusively for the use in the satellite STM, 

but their number was limited, as flight 

hardware was used as far as possible, e.g. 

solar panels, primary and secondary 

structures, tanks. The EM units were used to 

validate the design and to perform a prequalification, 

consisting of mechanical, 

thermal and electromagnetic compatibility 

tests. The EM units are manufactured with 

standard components. The QM units, 

equipped with High-Rel parts, served to 

perform the standard qualification. For all 

flight-model units, only acceptance tests will 

be performed. 

The concept of pre-qualification of the EM 

units provided a lot of flexibility in the 

QM/FM manufacturing schedule later in the 

programme. It made it possible in many 

cases to advance the FM unit 

manufacturing and the qualification units 

were then completed after FM delivery. 

Rolling-Spare Philosophy 

Since MSG is a multi-satellite programme, a 

rolling-spare philosophy has been adopted; 

for example, the QMs act as spares for 

FM-1. They will, however, be normally used 

on FM-2, with FM-2 units becoming 

available as spares for FM-2, after which 

13

An MSG/MOP 

comparison 

Exploded view of the 

MSG satellite 

MSG MSG 

12 channel enhanced imaging 3 channel imaging 

and pseudo sounding radiometers radiometer 

100 rpm spin- 100 rpm spin- 

stabilised body stabilised body 

Bi-propellant unified Solid apogee 

propulsion system boost motor 

500 W power demand 200 W power demand 

2000 kg in GTO 720 kg in GTO 

Design compatibility Flight qualified with 

with Ariane-4 (Spelda 10) & Ariane-5 Delta 2914, Ariane 1-3-4 

these units will be used on FM-3, while the 

FM-3 units will remain as the ultimate 

spares. 

Satellite Design 

The MSG concept is based on the same 

design principles as the Meteosat firstgeneration 

satellite and is also spinstabilised 

at 100 rpm. A cylindrical-shaped 

solar drum, 3.2 m in diameter, includes in 

the centre the radiometer (SEVIRI), and on 

top the antenna farm. The total height of 

the satellite, including the antenna 

assembly, is 3.74 m. 

The satellite itself is built in a modular way 

and is composed of the following elements: 

• The Spinning Enhanced Visible and 

Infrared Imager (SEVIRI) instrument, 

located in the central compartment of 

the satellite, ensures the generation of 

image data; formatting of image data is 

completed at satellite level before 

transmission to the ground. 

• The Mission Communication Package 

(MCP), including antennas and 

transponders, is in the upper 

compartment. It ensures the transmission 

of image data to the ground and the 

relay of other mission data. 

• The GERB (Geostationary Earth Radiation 

Budget) instrument. 

• The Geostationary Search and Rescue 

(GEOSAR) payload, which is made of a 

transponder with the capacity for 

relaying distress signals. 

• The satellite support subsystems. 

14 

The MSG satellite support subsystems 

consist of: 

• the Data Handling Subsystem (DHSS) and 

the associated Data Handling Software 

(DHSW), which splits into the Application 

Software (ASW) and the Basic Software 

(BSW) 

• the Electrical Power Subsystem (EPS) 

• the Attitude and Orbit Control Subsystem 

(AOCS) 

• the Unified Propulsion Subsystem (UPS) 

• the Telemetry, Tracking and Command 

Subsystem (TT & C) 

• the Thermal Control Subsystem (TCS) 

• the Structure Subsystem and the 

Mechanisms and Pyrotechnic Devices. 

S/L BAND TPA S BAND TTC 

UHF BAND EDA 

ANTENNA PLATFORM 

SEVIRI BAFFLE 

(and COVER) 

SOLAR ARRAY 

PROPELLANT TANKS 

COOLER 

LOWER 

CLOSING 

SUPPORT 

L BAND EDA 

SEVIRI & 

TELESCOPE 

UPPER 

STRUTS 

MAIN 

PLATFORM 

LOWER 

STRUTS 

CENTRAL 

TUBE 

SEVIRI 

SUNSHADE 

and COVER

For its initial boost into geostationary orbit 

as well as for station-keeping, the satellite 

uses a bi-propellant system. This includes 

small thrusters, which are also used for 

attitude control. The MSG solar array, built 

from eight curved panels, is wrapped 

around the satellite body. 

The support subsystems, Data Handling 

(DHSS), Power (EPS), Attitude and Orbit 

Control (AOCS) and the S-band transponders 

are located on top of the main platform, 

together with the Geostationary Earth 

Radiation Budget (GERB) experiment. The 

Unified Propulsion Subsystem (UPS) is located 

on the bottom side of the main platform. 

The antenna platform houses all elements 

of the Mission Communication Package 

(MCP), i.e. electronic units and antenna. 

The Meteorological Payload consists of the 

radiometer (SEVIRI) as the main instrument 

and a scientific experiment, GERB. 

The Mission Communication Package (MCP) 

includes: the raw data links, image 

dissemination link, Search and Rescue 

transponder and the telemetry/ 

telecommand transponders. 

The Electrical Power Subsystem (EPS) 

generates, stores, conditions and distributes 

the power for all subsystems, including 

thermal control and pyrotechnic functions. 

The following units are part of the EPS: 

Solar Array, Batteries, Power Conditioning 

Unit (PCU), Power Distribution Unit (PDU) 

and the Pyrotechnic Release Unit (PRU). 

The Data Handling Subsystem (DHSS), 

consisting of the Central Data Management 

Unit (CDMU) and two Remote Terminal Units 

(RTUs), serves the internal data exchange via 

an Onboard Data Handling (OBDH) bus. 

SEVIRI is directly connected to the OBDH 

bus, whereas all other subsystems are 

controlled and monitored via the RTUs. 

The Attitude and Orbit Control Subsystem 

(AOCS) comprises a Control Electronics Unit 

(AOCE), the Sun and Earth Sensors (SSU, ESU), 

an Accelerometer Package (ACU) and the 

Passive Nutation Dampers (PNDs). The AOCS 

directly commands the Unified Propulsion 

Subsystem (UPS). 

The UPS is a bipropellant system including 

the Liquid Apogee boost Motors (LAMs), the 

Reaction Control Thrusters (RCTs), the 

propellant- and pressurant tanks and all 

necessary valves, filters, pressure regulator, 

pressure transducers and the Gauging 

Sensor Units (GSUs). 

MSG’s mechanical subsystem includes the 

primary structure, the secondary structure 

(LAM support, solar-array fixation) and the 

SEVIRI cooler and baffle cover (which will be 

ejected prior to reaching the final 

geostationary orbit). 

15 

The electrical/ 

functional architecture 

of the satellite

3.2 AIT Programme 

The AIT programme is based on a threemodel 

philosophy, namely: 

• a Structural and Thermal Model (STM) 

• an Engineering Model (EM) 

• Flight Models (FM). 

Electrical integration and testing, which are 

performed on EM and FMs, are done as 

much as possible at subsystem level. At 

satellite level, these subsystems are 

assembled with a guiding principle of the 

necessary minimum of testing. 

The STM 

The main purpose for building an STM is to 

qualify the mechanical structure of the 

satellite, and to validate its thermal 

behaviour. This model serves also for 

mechanical interface verification and to 

establish mechanical procedures. The 

mechanical tests on the STM were 

successfully completed in spring 1999. 

16 

Overview of the major tests at satellite level 

It was subsequently dismantled to recover 

the flight elements from it. 

The EM 

The main purpose of the EM is to verify all 

of the satellite’s electrical interfaces, and to 

demonstrate that the satellite can meet the 

required performance goals. A second 

important task is to establish and validate 

test procedures and databases, together 

with the relevant EGSE. All satellite EM tests 

were performed in Alcatel’s facilities in 

Cannes (F). 

The FM 

The FM undergoes a series of tests to 

demonstrate that it is flight worthy, and that 

it fulfils the performance requirements. 

These tests are the same, or similar to tests 

performed on the STM and EM and will be 

performed at Alcatel in Cannes. 

The Main Test Programme 

– The Thermal-Balance Test was performed 

in the Large Solar Simulator at ESTEC (NL) 

STM EM FM 

• Thermal Balance / Thermal Vacuum Test √ √ 

• Vibration (Sine) √ √ 

• Acoustic Noise √ √ 

• Mass Properties Determination, incl. Balancing √ √ 

• Spin √ √ √ 

• Cover Release √ 

• Separation and Shock √ 

• Integration Test √ √ 

• Integrated System Test (IST 1) √ √ 

• Antenna Tests in CATR √ √ 

• EMC Test √ √ 

• SEVIRI Reference Test Ambient √ √ 

• IST 2 √

In a vacuum environment (1x10 -5 bar), 

various thermal cases were simulated and 

the temperature response of the satellite 

was compared with the predictions of 

the mathematical thermal model. This 

test was performed in spring 1998 with 

good results. 

– Mechanical Tests: Their purpose is to 

verify that the resonance frequencies of 

the satellite are as required by the 

launcher authority, and to 

demonstrate that the mechanical 

construction of the satellite is strong 

enough to withstand all of the 

mechanical forces that it will experience 

throughout its lifetime. 

– Integration Test: Its purpose is to 

establish correct functioning of a 

subsystem, and to verify its interfaces 

with other subsystems. 

– Integrated System Test: The IST verifies 

that the entire satellite functions correctly, 

and that the performance requirements 

can be achieved. 

– Antenna Tests in CATR: The tests 

performed on the Compact Antenna Test 

Range (CATR) are designed to 

demonstrate the performance and 

functioning of the antenna subsystem on 

the satellite. 

– EMC Test: This is the classical test to 

demonstrate electromagnetic 

compatibility of the satellite with its 

expected environment, with at least 6 dB 

design margin. Also included is a test to 

demonstrate non-sensitivity with respect 

to electrostatic discharge. 

– Spin Test: This test is performed with the 

satellite mounted on a spin table, 

rotating at its nominal operational speed 

of 100 rpm. This test validates all 

operations that depend on the spinning 

motion of the satellite, such as correct 

de-spinning of the L-band and UHF-band 

antennas, and the east-west scanning of 

the SEVIRI instrument. 

– SEVIRI Reference Test Ambient: This test 

verifies SEVIRI’s performance under 

ambient conditions. As such, it forms part 

of the IST, but because of the complex 

set-up with a dedicated optical system it 

is designated as a separate test. 

– SEVIRI Optical Vacuum Test: This test 

demonstrates the performance of the 

SEVIRI infrared channels, with the 

detectors operating at temperatures of 

85 to 95 K. To achieve this, the satellite is 

placed in a vacuum chamber, together 

with the same optical system that was 

used for the reference test. 

17 

Thermal-balance 

testing in progress in 

the Large Space 

Simulator at ESTEC in 

Noordwijk (NL)

EMC test in progress in 

the CATR 

General Test Approach 

Basically, the same set of tests is performed 

before and after the mechanical tests, to 

make sure that the mechanical forces 

experienced have no negative influence on 

the performance, with the exception of the 

spin test and the optical vacuum test. 

For subsequent FMs, the test programme is 

slightly more relaxed. The sine-vibration and 

thermal-balance tests are removed, because 

they are essentially design verification tests 

that are no longer required at that stage of 

the programme. 

To support the Assembly, Integration and 

Test (AIT) Programme, a suite of Ground 

Support Equipment (GSE) is needed. It 

consists of Mechanical GSE, Electrical GSE 

and Optical GSE. 

Mechanical GSE (MGSE) 

All items of a mainly mechanical nature 

belong in this group, but range from the 

satellite transport container (ca. 4 m x 

18 

4 m x 5 m), over various types of dollies 

(structures on which the satellite is 

mounted), via lifting devices, to simple 

masts on which to mount test antennas. 

Electrical GSE (EGSE) 

This group of seven computer systems 

contains all equipment needed to operate, 

control and monitor the satellite.They are: 

1. Overall Check-Out Equipment (OCOE) 

2. TM/TC Special Check-Out Equipment 

(SCOE) 

3. EPS SCOE 

4. AOCS SCOE 

5. RF SCOE 

6. Image SCOE 

7. Launch SCOE. 

Working closely together, their tasks range 

from supplying electrical power, to verifying 

the performance of the payload 

instruments. Each one has its own control 

computer, which in turn receives 

instructions from a central controller. 

Operation of the central computer is 

determined by the AIT engineers through 

direct manual input, or execution of preprogrammed 

sequences of commands. 

Optical GSE (OGSE) 

The OGSE comprises all equipment that is 

needed to provide input signals to the 

satellite’s optical sensors and instruments. 

3.3 Product Assurance 

A ‘quality product’ can be defined as one 

that meets the customers’ requirements – 

particularly in terms of performance, 

reliability, durability and usability. In order

to ensure that the MSG satellite is such a 

‘quality product’ meeting the customer’s 

agreed requirements, a set of proven 

activities, to be carried out during design 

inception up to launch, are brought 

together and detailed in a Product 

Assurance (PA) Plan. This ensures that 

quality is built-in right from the start of the 

project. 

The primary elements addressed in the PA 

Plan are: 

– Design/Qualification Reviews 

– Reliability and Safety 

– Critical Items Control 

– Parts, Materials and Processes 

– Software Quality Assurance 

– Audits 

– Production Control 

– Configuration & Documentation Control 

– Cleanliness and Contamination Control. 

Design/Qualification Reviews 

Each design review is a formal 

comprehensive audit of the MSG design, 

and is intended to optimise the design 

approach and achieve the required 

qualification and performances. 

The following satellite-level reviews are 

foreseen: 

– Preliminary Design Review (PDR) 

– Critical Design Review (CDR) 

– Qualification Results Review (QRR) 

– Flight-Acceptance Review (FAR) 

– Launch-Readiness Review (LRR) 

– Commissioning-Results Review (CRR). 

The PDR is a technical review of the then 

current maturity of the design. It also 

includes a PA status review. 

The CDR will freeze the detailed design, 

manufacturing processes and procedures 

in order to define the FM hardware 

baseline. 

A QRR is held to consider the collective 

evidence from tests, inspections, reviews 

and analyses to prove that requirements 

have been met with the margin specified. 

The FAR is held at the end of the FM test 

programme and will establish the 

flightworthiness of the satellite. The FAR 

also gives consent to ship to the launch 

site. 

The LRR is held at the launch site, six days 

prior to launch, to verify whether the whole 

system, including the satellite, the ground 

stations, the LEOP and the launcher are 

ready for launch. 

The CRR will establish the whole system 

after start-up and verify that all satellite 

systems are working in orbit according to 

their design specifications and releases 

routine operations 

Reliability and Safety 

The Reliability and Safety Plan addresses all 

areas that would compromise the life of the 

mission, or affect the staff and the 

environment prior to launch. 

Critical Items Control 

A Critical Items List (CIL) is produced by the 

Prime Contractor and all Subcontractors 

having design responsibility. The list 

includes all activities and precautions taken 

to minimise and control the risks relating to 

these items 

19

Parts, Materials and Processes 

All parts, materials and processes used in 

building the MSG satellite must be qualified 

for use in a space environment and meet ESA 

requirements. 

Software Quality Assurance 

The quality of the mission software is also of 

vital importance as any problem here could 

seriously affect the satellite’s operation. 

Audits 

The Prime Contractor is required to conduct 

audits of his own (internal audits) and of his 

subcontractors’ and suppliers’ (external audits) 

facilities, equipment, personnel procedures, 

services and operations in order to verify 

compliance with the PA requirements. 

Production Control 

Extensive controls are in place during the 

production of the various satellite models. 

These controls provide a fully documented 

overview of all areas, including assembly and 

test and have built-in traceability. Typical 

controls are: 

– Mandatory Inspection Points (MIPs): 

These take place at critical points during 

manufacture. 

– Test Readiness Review (TRR): These take 

place prior to formal acceptance testing 

of the related item. 

– Test Review/Delivery Review Board 

(TRB/DRB): These Boards review test 

results and manufacturing data and 

decide on suitability for delivery to the 

next stage of integration. 

– Material Review Boards (MRBs): These 

Boards are held when a major nonconformance 

has been found against the 

relevant requirements. 

20 

Configuration & Documentation 

Control 

– Configuration Control: Documentation 

for the MSG project is kept under formal 

change control. 

– Non-Conformance Reports (NCRs): 

Closure of major NCRs is essential before 

proceeding to the next level/integration/ 

test. 

– Data Packs: In order to formally complete 

a DRB, a full Acceptance Data Pack (ADP) 

must be approved by PA at the 

appropriate levels. This provides full 

traceability right back to component 

level, and is invaluable in tracking down 

possible causes of problems that may 

occur in the later stages of build/test. 

Cleanliness and Contamination 

Cleanliness is one of the driving elements 

for the satellite’s imaging-mission 

performance. SEVIRI is a contaminationsensitive 

optical and cryogenically cooled 

instrument. It has units that are built to 

different classes of cleanliness and this 

presents a difficult technical situation during 

environmental test phases. 

A contamination-budget assessment has 

been generated to predict the performance 

degradation due to contamination that may 

arise as a consequence of the on-ground 

activities. 

3.4 Image-Quality Ground 

Support Equipment (IQGSE) 

The Image-Quality Ground Support 

Equipment (IQGSE) for the Meteosat Second 

Generation (MSG) satellites is a computer

system for the processing and quality 

measurement of MSG images. The IQGSE 

software is coded in the C language with 

an X/Motif man/machine interface 

operating on a UNIX-based workstation. 

The IQGSE will be used for two different 

purposes: firstly to qualify on-ground the 

geometric image-quality performance of the 

MSG satellite system, and secondly to verify 

in flight the geometric image-quality 

performance of the MSG satellite system 

during the commissioning phase and other 

periods of the satellite’s seven-year design 

lifetime. 

The IQGSE architecture consists of five 

software modules. Its backbone is the 

Image Rectification Software (IRS), which 

computes and applies a high-accuracy 

geometric correction to the raw MSG 

images received from the ground segment. 

Concurrently with the image-rectification 

process, the IRS automatically measures 

absolute and relative landmark 

displacements for a given set of predefined 

landmarks. The IRS output comprises the 

rectified MSG image and the corresponding 

geometric image-quality file that contains 

the landmark processing results. The 

Landmark Catalogue Builder Tool (LCBT), 

the Image Quality Measurement Tool (IQMT) 

and the Performance Analysis Tool (PAT) 

support the IRS. The LCBT builds and 

maintains the landmark catalogue using the 

World Vector Shoreline database. The IQMT 

measures automatically or interactively 

the absolute and relative landmark 

displacements, while the PAT computes 

the image-quality figures of merit from the 

geometric image-quality file. Finally, there is 

the MSG Image and Data Simulator (MIDAS) 

in order to validate the Image Rectification 

Software before the launch of the first MSG 

satellite. 

The Image Rectification Software (IRS) 

Module comprises four main functions: the 

pre-processing, the navigation filter, the 

image rectification, and the landmark 

processing function. The pre-processing 

function converts the on-board time to 

Universal Time (UT) and determines the 

satellite spin period and the line start 

delays. Furthermore, it calculates the Sun-to- 

Earth centre angle by extracting the Earthto-space 

and space-to-Earth transitions, and 

performs the star detection. The navigation 

filter function determines a parameter state 

vector describing, for example, the satellite 

spin-axis attitude, the satellite orbit, the 

satellite rigid-body wobble, and the detector 

alignments within the focal plane. 

Eventually, the image rectification function 

performs the line-start jitter compensation 

and the image re-sampling. Simultaneously 

with the real-time rectification, the landmark 

processing function measures the rectified 

image quality on up to 1000 landmarks. 

21 

The Image 

Rectification Software 

(IRS) concept

4 PAYLOAD 

4.1 The Spinning Enhanced 

Visible and Infra-Red Imager 

(SEVIRI) 

The SEVIRI instrument is the primary payload 

of the MSG spacecraft. 

The SEVIRI Instrument 

Characteristics 

Spectral range: 

• 0.4 – 1.6 µm 

(4 visible/near infra-red channels) 

• 3.9 – 13.4 µm 

(8 infra-red channels) 

Resolution from 36 000 km altitude: 

• 1 km in high resolution for visible 

channels 

• 3 km in infra-red and visible channels 

Focal plane cooled to 85/95 K 

Earth scanning achieved by a combination 

of satellite spin (east-west) and mirror 

scanning (south-north). 

• One image every 15 minutes 

• 245 000 full images over 7-year nominal 

lifetime 

Instrument mass: 260 kg 

Dimensions: 

• 2.43 m height 

• 1m diameter (without Sun shield) 

• Power consumption: 150 W 

• Data rate: 3.26 Mbit/s 

SEVIRI Operating Principle 

The SEVIRI instrument’s functional 

architecture is based on four main 

assemblies: 

• the Telescope and Scan Assembly (TSA), 

including the Calibration Unit and the 

Refocusing Mechanism 

• the Focal Plane & Cooler Assembly (FPCA) 

• the Functional Control Unit (FCU) 

• the Detection Electronics (DE) including 

the Main Detection Unit (MDU), the 

Preamplifier Unit (PU) and the Detectors. 

The instrument’s operating principle can be 

summarised as follows: 

• The scan mirror is used to move the 

instrument Line-Of-Sight (LOS) in the 

south-north direction. 

• The target radiance is collected by the 

telescope and focused towards the 

detectors. 

• Channel separation is performed at 

telescope focal-plane level. 

• A flip-flop type mechanism is periodically 

actuated to place the calibration reference 

source into the instrument field of view. 

23 

The main SEVIRI 

instrument unit

Functional schematic 

and operating principle 

of SEVIRI 

The Earth-imaging 

principle 

• Imaging data are directly transferred from 

the MDU to the onboard data-handling 

subsystem. 

• SEVIRI function, control and 

telemetry/telecommand interfaces with 

the satellite are ensured by the FCU. 

The Bi-dimensional Earth Scan 

The basic purpose of the instrument is to 

take images of the Earth at regular intervals 

during a 15-minute image repeat cycle 

(involving a 12 min 30 sec Earth imaging 

phase and an up to 2 min 30 sec 

calibration and retrace phase). 

Earth imaging is obtained by a bidimensional 

Earth scan, combining the 

satellite spin and the scan mirror rotation: 

24 

• The rapid scan (line scan) is performed 

from east to west thanks to the satellite’s 

rotation around its spin axis. The latter is 

perpendicular to the orbital plane and is 

nominally oriented along the south-north 

direction. 

• The slow scan is performed from south 

to north by means of a scanning 

mechanism, which rotates the scan 

mirror in 125.8 µrad steps. A total 

scanning range of ±5.5 deg 

(corresponding to 1527 scan lines) is 

used to cover the 22 deg Earth-imaging 

extended range in the south-north 

direction, and 1249 scan lines cover the 

whole Earth in the baseline repeat cycle. 

The Telescope and Scan Assembly 

The Telescope and Scan Assembly includes 

the telescope optics, the telescope structure 

and the mechanism assemblies. 

The telescope’s basic optical layout is based 

on a three-mirror concept:

• M1: large Primary Mirror, concave 

aspherical, with 510 mm optical 

useful diameter 

• M2: Secondary Mirror, concave 

aspherical, of 200 mm diameter 

• M3: Tertiary Mirror, convex aspherical, of 

60 mm diameter. 

The required focal length (5367 mm) is 

obtained by successive magnification of the 

two mirrors M2 and M3. The total length 

of the telescope structure is 1.3 m. 

The Scan Mirror is located in front of the 

Primary Mirror, close to its focal plane, with 

a tilt of 45˚ relative to the optical path. The 

mirror has an elliptical shape (410 mm semimajor 

axis and 260mm semi-minor axis) and 

an elliptical central hole, which allows the 

optical beam to pass through after its 

reflection towards the primary mirror M1. 

All mirrors are of lightweight construction 

and manufactured from Zerodur. 

The telescope structure relies on the use of 

a central stiff base plate, which interfaces 

with the spacecraft via three isostatic 

mounts. The base plate is manufactured 

from a 70 mm aluminum honeycomb 

sandwich, including 4 mm-thick CFRP face 

sheets on each side. Each functional 

component is attached to the base plate 

through a dedicated support structure: 

• a stiff CFRP cone, providing the aperture 

to the spacecraft baffle and supporting 

the primary mirror M1 

• the Scan Assembly Support Structure, 

consisting of a stiff CFRP U-shaped frame 

and 8 CFRP struts, providing the support 

for the moveable scan mirror and its 

associated mechanisms 

• dedicated isostatic mounts to hold the 

Refocusing Mechanism (REM) located in 

the centre hole of the base plate. The 

M2/M3 mirror support structure 

interfaces with the REM top 

• a tripod carrying the Calibration 

Mechanism 

• a titanium strut arrangement (6 struts 

mounted at the lower side of the base 

plate) to keep the Focal Plane and Cooler 

Assembly in position. 

The main instrument electronics (MDU, FCU 

and PU electronic boxes) are located on the 

MSG main base plate. The SEVIRI Sun shield 

is directly mounted to the MSG spacecraft 

structural cone. 

The mechanical design of SEVIRI includes 

three mechanism assemblies: the Scan 

Assembly, the Calibration Unit and the 

Refocusing Mechanism. 

The Scan Assembly includes the Zerodur 

scan mirror, a scan support structure mainly 

manufactured from CFRP, and the scan 

assembly mechanisms, which are primarily 

composed of: 

• a linear spindle drive utilising a stepper 

motor with redundant windings 

25 

The mirror concept for 

SEVIRI

The Scan Assembly 

mounted on a shaker 

table for test purposes 

The Calibration Unit 

assembly 

• a kinematic link system which transfers 

the longitudinal movements of the linear 

spindle drive into rotations at scan mirror 

level 

• a set of angular contact ball bearings 

(dry-lubricated) allowing for small 

oscillatory rotations of the scan mirror 

• a set of springs attached to the mirror 

rotation axis to allow for spin load 

compensation in-orbit 

• a dedicated Launch Locking Device (LLD) 

to clamp the scan mechanism during 

launch. 

The main purpose of the Calibration Unit 

(CALU) is to allow the calibration of the 

infra-red channels of the radiometer, by 

inserting a Black Body Calibration Reference 

Source (CRS) into the optical beam at the 

M1 focal point. The CALU represents a flip- 

26 

flop type of mechanism based on a DC voice 

coil motor. To limit the shock loads when 

reaching the rest positions, dedicated shock 

absorbers are used. 

The Refocusing Mechanism (REM) allows for 

in-orbit focus adjustments (in 1.4 micron 

steps over a 2 mm range) by moving the 

M2/M3 mirror assembly along the 

instrument’s south-north axis. The REM 

features a stepper motor, a transmission 

gearbox and a roller screw providing the 

translation. The mechanical linear guide is 

provided by the elastic deformation of a sixbladed 

arrangement.

The Focal Plane and Cooler 

Assembly 

The Passive Cooler Assembly (PCA) is a twostage 

passive cooling device, composed of 

the Radiator Assembly (RA) and the Sunshield 

Assembly (SA), which provide the 

infra-red detectors with a cryogenic 

environment (basically 85 K in summer and 

95 K in winter). 

The Sun shield is used to avoid direct solar 

fluxes on the first- and second-stage radiator 

of the RA. Thanks to the design of the 

internal cone (elliptically shaped), the 

secondary flux on the second-stage radiator 

is already minimised. 

The PCA heat radiation towards cold deepspace 

is in the range 10 mW to 10 W. 

One of the RA’s most critical subsystems is 

the Detection Cold Wiring (DCW), which 

provides the electrical connection between 

the detectors located in the cold part (CIRO) 

and the warm part of the instrument (RA 

housing). The DCW needed to be optimised 

in order to comply with the electrical 

requirements whilst minimizing the thermal 

impact due to conductive losses (thermal 

gradient of about 200 K between cold and 

warm parts of the RA). Structurally, the 

CIRO is thermally de-coupled from the warm 

part by a set of low-conductive suspensions 

(12 GFRP struts) and a dedicated GFRP 

cone. 

The PCA is equipped with heaters, in order 

to allow for periodic decontamination of the 

instrument (operations to remove frozen 

contaminants from the cold surfaces). 

The Focal Plane Assembly’s Optical Benches 

(FPOBs) are designed to accommodate the 

12 channels of SEVIRI. The Benches consist 

27 

The Refocusing 

Mechanism 

The Radiator Assembly 

with Optical Benches

The Radiator Assembly 

during integration 

of two main assemblies: the VNIR and HRV 

Optical Bench (VHRO) for the 4 visible 

channels, and the Warm/Cold IR Optical 

Bench (WIRO/CIRO) for the 8 infra-red 

channels. The CIRO will be thermally 

regulated at 85 and 95 K depending on the 

solstices and on the cooler capabilities 

during MSG’s lifetime, whilst the VHRO is 

regulated at 20ºC. 

The FPOBs support the detectors and 

perform the appropriate imaging after the 

in-field beam separation at the telescope 

focal-plane level. Thus, most of the SEVIRI 

spectral, geometric and radiometric 

performances rely directly on the FPOB’s 

design and performance. 

28 

The Functional Control Unit (FCU) 

provides the SEVIRI command, control 

and interfaces with the MSG spacecraft’s 

on-board data handling subsystem. 

The FCU has three major sections: 

• the core section including the 

functional mode and sequence 

management 

• the mechanism section (electronics 

driving the mechanisms) 

• the heater and telemetry section 

dedicated to thermal power 

management as well as telemetry 

conditioning and management; 

The thermal control of the instrument is also 

managed by the FCU.

The Detection Electronics (DE) consist of the 

detectors, the Pre-amplifier Unit (PU) and 

the Main Detection Unit (MDU). 

The 12 SEVIRI channels have 8 Infra-Red (IR) 

detectors and 1 High Resolution detector in 

the Visible (HRV), 2 Visible and 1 Near IR 

(NIR). The IR detectors are all in mercurycadmium 

telluride, whereas the visible 

detectors are in silicon and the NIR detector 

is in indium-gallium arsenide. The detectors 

are shaped and sized to satisfy both the 

radiometric and imaging performance 

requirements of the SEVIRI instrument. 

The signal acquired by each detector of the 

42 chains is first amplified by the Preamplifier 

Unit (PU). The PU uses a general 

design with a modular approach common 

to all photovoltaic and photoconductive 

amplifiers. This subsystem consists of three 

assemblies: 

• The Cold Unit (CU) containing the frontend 

parts of the IRPV chains. This transimpedance 

amplifier common to all PV 

chains is implemented for impedance 

matching and for low-noise 

amplification. 

• The Warm Unit (WU) is devoted to the 

front-end parts of HRV/VNIR preamplifiers. 

• The PU main box contains the remaining 

electronics dedicated to shaping the 

analogue signal to the specified values, 

and includes telemetry/telecommand 

interfaces. 

The Main Detection Unit (MDU) contains 

the signal-processing electronics, including 

signal conditioning, anti-aliasing filtering, 

sampling and conversion of analogue 

29 

CIRO equipped with 

cold channels and 

wiring 

The hardware 

elements of the 

detection chain

signals into digital signals. The sampling 

delays are adjustable via telecommand, for 

all 42 chains of SEVIRI. The actual quantification 

is made inside the MDU by a 12-bit 

ADC, for an effective 10-bit resolution at the 

electronics output, after digital dynamicoffset 

and fine-gain corrections. Auxiliary 

data coming from the telemetries, which 

are needed for radiometry and image 

processing, are added to the detection data 

for image processing on the ground. 

A star-sensing function is implemented in 

the MDU. It is activated whenever the star- 

30 

sensing windows are telecommanded. No 

processing at SEVIRI level (filtering or 

dynamic offset correction) is applied to the 

star-sensing function. This raw data is sent 

to the spacecraft in the same way as any 

other auxiliary data. 

SEVIRI Performance Verification 

The on-board calibration process for the IR 

channels of the Imaging Radiometer 

consists of three steps: 

• measuring the cold deep-space radiance 

for the determination of the instrument 

self emission 

Major SEVIRI engineering-model radiometric and imaging performances: comparison of 

specifications and test results* 

Specifications Test Results Margins 

Radiometric Noise Specified per channel Compliant at BOL test Large margins 

Sampling Distance S/N 1 km for HRV, All channels compliant N/A 

3 km for the other channels. under worst case conditions 

Registration Errors Specified between channels Compliant Large margins 

in both E/W and S/N directions 

MTF and Image Quality Specified per channel See examples Sufficient margins 

with templates 

Radiance Response Specified per channel Compliant Large margins 

Spectral Response Specified per channel Compliant As specified 

and Stability within the template 

Scan Motion Specified for S/N Stable and Compliant Large margins 

scanning / pointing 

On-Board Calibration Specified to 0.6K Compliant Large margins 

accuracy at EOL 

*BOL = Beginning Of Life; EOL = End Of Life; MTF = Modulation Transfer Function (image-quality indicator); S/N and E/W = South/North (scan mirror line by line movement) and East/West 

(satellite revolution).

SEVIRI EM channel registration 

Units in E/W on S/N on E/W on IR S/N on IR E/W on IR, S/N on IR 

Km HRV/VNIR HRV/VNIR channels channels HRV/VNIR HRV/VNIR 

SSP (static) channels (static) (static) 

TP1 0.049 0.083 - - - - 

TP2 0.042 0.105 - - - - 

TP3 0.049 0.099 2.559 0.395 1.484 53.829 

TP4 0.027 0.097 2.438 0.383 1.494 53.506 

TP5 0.071 0.094 2.464 0.403 1.487 53.565 

TP6 0.057 0.096 2.456 0.409 1.493 - 

TP7 0.948 0.096 - - - - 

TP8 0.047 0.102 - - - - 

* Column 1 describes the Test Phases (TP); Columns 2 to 5 show the resulting registration error between Test Phases for both Visible and Infrared Channels. 

Column 6 and 7 describe registration between Visible and Infrared Channels. This shows a stable SEVIRI instrument when submitted to various thermal 

environments. Note: TP3 covers the SEVIRI Cold Operational (COP) phase at 85K, TP4 the SEVIRI Cold Operational phase (COP) at 95K, TP5 the SEVIRI Hot 

Operational (HOP) phase and TP6 the SEVIRI PCA Hot case. 

SEVIRI EM noise budget at Beginning of Life (BOL) 

Channel HRV VNIR VNIR NIR IR IR IR IR IR IR IR IR 

(µm) 0.6 0.8 1.6 3.9 6.2 7.3 8.7 9.7 10.8 12.0 13.4 

Specification 1.07 0.53 0.49 0.25 0.35 0.75 0.75 0.28 1.50 0.25 0.37 1.80 

(K) 

Prediction 0.47 0.13 0.14 0.07 0.14 0.28 0.14 0.11 0.36 0.12 0.17 0.47 

(K) 

Measured 0.43 0.16 0.14 0.07 0.11 0.19 - 0.07 0.21 0.07 0.11 0.23 

(K) 

* For the end-of-life assessment, about 30% to 50% margin has to be considered, depending on channels 

Left: Example of Spectral Response of SEVIRI IR 3.9 Channel (EM) 

Right: Scan Mirror Line of Sight (LOS) Evolution during Nominal Full Imaging (EM) 

31

SEVIRI engineeringmodel 

integration at 

satellite level 

• measuring the radiance coming from the 

on-board black body (at temperature 

T 0 ) resulting in an output (in counts) 

as measured by SEVIRI; the black-body 

true radiance is determined through the 

knowledge of the thermal and optical 

properties of the instrument 

• measuring the on-board black body at 

temperature T 0 +∆T with ∆T~20 K; this 

last measurement is performed to help in 

correcting the impact of the elements 

that are not in the beam path of the 

black body, namely the scan mirror and 

the M1 mirror and its baffle. 

32 

The VNIR channel calibration is based on a 

vicarious calibration consisting of measuring 

some known landmarks on Earth. 

SEVIRI Development Status 

The SEVIRI Engineering Model has 

successfully passed all instrument-level 

testing and has demonstrated that the 

design meets the specification. In the 

meantime, the SEVIRI EM has been 

integrated into the EM satellite, where 

environmental testing has started. 

The SEVIRI Proto-Flight Model (PFM) has 

completed instrument-level testing with the 

same success and has been shipped to 

Alcatel (F) for integration into the satellite.

4.2 The Mission 

Communication Package 

(MCP) 

MCP Antenna Subsystem 

The MSG telecommunications system has a 

number of tasks, each of which requires a 

particular antenna: 

• Reception of telecommands and 

transmission of housekeeping data. The 

TT&C S-band transponder is used for this 

task and is connected to a dedicated 

telemetry and telecommand antenna 

(TT&C antenna) 

• Transmission of the measured radiometer 

(SEVIRI) data, coming from the datahandling 

subsystem, to the primary 

ground station. The electronically despun 

antenna (EDA) is used for this task in 

L-band. 

• Reception of pre-processed images with 

associated data. A toroidal pattern 

antenna (TPA) operating in S-band is 

used for this task. 

• Transmission to users, using the L-band 

EDA antenna for low-resolution and highresolution 

data. 

• Receiving data from Data Collection 

Platforms (DCPs). The electronically 

switched circular array antenna uses the 

UHF-EDA at 402 MHz. 

• Transmission of the DCP data, using the 

L-band EDA antenna. 

• Receiving emergency (Search & Rescue) 

messages using the UHF-EDA at 

406 MHz. 

• Transmission of Search & Rescue 

messages, using the L-band EDA 

antenna. 

The TT&C antenna operating in S-band, is a 

low-gain wide-coverage antenna whose 

design had been optimised for MSG taking 

into account the much larger spacecraft 

body compared to the previous Meteosat 

satellite series. The new design makes use 

of four spiral conductors printed on a 

cylinder and fed in quadrature as the 

radiating elements. In the base of this 

antenna, various hybrids have been 

integrated to provide the required phase 

shifts for the spirals and another to provide 

the hot-redundant connection for the two 

TT&C transponders (both receiver sections 

are permanently on). 

The coverage of this antenna from the 

spinning satellite is from θ = 0˚ (satellite spin 

axis) to θ = 120˚ for all azimuth angles in 

right-hand circular polarisation. 

33 

The Mission 

Communication 

Package (MCP) 

antenna (FM1) at 

Alenia Aerospazio (I), 

with the TT&C antenna 

on top, and the S- and 

L-band Toroidal 

Pattern Antenna (TPA) 

inside the black 

cylindrical radome. The 

L-band Electronically 

Despun Antenna (EDA) 

can be seen in the 

middle, and the UHF- 

EDA in front of it 

The TT & C antenna

The Toroidal L and S-band antennas are 

narrow-band, reduced-height, slotted 

waveguide antennas, which provide 

toroidal patterns in the plane perpendicular 

to the spin axis. They are mounted side-byside 

inside a black-painted radome. The 

low-gain L-band TPA functions as back up 

for the high-gain, L-band electronically 

despun antenna in transmit mode. The Sband 

TPA acts as a receive-only antenna for 

the pre-processed high- and low-resolution 

data uplinked from the primary ground 

station. 

The L-band Electronically Despun Antenna 

(EDA) is used in transmit mode only to send 

the raw image data to the primary ground 

station and the processed data, received via 

the S-band, to the secondary users. As the 

satellite rotates at 100 rpm and the highgain 

antenna beam needs to be aimed at 

the ground continuously, an electronic 

means of despinning this beam in the 

opposite direction to the satellite’s rotation is 

implemented. This antenna is composed of 

32 columns of 4 dipoles each, and is 

mounted in a cylindrical way close to the 

top of the satellite. 

The transmit beam is built up from four or 

five active columns, which are fed by an 

array of: one 4-Way Power Divider (4WPD), 

4 Variable Power Dividers (VPDs), and 8 

Single-Pole Four-Way PIN diode switches 

(SP4T). The VPD allows the RF transmit 

signal to be split into two output signals of 

constant phase, but with seven 

programmable output-level ratios between 

the two outputs. The 8 outputs from the 

VPDs are fed via 8 electronic switches 

(SP4Ts) to the feed boards of the 

34 

32 antenna columns. By switching the right 

amount of power to the right column and 

being synchronised with the satellite spin 

rate, an antenna beam is created which 

appears to be stationary with respect to the 

ground. A high-gain (~ 12 dB) antenna 

beam is thus available, easing the groundstation 

requirements for the secondary user 

community. 

The UHF-band EDA Antenna: To receive the 

meteorological data from the Data Collection 

Platforms (DCPs) operating in the UHF band 

and the newly implemented Search & 

Rescue mission on MSG, an electronically 

switched UHF array of 16 crossed dipoles 

was selected. These dipoles are positioned in 

front of the L-band EDA, which at a distance 

of 3/4 λ acts as a reflector for the UHF array. 

A simplified beam-forming network is 

employed, whereby the outputs of the 

dipoles are connected to the inputs of four 

4-way electronic switches, which in turn are 

connected to the inputs of a 4-way power 

combiner. Of the 16 dipoles, four are used 

to form the beam whereby the next dipole is 

selected every 22.5˚ synchronised with the 

satellite’s spin rate. 

To control and supply all of the complex 

timed switching for the various active 

elements of the antenna subsystem, a 

dedicated equipment item known as the 

Common Antenna Control Electronics 

(CACE) is used. This equipment receives 

synchronisation signals from the datahandling 

subsystem and generates the 

correctly timed drive signals for the SP4Ts 

and VPDs in the antenna subsystem. Apart 

from the normal despun mode, this 

equipment also allows the antenna to be

put into a fixed-beam mode, which permits 

the antenna beam pattern to be measured 

on the ground or in orbit. 

MCP Transponder Subsystem 

On board the satellite, the MCP 

Transponder Subsystem’s tasks are the 

reception, amplification and transmission of 

the following channels: 

• Raw Data channel: down-linking to the 

Primary Ground Station (PGS) of the 

SEVIRI (and GERB when applicable) raw 

data stream, plus auxiliary/ancillary 

information received from the Data 

Handling Subsystem. 

• HRIT channel: high-data-rate 

dissemination to the user community 

(High-Rate User Stations, or HRUSs) of 

processed meteorological data and 

images received from the PGS. 

• LRIT channel: low-data-rate dissemination 

to the user community (Low-Rate User 

Stations, or LRUSs) of processed 

meteorological data and images received 

from the PGS. 

• DCP channel: relay of messages from the 

Data Collection Platforms to the PGS for 

further distribution. 

• Search & Rescue channel: relay of distress 

signals from emergency beacons on the 

visible Earth’s disc to dedicated ground 

stations (COSPAS/SARSAT network). 

The raw data signal coming from the Data 

Handling Subsystem is fed into the Raw 

Data Modulator (internally redundant) 

equipment, which performs the QPSK 

modulation before entering the 

Intermediate Frequency Processor (IFP). 

The IFP also receives the HRIT and LRIT 

signals coming from the S-band antenna 

via the S-band filter and the S-band receiver 

(two in cold redundancy) which contain the 

necessary low-noise amplification and 

frequency down-conversion. The IFP 

equipment, which operates in cold 

redundancy, filters and up-converts the 

three signals separately and amplifies them 

to a selected output level or with a certain 

fixed received-signal (RD, HRIT and LRIT) 

gain set by ground command. The output 

signals of the IFP drive the Solid-State Power 

Amplifiers (SSPAs) directly to their chosen 

operating points. 

The multi-carrier DCP channel, which can 

be composed of up to 460 individual 

carriers, enters the transponder together 

with the Search and Rescue signal via the 

UHF filter and feeds the two UHF receivers 

(configured in cold redundancy). They 

perform the low-noise amplification 

and frequency up-conversion to the 

corresponding down-link frequency in 

L-band. The DCP signal is then forwarded 

to the SSPA matrix for further amplification. 

The SSPA matrix is composed of four SSPAs 

(output power about 10 W per amplifier) in 

a 4/3 redundancy scheme. One SSPA is 

allocated to the HRIT channel, one is used 

by the RD and LRIT channels simultaneously, 

one is dedicated to the DCP channel, and 

35 

The MCP subsystem 

for the FM1 satellite 

being integrated and 

tested at Alenia 

Aerospazio (I)

The MCP block 

diagram 

TT&C transponder 

block diagram 

36 

MCP communication-link characteristics and associated frequencies 

Raw Data HRIT LRIT DCP S&R 

Up-link Not 2015.65 2101.5 402.06 406.05 

frequency (MHz) applicable 

Down-link 1686.83 1695.15 1691.0 1675.281 1544.5 

frequency (MHz) 

Useful signal 5.4 1.96 0.66 0.75 0.06 

bandwidth (MHz) 

Bit rate 7.5 Mbps 2.3 Mbps 290 kbps 100 bps 400 bps 

Modulation QPSK QPSK BPSK PM PM

The main performance parameters of the TTC transponders 

Receiver 

• Up-link frequency 2068.6521 MHz MSG-1 

2067.7321 MHz MSG-2 

2069.5729 MHz MSG-3 

• Carrier acquisition range –128 dBm to –50 dBm 

• Telecommand operation range –110 dBm to –50 dBm 

• Telecommand modulation scheme PM of subcarrier on up-link carrier 

• Telecommand subcarrier 8 kHz 

• Bit rate 1000 bps 

• Noise figure 3 dB 

Transmitter 

• Down-link frequency (two modes of 2246.5 MHz MSG-1 

operation, coherent or non-coherent 2245.5 MHz MSG-2 

w.r.t. the up-link frequency) 2247.5 MHz MSG-3 

• Output power 3 W 

• Telemetry modulation scheme PM of subcarrier on down-link carrier 

• Telemetry subcarrier 65.536 kHz 

• Bit rate 8192 bps 

Ranging Channel 

• RNG tone capability 100 – 300 kHz 

• RNG channel video bandwidth 650 kHz 

Power Consumption 

• 2 Rx ON, 2 Tx OFF 12.4 W 

• 2 Rx ON, 1 Tx ON 32.4 W 

Subsystem Mass 7900 g 

37 

An MSG TT&C 

transponder

the remaining redundant SSPA can be used 

by any of the other channels in case of 

failure. 

The Search and Rescue signal is preamplified 

by the UHF receiver and then 

further filtered, frequency up-converted and 

power-amplified in the S&R Transponder. 

The objective is to provide support to the 

international COSPAS-Sarsat humanitarianoriented 

Search and Rescue Organisation. 

After power amplification, all of the 

channels (RD+LRIT, HRIT, DCP and S&R) are 

filtered and combined in the output 

multiplexer (OMUX), before being fed to the 

Antenna Subsystem. 

TT&C Subsystem 

The Telemetry, Tracking and Command (TTC) 

Subsystem consists of two S-band 

transponders and performs the following 

functions: 

• Reception and demodulation of the uplink 

command and ranging subcarriers of 

the S-band signal transmitted by the 

ground control station. 

• Delivery of the telecommand video signal 

to the on-board Data Handling 

Subsystem. 

• Modulation of the down-link carrier by 

the received and demodulated ranging 

signal and the telemetry signals received 

from the on-board Data Handling 

Subsystem. 

• Power amplification and delivery of the 

S-band down-link carrier to the Antenna 

Subsystem. 

• The down-link carrier can be generated 

coherently or non-coherently with respect 

38 

to the up-link carrier received from the 

ground station. 

The TTC Subsystem is composed of two 

identical transponders, each consisting of 

several modules packaged in a single unit. 

The receiver and transmitter of each 

transponder are electrically independent, 

except for the necessary interconnections to 

perform the ranging operations. The 

receivers of the transponders are always ‘on’ 

at any time during the satellite’s lifetime, 

while the transmitters are operated in cold 

redundancy. 

4.3 The Geostationary 

Earth Radiation Budget 

Experiment (GERB) 

MSG satellite resources allow for the 

accommodation of an Announcement of 

Opportunity instrument. The ensuing flight 

opportunity has been taken up by a 

European consortium (led by the UK 

Natural Environmental Research Council 

acting through the Rutherford Appleton 

Laboratory), which has developed and 

manufactured a new optical instrument, the 

GERB. With a three-mirror telescope and all 

supporting functions, GERB will measure 

the components of the Earth’s Radiation 

Budget (ERB), which is the balance 

between the incoming radiation from the 

Sun and the outgoing reflected and 

scattered solar radiation plus the thermalinfrared 

emission to space. 

Observations from space have a central role 

in understanding the Earth’s Radiation 

Budget since they are quasi-global. GERB

will measure energies leaving the Earth over 

the geographical region seen by MSG, 

thereby exploiting the excellent temporal 

sampling possible from geostationary orbit. 

These observations are the first of their kind 

and will make an important contribution to 

the enhancement of the climate simulation 

models (diurnal cycle), with strong practical 

relevance to global climate change, food 

production and natural-disaster prediction. 

GERB consists of two units: 

The Instrument Optical Unit (IOU) which is 

very compact (56 x 35 x 33 cm 3 ), and 

includes essentially: 

• the telescope (three-mirror anastigmatic 

system) 

• the de-scanning mirror for staring at 

appropriate targets 

• the detector (a linear blackened 

thermoelectric array of 256 elements) 

with its signal-amplification and 

processing circuitry (including ASICs and 

a DSP) 

• the quartz filter mechanism used to 

switch the measurement into alternate 

wavebands (total and shortwave) 

• the calibration devices (black body and 

solar diffuser) 

• the passive thermal design. 

The Instrument Electronic Unit (22 x 27 x 

25 cm 3 ), which on one side conditions 

power and signals from MSG to further 

distribute them to the optical unit, and on 

the other collects and formats data 

generated by the IOU before transmitting it 

39 

Components of the 

Earth’s Radiation 

Budget 

The Instrument Optical 

Unit (bottom left) 

The GERB flight 

model (below)

to MSG (owing to its microprocessor, GERB 

has a high level of autonomy). 

The radiometric performance is obtained 

after adequate calibration: 

• On the ground, the instrument has been 

subjected to an extensive characterisation 

programme under vacuum. 

• On board, a solar-illuminated integrating 

sphere and a black-body device with 

known characteristics are implemented in 

the optical unit. 

The scan mirror, which rotates counter to 

the satellite’s spin direction, allows the 

telescope to point successively at the black 

body, the Earth, and the integrating sphere 

within each MSG period. Therefore – 

considering deep-space views also – a highly 

accurate correction of each GERB Earth pixel 

measurement can be performed on the 

ground. 

4.4 The Search and Rescue 

(S&R) Mission 

In addition to serving the primary 

meteorological missions, MSG is also 

equipped with a transponder for the 

Geostationary Search and Rescue service of 

the COSPAS-Sarsat organisation. 

40 

Performance characteristics of the GERB instrument 

Wavebands Total 0.32 – 30µm 

Shortwave (SW) 0.32 – 4µm 

Longwave (LW) 4 – 30µm 

Radiometry SW LW 

Absolute Accuracy

S&R transponder block using a SAW filter 

operating in the up-link frequency band. 

The COSPAS-Sarsat S&R frequencies are not 

very different from those of the Meteosat 

data links, and with just a little extra 

development effort the S&R requirements 

have been accommodated on MSG. 

Nonetheless, since S&R is not part of the 

meteorological objectives of the MSG 

programme, it was agreed to implement this 

payload subject to some constraints, namely: 

• no interference with the meteorological 

missions 

• switch-off in the event of a power 

shortage 

• minimum mass and cost. 

These constraints have been fulfilled by 

making the S&R transponder nonredundant. 

Both the UHF receive antenna 

and the L-band transmit antenna provide 

coverage of the full Earth as seen from 

longitude 0.0°. The geographical area 

covered complements the existing 

Cospas-Sarsat geostationary coverage very 

well. 

41 

The geographical area 

covered by MSG for 

Search and Rescue

5 SATELLITE SUBSYSTEMS 

5.1 The Structure 

The MSG satellite is spin-stabilised. The body 

is a cylindrical-shaped drum, 3.218 m in 

diameter. The total height of the satellite, 

including the antenna assembly, is 3.742 m. 

The outer skin is dedicated to the fixed 

solar array. The internal configuration is 

built around the SEVIRI instrument, 

including a double-stage passive cooler 

accommodated on the lower part of the 

spacecraft. 

The MSG satellite structure consists of two 

main parts: the Primary Structure (191.6 kg) 

providing support for payload and most 

subsystems, and the Secondary Structure 

(27.5 kg) providing support for UPS (Unified 

Propulsion System) and EPS (Electronic 

Power Subsystem) equipment. The SEVIRI 

baffle (15.5 kg) is a light structure that 

protects the instrument’s field of view from 

spurious radiation or pollution. On the 

ground and during launch, a cover protects 

the cooler from pollution. 

Primary Structure 

The main elements of the Primary Structure 

are: 

• the Service Module Structure providing 

support for payloads and for the main 

part of support subsystems equipment 

• the Antenna Platform to accommodate 

the MCP subsystem equipment. 

The Service Module Structure consists of : 

• The Conical Central Tube, based on a 

stringer-stiffened shell design and 

equipped with three rigid interface rings 

for attachment with: 

– launcher adapter and lower struts at 

the lower ring 

– main platform at the upper interface 

ring 

– propellant-tank supports at the upper 

and intermediate rings. 

The Central Tube provides fixation for 

part of the propulsion subsystems pipework, 

and additional interfaces for the 

fixation of the thermal lower closing 

support and umbilical connectors, as 

well as support for the launcherseparation 

actuators. 

• The Main Platform, fixed on the Central 

Tube, manufactured in sandwich form 

with aluminium skins, provides interfaces 

for the SEVIRI instrument and 

accommodates part of the propulsion 

subsystem units on its lower face. 

• A set of lower struts fixed on the Central 

Tube support the main platform edges. 

Two additional struts support the main 

platform, below the two heavy batteries. 

• A set of upper struts connect the Service 

Module to the Antenna Platform. 

43 

MSG Structure: 

Antenna Platform, Main 

Platform, SEVIRI 

Sunshade, Lower 

Support Closing Ring 

and Upper & Lower 

Struts

The Antenna Platform is manufactured in 

sandwich form with aluminium skins, 

providing interfaces for the accommodation 

of both MCP transponders on its lower face, 

and Antenna Assembly on its upper face. 

Secondary Structure and Baffle 

The function here is to provide intermediate 

supports for Unified Propulsion System (UPS) 

and Electrical Power System (EPS) equipment: 

UPS Secondary Structure 

• Two LAM supports, each constituted by 

three pairs of struts providing the LAM 

for iso-static and rigid mounting, 

alignment accuracy and stability 

• Two helium-tank supports, each 

constituted by one tripod and one bipod, 

providing the helium tanks with iso-static 

and rigid mounting. 

• Two E/W (east/west) and two N/S 

(north/south) thruster supports, 

constituted by structural brackets for rigid 

mounting, with vertical adjustment 

capabilities (E/W thrusters). 

• The E/W and the N/S thruster supports 

are fixed, respectively, to the Main 

Platform and to the Antenna Platform. 

EPS Secondary Structure 

• SAP (Solar Array Panel) supports, 

constituted for each of the eight SA 

panels by a set of 6 brackets providing 

the SA panels with rigid and iso-static 

mounting, to allow their thermo-elastic 

dilation. Of the six brackets used for each 

SAP, two are fixed to the Antenna 

Platform and two to the Lower Closing 

Support. 

• A Lower Closing Support (based on a 

light, profiled structure) provides fixation 

44 

to the Thermal Lower Closing Support, 

which sustains the LAM Thermal Closing, 

the valves, and also the lower SAP 

supports. 

SEVIRI Baffle 

The SEVIRI Baffle is a light structure, 

protecting the instrument’s field of view 

from spurious radiation or pollution, and is 

constituted by a main body with three 

structural frames. The main body’s form fits 

the shape of the SEVIRI optical beam, and 

consists of: 

• A metallic envelope, which is an 

assembly of two curved thin shells and 

two lateral iso-grid plates for rigidity 

purposes. 

• Two optical vanes fixed at the end of the 

metallic envelope. 

• A thermal-control interface flange. 

The thermo-optical and optical performances 

are ensured by the optical vanes and 

black-paint coating inside the main body. 

The three frames stiffen the SEVIRI entry 

baffle, and ensure its interface with the 

main platform cover and mechanism. 

Materials 

Primary Structure Materials 

• Aluminium alloy for most of the 

structural parts and machined parts 

(rings, struts brackets) or sheets for the 

Central Tube skins or platform skins. 

• Carbon-Fibre Reinforced Plastic (CFRP) for 

struts of propellant tank supports. 

• Titanium for the most loaded brackets of 

the propellant-tank support struts. 

• Other materials for structural part 

assembly (titanium bolts for strut fittings) 

or bonding (adhesive).

Secondary Structure Materials 

• Aluminium alloy for most of the structural 

parts. 

• Other materials for structural part 

assembly (titanium bolts for strut fittings) 

or bonding (adhesive). 

5.2 The Unified Propulsion 

System (UPS) 

The first-generation Meteosat was equipped 

with two independent propulsion systems. A 

solid-propellant apogee boost motor, MAGE- 

1, and a small hydrazine propulsion system 

served for orbit, attitude, spin and nutation 

control. The MSG UPS combines the two 

propulsive tasks in one common tankage 

and feed system, and it will be a world first 

for a UPS to operate at under 100 rpm. The 

incorporation of a Propellant Gauging 

Sensor Unit (GSU) is an innovative element, 

allowing the user to have an accurate 

knowledge of the propellant remaining 

during the last three years of the mission. 

The significantly larger mass of MSG, 

weighing in at about 2000 kg compared to 

the 720 kg of the first-generation satellite, 

has led to the implementation of a pressureregulated 

bi-propellant propulsion system 

operating with Mono-Methyl Hydrazine 

(MMH) as the fuel and nitrogen tetroxide 

(MON-1) as the oxidiser. This not only 

provides the higher total impulse required 

for the MSG mission, but also leads to an 

improvement in the specific impulse: a 7 % 

increase comparing the LAM with the 

previous solid ABM, and a 15% increase 

when comparing the bi-propellant Reaction 

Control Thrusters (RCTs) with the hydrazine 

monopropellant thrusters used previously. 

The main requirements for the UPS are to 

inject the satellite into geostationary orbit 

after its release from the Ariane launcher. 

This will consume about 83% of the total 

loaded propellant of 976 kg contained in 

four spherical tanks of 750 mm diameter. 

Besides the spin-rate control and the 

attitude manoeuvres, most of the propellant 

will be consumed by inclination control 

(11% of total propellant mass) and 

east/west manoeuvres (4%) throughout the 

seven years of nominal operation. 

The UPS is comprised of the following key 

equipment: 

• two 400 N LAMs 

• six 10 N RCTs 

• eleven fill and drain valves 

• four propellant tanks 

• two latch valves 

• two pressurant tanks 

• three pressure transducers 

• four gauging sensors. 

The design of the UPS, the choice of 

equipment, the manufacturing tools and 

procedures are based on the experience 

acquired during the Spacebus projects. 

Nevertheless, a significant analytical design, 

test and assembly preparation effort was 

required, to adapt the well-known 

integration approach to the totally different 

MSG configuration. 

Most of the equipment is of European 

origin, with only the latching valves, the 

RCT flow-control valves and the propellant 

filter cartridges being procured from the 

USA. The UPS has a mass of 94 kg and is 

operated by the Attitude and Orbit Control 

45

The UPS layout on the 

underside of the Main 

Platform 

Electronics (AOCE). A specially built UPS unit 

tester allows self-standing checkout and 

operation of the UPS at equipment and 

satellite level via a skin connector. 

Subsystem Design 

The main design driver for the UPS is the 

spin environment and the payload cooler 

accommodation in the central cone of the 

Primary Structure. This configuration 

necessitated the placement of two LAMs at 

a radius of 1200 mm from the spin axis. 

It was decided to accommodate the UPS on 

the lower face of the main platform, which 

was not occupied by any other equipment. 

The three main subassemblies, the 

Pressurant Control Panel (PCP) and the two 

46 

Propellant Isolation Assemblies (PIAs) for 

oxidiser and fuel are located between the 

propellant-tank cutouts. The four propellant 

tanks, the two pressurant tanks and the two 

LAMs are mounted via struts and brackets to 

the central cone. The fact that the axial 

thrusters (N/S thrusters) are located on the 

antenna platform required a staggered 

integration sequence, which led to the 

provision of screw joints on pipes leading 

from the main platform to the antenna 

platform. The position of the radial thrusters 

(E/W thrusters) has been chosen such that 

the diagonal use of one upper and one 

lower thruster will always have the satellite 

centre of gravity between them throughout 

the mission. 

The routing of the 90 m of quarter-inch 

titanium tubes needed careful consideration 

regarding the launch and spin environment, 

to avoid tanks being filled-up or depleted 

unsymmetrically or propellant being trapped 

in pressurant lines during initial spin-up. 

Two carbon-fibre-wrapped helium tanks 

(max. operating pressure 275 bar, volume 

35 l) supply – via pyrotechnic valves, a 

pressure regulator and check valves – the 

four propellant tanks (max. operating 

pressure 22 bar, volume 219 l). The 

propellant tanks supply the six RCTs 

arranged in two redundant branches via 

two latching valves and the two LAMs via 

four pyro valves. Minimum fracture safety 

factors were used to optimise tank mass. 

For fill and drain purposes, the propellant 

tank valves are located on the Lower 

Closing Support structure, thereby allowing 

for optimum draining.

Due to the non-availability of a European 

two-stage regulator, it was decided to 

operate the RCTs for the initial spin-up and 

attitude manoeuvres in a pre-blow-down 

mode from the propellant tanks (12 – 8 

bar), prior to pressurisation to 18.5 bar for 

the apogee manoeuvres. This will avoid any 

leakage-related critical pressure increases. 

After station acquisition, the LAMs and the 

pressurisation part will be isolated by firing 

the normally open (NO) pyro valves. 

In order to allow maximum propellant 

utilisation, a high gauging-accuracy 

requirement was specified, leading to the 

design and development of a very precise 

capacitive propellant Gauging Sensor Unit 

(GSU), which is built into the propellant 

tanks. The qualification testing has shown 

that an accuracy of ± 0.05% of total tank 

volume can be achieved for the last three 

years of mission. Such a performance has 

never previously been achieved on a satellite 

propulsion system and is 30 times better 

than with existing techniques. It is important 

in so far as it allows accurate planning for 

the final de-orbiting manoeuvre. 

5.3 The Attitude and Orbit 

Control System (AOCS) 

Like the first-generation Meteosats, MSG 

is equipped with a similarly designed 

subsystem whereby the attitude, nutation, 

spin rate and reference pulse are generated 

by specific Sun, Earth and acceleration 

sensors. Many of the off-the-shelf equipment 

items have required changes and additional 

performance testing. The main processing 

unit of the Attitude and Orbit Control 

Electronics (AOCE) and the Passive Nutation 

Damper (PND) are MSG-specific 

developments. 

The changes introduced within the AOCS 

include: 

• synchronisation pulse generation in 

eclipse 

• stable satellite, inertia ratio > 1 

• multi-burn apogee manoeuvres 

• active nutation damping using a microcontroller 

• interface with data-handling software 

• Passive Nutation Damper tuned for 

geostationary orbit (GEO). 

The first change means that the AOCS has 

to provide for a satellite synchronisation 

pulse in eclipse, while the second means 

that the Active Nutation Damping (AND) in 

Geostationary Transfer Orbit (GTO) is 

primarily required in case a liquid apogee 

boost motor fails. This is less critical than on 

the first-generation spacecraft, where the 

non-stable configuration (solid apogee 

boost motor and satellite) needed 

continuous surveillance and active nutation 

control until separation of the boost motor. 

It was decided to limit the utilisation of the 

47 

The first UPS flight 

model (FM1) on its 

transport and 

integration jig

Left: The Attitude and 

Orbit Control 

Electronics (AOCE) unit 

Right: The Attitude 

Sensor Assembly (ASA) 

unit: left to right, 2 

ACUs, connector 

bracket and ASA 

alignment mirror, ESU 

(3 telescopes), and SSU 

(meridian and skew slit) 

AND to GTO and to fine-tune the PNDs for 

the GEO inertia ratios, to achieve optimum 

performance. 

The AOCS configuration with the AOCE and 

the Attitude Sensor Assembly (ASA) consists 

of the following equipment: 

1 x AOCE, internally redundant 

1 x Sun Sensor Unit (SSU), internally 

redundant 

1 x Earth Sensor Unit (ESU), 

3 channels 

2 x Accelerometer Units (ACUs) 

1 x Attitude Sensor Bracket (ASB), 

equipped with connector bracket, 

harness, alignment mirror and 

bonding straps 

2 x Passive Nutation Dampers (PNDs). 

Two sets of System Checkout Equipment 

(SCOE), derived from the subsystem 

electrical ground-support equipment, were 

provided for satellite and launch-activity 

support. Except for the AOCE and PND, the 

48 

other equipment had previously been used 

on scientific and telecommunication 

satellites. The ACU was used on the firstgeneration 

Meteosats. 

The total mass of the AOCS is 16 kg and 

its power consumption (mode-dependent) 

varies between 8 and 14.5 W, with a 

maximum peak during UPS Liquid 

Apogee Motor commanding of 70 W for 

100 ms. 


The major tasks to be fulfilled by the AOCS 

are: 

– Attitude Measurement 

• Sun aspect angle 

• Earth aspect angle 

• Nutation angle and frequency (GTO). 

– Satellite Synchronisation Pulse 

Generation 

• SSP 1, Sun Synchronisation Pulse 

• SSP 2, Earth Synchronisation Pulse 

• Spin rate.

– Nutation Damping 

• Active nutation damping, axial RCTs in 

closed loop (GTO) 

• Passive nutation damping, PNDs 

(GEO). 

– Operational interface with the UPS 

• Monitoring of UPS sensors 

• RCT, LAM and Latching Valve 

command generation and control 

• Monitoring of command duration and 

number of pulses generated. 

Besides servicing these main tasks, the 

AOCS interfaces with the Data Handling 

and the Electrical Power Systems. Integrated 

in one box, AOCE-B is cold-redundant to 

AOCE-A. Significant cross-strapping is 

provided between all sensors and actuators 

within the AOCE. The temperature 

monitoring of the RCTs and LAMs and the 

coils of the latching valves have their own 

redundancies. 

For fast failure identification, it was decided 

to equip the combustion chambers of RCTs 

and LAMs with thermocouples. As a 

thermocouple always needs the reference 

temperature at its junction to the normal 

harness, the thermocouple wires have been 

routed to the AOCE, where this transition is 

performed within the connectors. 

Thermistors installed inside the AOCE close 

to these connectors measure the required 

reference temperature. The thermocouple 

output is then transformed into a standard 

analogue output. The AOCE also provides 

monitoring of secondary voltages, current 

and converter temperatures for 

housekeeping purposes. 

As the spinning of the MSG satellite 

provides a self-stabilised attitude, it was 

decided that most of the manoeuvres 

(except AND) should be open-loop and 

ground-controlled, with two ground 

stations available during GTO. This 

minimises the on-board monitoring and 

reconfiguration effort. In order to further 

protect the satellite against propulsioninduced 

effects (leakage and spurious 

firing), the latching valves will be closed 

after manoeuvres. Action blocks in the 

monitoring and recovery function (Data 

Handling Subsystem software) are therefore 

limited to reacting to spin-rate anomalies, 

synchronisation loss and invalid sensor 

pulses. 

The UPS provides for two redundant 

branches, which are cross-strapped to both 

AOCEs. Three RCTs can provide all necessary 

control torques. For east/west manoeuvres, 

the diagonal radial thrusters are used (e.g. 

R1 and R3) in pulse mode. 

49 

The AOCS/UPS 

actuator arrangement: 

green indicates the 

nominal thruster 

branch, and red the 

redundant branch

Main AOCS requirements and performance 

Function Requirement Result Remark 

5.4 The Electrical Power 

System (EPS) 

The electrical power system is formed from 

five separate elements: a solar-array 

photovoltaic energy source; two nickelcadmium 

storage batteries; a Power Control 

Unit (PCU); a Power Distribution Unit (PDU); 

and a Pyrotechnic Release Unit (PRU). In 

sunlight, the power is generated by solar 

cells. Peak power loads, which exceed the 

solar array’s capabilities, are supplied from 

the batteries through battery-discharge 

regulators. During eclipse operations, all 

power is supplied by the batteries. At the 

end of each eclipse period, the batteries are 

recharged from the solar array. The solar 

array and batteries are designed and sized 

for a 10-year mission lifetime, including 

reliability and failure-tolerance requirements. 

50 

Synchronisation Pulse 

Sun, SSPI < 0.05 deg < 0.046 deg outside the central region of the Sun Aspect Angle 

< 200 ns, jitter < 136 ns under common mode noise and AOCE jitter 

Eclipse, SSP2 < 0.18 deg < 0.175 with 1 rpm spin rate variation into eclipse, 

excluding Earth radiance error 

Active and Passive Nutation Damping (AND & PND) 

AND in GTO 5 to 0.15 deg 5.2< τ < 10 min at 55 rpm and inertia radio 

in < 10 min 1.2 < λ < 1.35 

PND in GEO 0.01 deg to 2 arcs τ < 4 min 50% margin for λ = 1.1 at 40˚C 

in < 5 min and λ = 1.25 at 5˚C 

Spin-Rate Measurement 

GTO, 5-100 rpm < 1 rpm, 5-30 rpm < 0.001 rpm no nutation, no eclipse 

< 0.1 rpm, 30-100 rpm < 0.002 rpm no nutation, no eclipse 

GEO, 99-101 rpm < 0.01 rpm < 0.0002 rpm no nutation, no eclipse 

< 0.1 rpm, 30-100 rpm < 0.05 rpm no nutation, no eclipse 

Spin-Axis-Orientation Measurement 

GTO < 0.03 deg < 0.273 deg all ESUs, using on ground ESU calibration mode, 

no nutation, no wobble 

GEO < 0.1 deg < 0.05 deg no nutation, incl. wobble error 

Nutation Determination 

GTO < 0.01 deg resolution 0.001-0.0023 deg for 0.01-5 deg at 55 rpm 

GEO < 0.003 deg < 0.0023 deg for 0.003-0.12 deg at 100 rpm 


Solar Array 

The solar array is the satellite’s primary 

source of power. It consists of eight curved 

solar-array panels mounted around the 

body of the spacecraft. Seven panels are 

identical and interchangeable standard solar 

panels, and one is a special panel with a 

large cut-out window serving as the SEVIRI 

instrument-viewing aperture. A total of 

7854 high-efficiency silicon solar cells are 

attached to the eight panels. Each 60 mm x 

32 mm cell is covered by a cerium-doped 

cover glass, which has an indium-tin oxide 

coating to make it conductive and thereby 

prevent surface charging. 

Sixty-six cells are interconnected in series to 

obtain a string output of 30 V via a single 

series blocking diode. There are one

hundred and nineteen solar-cell strings 

around the circumference of the spacecraft 

and these are connected in parallel. About 

one third of these solar cells view the Sun at 

any moment during the satellite’s spin cycle. 

The solar array will provides a power output 

of more than 600 W for up to 10 years in 

geostationary orbit. The complete solar array 

weighs 76 kg. 

Batteries 

Secondary power, for peak loads and eclipse 

operations, is provided from two 29 Ah 

nickel-cadmium batteries. Each battery has 

16 series-connected cells, provides an 

average of 20 V, and has an energy capacity 

of 580 Wh. Passive thermal control via a 

radiator plate for cooling plus active 

thermistor-controlled heaters will maintain 

the batteries within their operational limits 

throughout the mission. 

As the batteries are mounted close to the 

spacecraft’s outer circumference, they 

experience a force of 18 g due to its 

100-rpm spin rate. This force was found 

to have a detrimental effect on the 

capacity of the battery cells. Extensive 

investigations and tests established that 

capacity loss is minimised if the smallest 

dimension of the cell is orientated in the 

direction of the acceleration force. 

Each battery weighs 27.5 kg, giving a total 

battery mass of 55 kg. 

The Power Control Unit (PCU) 

The PCU provides centralised management 

of the 27 V power bus. It converts the 

energy from the solar array and the 

batteries into a regulated bus voltage. The 

unit contains a six-section shunt solar-array 

regulator, four battery-charge regulators and 

six battery-discharge regulators. Included in 

this equipment are the spacecraft power-bus 

and battery-cell management and 

protection functions, telemetry and 

telecommand housekeeping functions, and 

ground-support umbilical interfaces. 

The PCU has been designed with a high 

level of autonomy, redundancy and 

modularity to protect against failures or any 

failure-propagation modes. Reliability and 

failure tolerance is necessary due to the 

intrinsically singular nature of the bus 

voltage supply and the limited energy 

resources available from the solar array and 

batteries. The PCU weighs 23.5 kg. 

The Power Distribution Unit (PDU) 

The PDU is the interface equipment 

between the power subsystem and the 

other spacecraft subsystems and payloads. It 

distributes the power to all spacecraft loads 

through 42 current-limiting switches. These 

switches protect the power bus from 

overcurrent failures in any of the spacecraft 

51 

One of MSG’s two 

nickel-cadmium 

batteries

The Power Distribution 

Unit (PDU) (left) 

The Pyrotechnic 

Release Unit (PRU) 

(right) 

equipment. It also ensures that other 

equipment does not experience powersupply 

disruption during failure recovery. In 

addition, there are 54 simple transistor 

switches in the PDU for thermal-control 

heater on/off switching. Main and 

redundant auxiliary converters supply the 

telemetry circuitry, which monitors the 

power to the users. Bus and battery cell 

undervoltage protection by the shedding of 

non-essential loads is also incorporated. 

Additionally, power-up ‘on-switching’ and a 

time-limited ‘on-retriggering’ of essential 

spacecraft equipment loads, in order to 

restore DC power for a limited number of 

‘on-retry’ attempts, is included. The PDU 

weighs 9.7 kg. 

The Pyrotechnic Release Unit (PRU) 

The PRU conditions and safely distributes the 

energy to ignite either singly, or a maximum 

of three simultaneously, of the 

32 pyrotechnic initiators used within the 

satellite. The PRU is powered directly from 

the two spacecraft batteries and supplies a 

pulse current of 5 A amplitude and 25 ms 

duration. On-ground and launcher safety 

requirements have imposed a minimum of 

52 

three levels of protection-inhibits between 

the power source and the initiators to 

ensure that inadvertent firing of these safetycritical 

devices cannot occur. This 

requirement is fulfilled by series relays and 

current limiters placed between the batteries 

and the initiators. The PRU weighs 5 kg. 

5.5 Data Handling & 

Onboard Software 

Data Handling Subsystem (DHSS) 

The MSG Data Handling Subsystem (DHSS) 

consists of three physical units: the Central 

Data Management Unit (CDMU), and the 

two Remote Terminal Units (RTUs). The 

three units are interconnected via the serial 

standard OBDH data bus. One RTU is 

located on the spacecraft’s main platform 

and monitors the equipment mounted 

there. The other is located on the upper 

platform and monitors the MCP subsystem. 

The two RTUs are identical except for their 

OBDH bus terminal addresses, which can 

be set by external address plugs. 

Also connected to the OBDH bus is the

FCU of the SEVIRI subsystem. The FCU 

incorporates a dedicated OBDH interface, 

the Remote Terminal Interface (RTI). 

The CDMU is master on the OBDH bus, and 

controls all traffic on the bus. Commands 

and acquisitions are thus sent out from the 

CDMU to the different subsystems of the 

satellite via the RTUs, or via the SEVIRI FCU. 

Characteristics of the Data-Handling Subsystem (DHSS) 

The CDMU is also equipped with a programmable 

Central Reconfiguration Module 

(CRM), which can reconfigure the DHSS 

(including the CDMU) upon reception of 

several different alarm signals. Most of these 

alarms are generated by the CDMU itself 

(e.g. on detection of a non-correctable 

memory error or a memory-protection 

violation), but there is also one external 

CDMU RTU UP RTU MP 

Dimensions: 

Length 340 mm 250 mm 250 mm 

Width (depth) 234 mm 234 mm 234 mm 

Height 287.5 mm 203 mm 203 mm 

Mass 11.3 kg 7.3 kg 7.3 kg 

Mean Power (typical value) 19 W 7.7 W 7.7 W 

Peak Power 34.5 W 19.7 W 19.7 W 

(worst case, over 1 ms) 

Reliability 0.990 0.997 0.997 

Subsystem reliability (with CRM) 0.975 

53 

Architecture of the 

Data-Handling 

Subsystem (DHSS)

system alarm relating to the satellite ‘safe 

mode’. 

The DHSS provides the following basic 

functions: 

• Decodes and distribute telecommands 

through a hot-redundant telecommand 

(TC) chain using TC packets formatted 

according to the ESA Packet 

Telecommand Standard. All TCs, except 

high-priority TCs (handled by TC decoder 

hardware), are forwarded directly to the 

onboard software. 

• Acquires and encodes telemetry data 

(S-band) at a rate of 8192 bps, including 

coding, formatted according to the ESA 

Packet Telemetry Standard. 

• Acquires and encodes payload data 

(L-band) at a speed of 2 x 3.75 Mbps, 

including coding. Three Virtual Channels 

are used: on VC0 and VC1, packet 

headers are generated by Basic Software 

(BSW), while on VC7, all telemetry 

packets are generated by Application 

Software (ASW), except for idle packets. 

• Provides an On-Board Time (OBT) 

function using a high-precision TCX 

oscillator. 

• Distributes a set of dedicated clocks 

including: AOCE clock, CACE clock, MCP 

clock, and Payload (SEVIRI/GERB) Master 

Clocks. 

• Controls the reading out of raw data 

from the payload (SEVIRI and GERB) 

using CTS/RTS hand-shaking signals. 

• Provides a processing capability for ASW 

control functions. 

• Provides command and monitoring 

capabilities for other subsystems 

through RTU input/output channels. 

54 

Onboard Software 

All spacecraft command and control, all 

autonomous functions like onboard failure 

handling and thermal control, as well as all 

telemetry acquisition and reporting and all 

but the most basic telecommanding is 

centralised in the MSG onboard software. 

The fast rotation of the satellite imposes 

special requirements on the exact 

synchronisation of the payload data 

acquisition, which directly influences the 

quality of the image. The overall mission 

requirement of 24 h autonomy in orbit 

requires a relatively comprehensive onboard 

failure detection, isolation and recovery 

(FDIR) setup. These functions, together with 

a telemetry and telecommand 

implementation, which fully complies with 

the ESA Packet Telemetry/Telecommand 

Standards, and the higher-level, applicationoriented 

requirements defined in the ESA 

Packet Utilisation Standard, are among the 

most demanding and drive the software 

design to a large extent. 

Since several of these tasks are 

asynchronous in nature, use of a 

conventional scheduler-based operating 

system, which activates tasks periodically in 

a fixed sequence, is not possible. Instead, 

the software is based on a pre-emptive 

multitasking kernel, which supports 

asynchronous task activation. Abandoning 

the relative simplicity of a scheduler-based 

system comes at the expense of increased 

software complexity and the need for 

special software verification and testing 

efforts, which led in turn to an extended 

and comprehensive software verification 

and validation programme.

The software is split into two major blocks, 

reflecting the difference between a lowerlevel 

hardware-oriented operating-system 

kernel, the Basic Software (BSW), and 

higher-level application-oriented functions, 

which together form the Application 

Software (ASW). The complete onboard 

software suite runs on a 31750 

microprocessor in the Central Data 

Management Unit (CDMU). The software 

was written in ADA, with only very limited 

use of Assembler code for time-critical 

procedures. 

The onboard software is stored in 56 kW of 

ROM and copied into 96 kW of RAM upon 

initialisation. The memory margins are 11% 

and 20%, respectively. r 

0000h 0000h 

Code 

8000h 

B000h B000h 

Unused E000h 

FFFFh 0000h 

Logical RAM 

Operand Area 

64 Kw 

0000h 

8000h 

B000h 

E000h 

FFFFh 

Logical RAM 

Instruction Area 

64 Kw 

Data 

Unused 

Constants 

Free RAM 

17FFFh 

Physical RAM 

96 Kw 

Code 

Constants 

Free RAM 

Data 

55 

The onboard 

hardware/software 

context 

Memory map of MSG’s 

onboard software 

Instruction 

Area 

64 Kw 

Operand Area 

64 Kw 

HW 

Protected 

Area 

exact 

PROM 

copy 

56 Kw 

(code and 

constants)

Meteosat Second Generation - ESA

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