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Mars ExpressMars ExpressPlanned achievements: first European Mars orbiter & landerLaunch date: planned for 1 June 2003 (11-day window), Mars arrival

26 December 2003Mission end: orbiter after 1 martian year (687 Earth days), 1-year extension

possible; lander 180 Earth days Launch vehicle/site: Soyuz-Fregat from Baikonur Cosmodrome, KazakhstanLaunch mass: 1120 kg (science payload 113 kg, lander 60 kg)Orbit: heliocentric, followed by 259x11560 km, 86.35°, 7.5 h about Mars for first

440 days, followed by 298x10107 km, 86.35°, 6.7 hPrincipal contractors: Astrium SAS (orbiter), Astrium UK (lander), Martin-Baker

Aircraft Co (EDLS). Phase-B January-November 1999, Phase-C/D January2000 - November 2002

The Mars Express orbiter andBeagle 2 lander have key roles in theinternational exploration programmeplanned for the Red Planet over thenext two decades.

Some of the orbiter’s instrumentswere originally developed for Russia’sill-fated Mars-96 mission. Nowupgraded, they will provide remotesensing of the atmosphere, groundand up to 5 km below the surface.The information will help to answermany outstanding questions aboutMars, such as what happened to thewater that once flowed freely, and didlife ever evolve?

Beagle 2 will be the first lander sinceNASA’s two Viking probes in 1976 tolook specifically for evidence of pastor present life. No other Mars probeis making exobiology so central to itsmission.

Mars Express was conceived in 1997as the first Flexible low-cost missionin the Horizons 2000 programme. Itwill cost the Agency no more than€150 million (1996 rates) – only abouta third of the cost of similar previousmissions. Despite that modest level,however, its future hung in thebalance because of the steady erosionof ESA’s science budget since 1995.In November 1998, the ScienceProgramme Comittee approved it on

the basis that it did not affectmissions already selected,particularly Herschel and Planck.Following the Ministerial Council of11/12 May 1999 approving thefunding, the SPC gave the finalgo-ahead on 19 May 1999.

ESA is funding the orbiter, launchand operations. The scienceinstruments are being providedseparately by their home institutes;the lander is a cooperative venture by

http://sci.esa.int/marsexpress/

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The Beagle 2 lander is ejected from theorbiter. Bottom left: Mars Express in launchconfiguration on the Fregat upper stage.

ESA and the UK Beagle 2 consortium.The science Announcement ofOpportunity was released inDecember 1997 and 29 proposalswere received by the 24 February1998 deadline, including three for alander, which was treated as aninstrument. The SPC made theselection at the end of May 1998.Beagle 2 cost is €50 million, of whichESA agreed in 2000 to contribute€24 million in return for managementoversight and increased return to theEuropean scientific community.

The spacecraft is being builtunusually quickly to meet the tight11-day launch window during theparticularly favourable Marsopportunity of 2003. Savings arebeing made by reusing existinghardware, adopting new managementpractices, shortening the time fromoriginal concept to launch, andprocuring the most cost-effectivelauncher available. Maximum use isbeing made of off-the-shelf andRosetta technology – 65% of thehardware is at least partially derivedfrom the Rosetta cometary missionthat will also depart in 2003.

ESA is delegating tasks to AstriumSAS in Toulouse (F) that previouslywould have been performed by theproject team at ESTEC. In particular,Astrium is managing the

orbiter/payload and orbiter/launchertechnical interfaces. The period fromconcept to awarding the design anddevelopment contract was cut fromabout 5 years to little more than1 year. Astrium SAS won the€60 million fixed-price prime contractin December 1998 in competitionwith consortia led byAlenia/Aerospatiale and Dornier. ThePhase-B/C/D design anddevelopment phase will take less than4 years, compared with up to 6 yearsfor previous similar missions.

Recent missions have raised manyquestions about Mars. What forcescreated the spectacular landscapefeatures? When did they stop – or arethey still active? Was early Marsreally warm and wet? If so, where didthe water and atmosphere go? Did lifeevolve there? And is primitive life stillthriving, perhaps in undergroundaquifers? Mars Express will help toprovide answers by mapping thesubsurface, surface, atmosphere andionosphere from orbit and conducting

http://www.beagle2.com

Mars Express Science Goals

– image the entire surface at high

resolution (10 m/pixel) and selected areas

at super-high resolution (2 m/pixel)

– map the mineral composition of the

surface at 100 m resolution

– map the composition of the atmosphere

and determine its global circulation

– determine the structure of the subsurface

down to a few kilometres

– determine the effect of the atmosphere on

the surface

– determine the interaction of the

atmosphere with the solar wind

Beagle 2 lander:

– determine the geology and mineral

composition of the landing site

– search for life signatures (exobiology)

– study the weather and climate

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observations and experiments on thesurface. Investigations will alsoprovide clues as to why the north isso smooth while the south is sorugged, how the Tharsis and Elysiummounds were raised and whetherthere are still active volcanoes. Notonly does Mars have the largestvolcanoes and deepest canyons in theSolar System, it also shows evidencefor the most catastrophic floods.Large channels carved by these floodsdrain into the northern plains,lending support for the existence ofan ancient ocean over most of thenorthern hemisphere. Valley networksthat criss-cross the southernhighlands were also probably formedby water. And many craters,especially at high latitudes, aresurrounded by fluidised ejecta. Thissuggests there was undergroundwater or ice at the time of impact,and possibly more recently.

If water was largely responsible forthese features, however, it has longsince disappeared: most of theevidence is more than 3800 millionyears old. Today, atmosphericpressure at ground level is only about1% that on Earth. So where did thegases and water go and why? Each ofthe orbiter’s seven instruments willcontribute towards the answer.

The water could have been lost tospace and/or trapped underground.Four orbiter experiments (ASPERA,SPICAM, PFS, MaRS) will observe theatmosphere and reveal processes bywhich water vapour and otheratmospheric gases could haveescaped into space. Two (HRSC,OMEGA) will examine the surface andin the process add to knowledgeabout where water may once haveexisted and where it could still lieunderground. One (MARSIS) will lookfor underground water and ice. Thisis the first time that a ground-penetrating radar has been used inspace.

Beagle 2 will look for signatures of lifeon Mars, whether long-dead or still-living, by measuring the 12C/13C ratioin the rock. On Earth, manybiological processes favour 12C, so ahigh ratio is taken as evidence of lifeand has been found in rocks up to4000 million years old, even wheregeological processing has occurred.The hope is that the same occurredon Mars.

1 MARSIS2 HRSC3 OMEGA4 SPICAM5 PFS6 ASPERA7 MaRS

6

1

23

5

4

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On Earth, life produces anothersignature: methane. This gas has avery short lifetime on Mars because ofthe oxidising nature of theatmosphere, so its presence wouldindicate a replenishing source, whichmay be life, even if it is buried.

The only previous landers to lookdirectly for evidence of past life wereNASA’s Vikings in 1976. However,Mars’ harsh, oxidising atmospherewould almost certainly havedestroyed any such evidence on thesurface. Beagle 2 will surmount thisproblem by using a ‘mole’ to retrievesoil samples from as deep as 1 m. Itscorer and grinder will also expose theinterior of rocks for study.

The plan is for Mars Express torelease the lander 5 days before itsarrival at Mars. Beagle 2 enters theatmosphere at more than20 000 km/h protected by aheatshield. When its speed hasreduced to about 1600 km/h, theheatshield is jettisoned and theparachutes deploy. Finally, large gasbags inflate to protect it as it bouncesto a halt, 5 min after entry.Immediately, the bags are ejected, itsclamshell outer casing springs open,solar panels unfurl and camerasbegin operating. The first few daysare spent running pre-programmedsequences, imaging the site andrunning the environmental sensors,preparing for the very detailed rockand soil analysis.

Beagle 2 will land on Isidis Planitia, alarge, flat basin straddling thenorthern plains and ancienthighlands. The site (10.6ºN, 270ºW) isat the maximum latitude for the Sunto provide sufficient warmth andpower in early spring. It is not toorocky to threaten a safe landing (butenough to be interesting), has few

Most of Beagle 2’sinstruments are carriedon its arm.

Beagle 2’s 95x500 km landing ellipse in theIsidis Planitia. The sitewas selected inDecember 2000.

Sampling“Spoon’’

Stereo Camera(Left)

Rock Corer-Grinder

MössbauerSpectrometer

Microscope

Stereo Camera(Right)

X-RaySpectrometerMoleElectronics

Wind Sensor

Wide AngleMirror

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steep slopes down which the probemay have to bounce, and is not toodusty. Isidis Planitia is low enough toprovide sufficient depth ofatmosphere to allow the parachutesto brake the descent. Also, the regionis a sedimentary basin where tracesof life could have been preserved.

Most of Beagle 2’s experiments are onits PAW (Position AdjustableWorkbench) at the end of a roboticarm. They include stereo cameras,microscope, two spectrometers

(Mössbauer and X-ray), thecorer/grinder and mole. PAW alsohas a lamp for illumination, a brushand a sample scoop in case of molefailure.

Shortly after landing, the stereocamera will record panoramic shotsof the site, followed by close-upimages of near-by soil and rocks ascandidates for further analysis. Whena suitable rock has been chosen, PAWwill rotate until the grinder ispositioned to remove the weathered

Mars Express Scientific Instruments

HRSC Super/High-Resolution Stereo Colour Imager. Push-broom scanning camera, 9 CCDs. 10 m global res, 2 m res selected areas. 22 kg, 40 W, 10 Mbit/s. PI: Gerhard Neukum, Institut für Weltraumsensorik und Planetenerkundung (Berlin, D). Participating countries: D, F, RU, US, FIN, I, UK

OMEGA Visible/IR Mineralogical Mapping Spectrometer. Mineral mapping at 100 m res,Specific surface mineral and molecular phases, mapping photometric units, atmospheric particles. 0.5-5.2 µm. 30 kg, 42 W, 500 kbit/s.PI: Jean-Pierre Bibring, Institut d’Astrophysique Spatiale (Orsay, F). Participating countries: F, I, RU

PFS Planetary Fourier Spectrometer. Atmospheric composition & circulation, surface mineralogy, surface-atmosphere interactions. 1.2-45 µm. 32 kg, 35 W, 33 kbit/s.PI: Vittorio Formisano, Istituto Fisica Spazio Interplanetario (Frascati, I). Participating countries: I, RU, PL, D, F, E, US

MARSIS Subsurface Sounding Radar/Altimeter. Subsurface structure from 200 m to few km, distribution of water in upper crust, surface roughness and topography, ionosphere. 40 m-long antenna, l1.3-5.5 MHz RF waves, 17 kg, 60 W, 30 kbit/s.PI: Giovanni Picardi, Universita di Roma La Sapienza (Rome, I). Participating countries: I, US, D, CH, UK, DK

ASPERA Energetic Neutral Atoms Analyser. Upper atmosphere interaction with interplanetary medium & solar wind (lack of a magnetic field is believed to have allowed the solar windto sweep away most of Mars’ atmosphere); near-Mars plasma and neutral gas environment. 9 kg, 12 W, 6 kbit/s. PI: Rickard Lundin, Swedish Institute of Space Physics (Kiruna, S). Participating countries: S, D, UK, F, FIN, I, US, RU

SPICAM UV and IR Atmospheric Spectrometer. Water (1.30 µm), ozone (250 nm UV) & dust profiles of atmosphere by stellar occultation. 6 kg, 12 W, 6 kbit/s. PI: Jean-Loup Bertaux, Service d’Aeronomie du CNRS (Verrieres-le-Buisson, F). Participating countries: F, B, RU, US

MaRS Radio Science Experiment. Atmospheric density, P & T profiles, ionospheric electrondensity profiles (passage of radio waves through atmosphere), surface dielectric and scattering properties (reflection of radio waves) and gravity anomalies (orbital tracking). PI: Martin Paetzold, Köln University (D). Participating countries: D, F, US, A

Beagle 2 360º panoramic stereo camera & microscope (surface textures, 4 µm res; Mullard Space Science Lab.), X-ray spectrometer (in situ chemical composition; Leicester Univ.), Mössbauer spectrometer (Fe minerals, inc. carbonates, sulphates, nitrates; MPI fürChemie), Gas Analyser Package (GAP; carbon, gas isotopes, trace constituents; Open Univ.), Environmental Sensor Suite (T, P, wind speed/direction, 200-400 nm UV flux, solar protons/cosmic rays, dust impacts; 0.180 kg; Leicester Univ./Open Univ.), mole & corer/grinder (DLR Inst. Space Simulation & Polytechnic Univ. of Hong Kong). Lander surface mass 30 kg (instruments 17 kg, 100 W, 128 Kbit/s). PI: Colin Pillinger, Open Univ. (Milton Keynes, UK). Participating countries: UK, D, F, HK, CH

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GAP has 12 ovens in which samplescan be heated gradually in thepresence of oxygen. The carbondioxide generated at eachtemperature will be delivered to amass spectrometer to measure itsabundance and the 12C/13C ratio.The temperature at which thecarbon is generated will reveal itsorigins, as different carbon-bearingmaterials combust at differenttemperatures. The massspectrometer will also look formethane in atmospheric samples.

surface. PAW can then position themicroscope or spectrometers toanalyse the freshly exposed material.

When a rock looks particularlyinteresting, a sample will be drilledout with the corer and taken to thegas analysis package (GAP) inside thelander by the robotic arm. The molewill collect soil samples and deliverthem to the GAP in a similar way.The PAW will rotate until the mole ispositioned so that it can burrowunderground to collect the samples.

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A variety of tiny sensors scatteredabout the lander will measuredifferent aspects of the environment,including: atmospheric pressure, airtemperature and wind speed anddirection; UV radiation and oxidisinggases such as ozone and hydrogenperoxide; dust fall-out and thedensity and pressure of the upperatmosphere during descent. Thestereo camera will help to construct a3D model of the area within reach ofthe arm. As PAW cannot be operatedin real-time from Earth, this modelwill be used to guide the instrumentsinto position alongside target rocksand soil. The microscope will pick outfeatures a few 0.001 mm across inrock surfaces exposed by the grinder.

The Mössbauer Spectrometer willinvestigate the mineral composition ofrocks by irradiating exposed rocksurfaces and soil with gamma-raysfrom a 57Co source, and thenmeasuring the spectrum of thegamma-rays reflected back. Inparticular, the nature of iron oxidesin the pristine interior and weatheredsurface of rocks will help todetermine the oxidising nature of thepresent atmosphere.

The X-ray spectrometer willmeasure the elements inrocks by bombarding theirexposed surfaces with X-raysfrom four radioactive sources(two 55Fe and two 109Ca). GAP willestimate rock ages by measuring theratio of 40K (spectrometer) to 40Ar.

International collaboration, eitherthrough participation in instrumenthardware or scientific data analysis,is significantly enhancing themission. NASA’s major contributionto MARSIS is an example. Also,arriving at Mars at the very beginningof 2004, Japan’s Nozomi spacecraftwill follow Mars Express. Themissions are highly complementary.From its highly elliptic equatorialorbit, Nozomi will focus on the upperatmosphere as well as the interactionof the solar wind with the ionosphere.Beagle will land at about the sametime as NASA’s Mars Rover. Theagencies are arranging to enable thelanders to use each other’s orbiter asback-up for relaying data and othercommunications to Earth. MarsExpress is also requesting the use ofNASA’s Deep Space Network forcommunications with Earth duringparts of the mission. Five of the MarsExpress instruments (OMEGA, PFS,ASPERA, HRSC, SPICAM) aredescendants of instruments originallybuilt for the Russian Mars-96mission. Each of the seven orbiterinstrument teams has Russianco-investigators.

OrbiterConfiguration: box-shaped bus1.5x1.8x1.4 m of conventionalaluminium construction. Dry mass680 kg.

Attitude/orbit control: orbit correction& Mars insertion by single 400 N

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NTO/MMH thrusters,attitude control by

8x10 N hydrazinethrusters (420 kg in 2 tanks

totalling 580 L) and4x12 Nms reaction wheels.

Pointing accuracy 0.15º supportedby 2 star trackers, 6 laser gyros, 2coarse Sun sensors.

Power system: twin 4-panel Si solarwings derived from Globalstartotalling 11.42 m2 provide 650 W atMars (500 W required). Supported by3x22.5 Ah Li-ion batteries.

Thermal control: aluminium/tin alloyblankets keep interior at 10-20ºC.

Communications: via 1.6 m-dia 65 WX-band HGA to 34 m New Norcia,Perth ground station at up to230 kbit/s from 12 Gbit SSR. UHFantenna receives Beagle 2 data.Processed data will be placed inpublic archive at ESTEC after6 months. Controlled from ESOC.

Beagle 260 kg (30 kg on surface, 17 kgscience instruments). Lander 66 cm-dia, 22 cm-high, primary structurecarbon-fibre skin on aluminiumhoneycomb. Powered up shortlybefore ejection from orbiter, 5 daysout from Mars. Comprises lander andEntry, Descent & Landing System(EDLS), drawing heavily on Huygensheritage. EDLS provides a frontshield/aeroshell and backcover/bioshield. Mortar fires throughpatch in back shield to deploy 3.2 m-dia drogue ’chute and then 7.5 m-diamain ’chute; front shield is releasedpyrotechnically. Three 2 m-dia gas-bags inflated. On contact, ’chutereleased for lander to bounce away.Coming to rest, a lace is cut for thebags to open. Beagle’s clam-shell lidopens to begin the science phase.

Most science instruments mountedon arm with 75 cm reach: stereocamera, microscope, twospectrometers (Mössbauer and X-ray),lamp, mole and corer/grinder.

DLR’s Pluto (planetary undersurfacetool) mole can crawl 1 cm in 6 s byusing spring compression to propel adrive mass. Collects sample in tipcavity, wound back in from up to 3 mby power cable for sample delivery toinstruments. Grinder/corer exposesfresh rock surfaces and can drill1 cm deep for 2 mm-dia 60 mgsample.

Power provided by GaAs arraytotalling 1 m2 on 5 panels, supportedby 200 Wh Li-ion battery. 128 Kbit/sdata link by 400 MHz UHF 5 Wtransmitter via patch antennas toorbiter. Goal for surface operations is180 Earth days.

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CryoSatCryoSatPlanned achievements: first thickness-change measurements of Arctic iceLaunch date: planned for April 2004Mission end: after 3.5 years Launch vehicle/site: to be selected (Rockot/Dnepr-class)Launch mass: about 737 kg (62 kg SIRAL)Orbit: planned 717 km circular, 92º, 369-day repeat, 30-day subcycle (validation

phase 711 km, 3-day repeat)Principal contractors: Astrium GmbH (prime, system, platform, AIV)), Alcatel Space

Industries (SIRAL); Phase-A Feburary - July 2000; Phase-B July 2000 -January 2001; System Design Review February 2001; Phase-C/D July 2001 -September 2003; Critical Design Review April 2002

Cryosat will monitor changes in thethickness of the polar ice sheets andof floating sea ice. TheAnnouncement of Opportunity for thefirst Earth Explorer Opportunitymission in ESA’s Living Planetprogramme was released on 30 June1998; 27 proposals were received byclosure on 2 December 1998. On 27May 1999, ESA approved Cryosat. Asan Opportunity mission, it isdedicated to research andheaded by a Lead Investigator(LI): Prof. Duncan Winghamof University College London(UK). The LI is responsiblefor overall mission design,implementation and dataprocessing; ESA has overallmission responsibility

Rising temperatures meanthat we face the widespreaddisappearance over the next80 years of the ice coveringthe Arctic Ocean. The effectswill be profound not only inthe Arctic. Warm wintertemperatures in Europe resultfrom ocean currents that areaffected by fresh water fromprecipitation and Arctic ice meltwater,and both may increase in a warmingclimate.

Cryosat will measure variations in thethickness of the Arctic sea ice andthe ice sheet, ice caps and glaciersthat ring the Arctic Ocean. The team

is adapting existing Europeanaltimeter flight hardware to usesynthetic aperture andinterferometric techniques. In theSynthetic Aperture Radar (SAR)mode, Cryosat’s radar sweeps acrossthe groundtrack, repeatedly coveringany one location as it moves along itsorbit. The data from all beams arecombined to give a single height

measurement for that location.Later measurements will show if

there has been a change inheight. Of course, thisrequires that CryoSat’s orbitis known accurately. Wherehigher resolution is required,

over the steep edges of icesheets, the second radar isadded to create aninterferometer across the trackfor the ‘SARin’ (SARinterferometer) mode.

Sea-ice thickness plays acentral role in Arctic climate:

it limits how much the winterArctic atmosphere benefitsfrom heat stored in the oceanthe previous summer. Heat flux

of more than 1.5 kW/m2 from theopen ocean may be reduced by afactor of 10-100 by ice. Secondly,fluctuations in sea-ice mass affecthow ocean circulation is modified byfresh water. Presently, half of thefresh water flowing into theGreenland Sea – some 2000 Gt/yr –comes from the wind-driven ice floes

The UK Hadley Centre predicts a thinning ofArctic ice from 5 m to

1 m if the carbon dioxidelevel quadruples.

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from the Arctic Ocean. Finally,sea-ice thickness is very sensitive

to ice thermodynamics, how icedeforms under stress, and heat fromthe air and ocean – 10 W/m2 will meltaround 1 m of ice in a year.

It is possible that an irreversiblechange in Arctic sea-ice mass isalready underway. Existing thicknessmeasurements already suggest animportant trend in Arctic climate –but it is equally possible that this ismerely an unremarkable short-termchange. The measurements are stillscattered too thinly in space and timeto tell. Sea-ice thickness haspreviously been measured by drilling,by sonar observations of ice draftfrom submarines operating beneaththe pack ice, or in one case, from amoored sonar array across the FramStrait. Measurements haveaccumulated over the years, but arerarely of the same locations. Cryosat,by repeatedly sampling 70% of icefloes, will provide an authoritativeview of the fluctuations.

Cryosat will do more than observe thefloating ice of the Arctic Ocean.Obvious sources for the watercausing the 18 cm rise in sea levellast century are the ice sheets andglaciers on land. Observations fromERS, Seasat and Geosat indicate thatthe great central plateaus of theAntarctic and Greenland ice sheetsare stable so, if these ice sheets arecontributing to the rising sea level,the changes must be happening attheir edges. The improvement inresolution from Cryosat’s radar,coupled with its interferometriccapability, will make continuousmeasurements of the ice sheetmargins and smaller ice caps possiblefor the first time.

These measurements translate intothe need to measure thicknesschanges in Arctic sea ice of 105 km2

with an accuracy of 3.5 cm/yr

http://www.estec.esa.nl/explorer

(1.6 cm/yr will be achieved); in ice sheetscovering 104 km2 of 8.3 cm/yr (3.3 cm/yr willbe achieved); and in ice sheets of 13.8x106 km2

of 0.76 cm/yr (equivalent to a loss of 92 Gt ofwater each year; 0.17 cm/yr will be achieved).

Satellite configuration: total length 4.6 m, width2.34 m. Box-shaped aluminium bus withextended nose for SIRAL. End/side platescomplete main body and support solar array;nadir face as radiator. Controlled by ERC-32single-chip processor via MIL-1553B bus.

Attitude/orbit control: 3-axis nadir pointing to0.2/0.2/0.25º roll/pitch/yaw by three 30 Am2

magnetorquers supported by redundant sets ofeight 10 mN cold-gas nitrogen blowdownthrusters (plus four 40 mN for orbit adjust);32 kg nitrogen in single central sphere. Attitudedetermination by 3 star trackers, 3 fluxgatemagnetometers, and Earth/Sun sensors.Precise orbit determination by DORIS (5 cmradial accuracy; 30 cm realtime), LaserRetroreflectors (cm accuracy) and S-bandtransponder.

Power system: two GaAs panels totalling 9.4 m2

on upper bus provide 525 Wh (SIRAL requires132 W; 99 W in low-res mode). Supported by60 Ah Li-ion battery.

Communications: controlled from ESOC viaKiruna (S). 320 Gbit downlink each day from128 Gbit solid-state mass recorder at100 Mbit/s QSPK 25 W 8.100 GHz. 2 kbit/s TCat 2026.7542 MHz. 4 kbit/s TM at 2201 MHz.

CryoSat Payload

SIRAL

Cassegrain antenna emits 44.8 µs Ku-band pulses, bandwidth320 MHz. Three operating modes: Low-Resolution Mode (LRM) 1pulse in 44.8 µs burst, 51 kbit/s; SAR mode 64 pulses in 3.8 msburst, 12 Mbit/s, for ice floes & ice sheet interiors at < 1 km res;SARin mode 64 pulses in 3.8 ms burst, received by bothantennas, 2x12 Mbit/s, for ice sheet edges at 250-300 m res.

DORIS

Doppler Orbitography & Radio-positioning Integrated by Satellitedetermines the orbit with 5 cm accuracy. It receives 2.03625 GHz& 401.25 MHz signals from ground beacons and measures theDoppler shift every 7-10 s. 91 kg, 16.7 kbit/s, 42 W.

LRR

Cluster of 9 laser reflectors provide orbit determination with cm-accuracy using laser ground stations. Backup to DORIS.

CryoSat Phase-B design.

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AATVTVPlanned achievements: ISS resupply/reboostLaunch date: first planned for September 2004, then every 15 monthsMission end: nominally after 6 monthsLaunch vehicle/site: Ariane-5ESV from Kourou, French GuianaLaunch mass: 20 500 kg (7372 kg payload)Orbit: as ISS (typically 400 km, 51.6°), via 300x300 kmPrincipal contractors: EADS-Launch Vehicles (development), Astrium GmbH

(integration, Propulsion & Reboost Subsystem), Alenia Spazio (Cargo Carrier),Astrium SAS (avionics, software), Oerlikon Contraves (structure), Alcatel BellTelephon (EGSE)

In combination with the Ariane-5launcher, the Automated TransferVehicle (ATV) will enable Europe totransport supplies to theInternational Space Station. It willdock with Russia’s Zvezda moduleafter a 2-day autonomous flight usingits own guidance, propulsion anddocking systems. Its 7.4 t payloadwill include scientific equipment,general supplies, water, oxygen andpropellant. Up to 4 t can bepropellant for ATV’s own engines toreboost the Station at regularintervals to combat atmospheric drag.Up to 860 kg of refuelling propellantcan be transferred via Zvezda toZarya for Station attitude and orbitcontrol. Up to 5.5 t of dry cargo canbe carried in the pressurisedcompartment.

ATV offers about four times thepayload capability of Russia’sProgress ferry. Without ATV, onlyProgress could reboost the Station.Both technically and politically, it isessential that the Station can call onat least two independent systems.

An ATV will be launched on averageevery 15 months, paying Europe’s8.3% contribution in kind to theStation’s common operating costs. Itcan remain docked for up to6 months, during which time it willbe loaded with Station waste beforeundocking and flying into Earth’satmosphere to burn up.

Following launch from the Ariane-5complex in French Guiana, themission will be controlled from theATV Control Centre in Toulouse (F).Docking manoeuvres will becoordinated with NASA’s SpaceStation Control Center in Houstonand with Russia’s control centre nearMoscow, which oversees all theStation’s Russian modules.

ATV’s docking mechanism is beingprovided by Russia in exchange forESA’s Data Management System(DMS-R) for Zvezda. A similar DMS isbeing used in Columbus and ATV.The docking system has long beenused on Russia’s stations and Soyuzand Progress craft. A probe engagesthe receptacle on Zvezda and is

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slowly retracted until the 1.3 m-diafaces and their electrical andhydraulic connectors mate. Eighthooks on each face are closed tocomplete hard docking. Zvezda’s80 cm-dia hatch is opened by thecrew and a long tool is used tounlock ATV’s hatch. Finally, 16clamps are installed for rigidityacross the docking collars.

ATV development was confirmed atthe October 1995 ESA MinisterialCouncil meeting in Toulouse.Phase-B2 began in July 1996. The€408.3 million (1997 conditions)Phase-C/D fixed-price contract wassigned with EADS-Launch Vehicleson 25 November 1998. PDR wascompleted in December 2000; CDR isdue in 2003. ESA and Arianespace inJune 2000 signed a €1 billioncontract to launch nine ATVs over aperiod of 10 years. They will beproduced and operated by industryunder a single contract(encompassing the launch contract)to be awarded in 2001.

Ariane-5 injects ATV into a300x300 km, 51.6° transfer path.Orbit circularisation and phasing itwith the Station then take about50 h. At the first apogee, ATV raisesperigee to 400 km to stabilise theorbit. All ATV operations aremonitored from Toulouse via NASA'sTracking & Data Relay Satellite(TDRS) system. Following the perigee-raising manoeuvre, a series ofreconfiguration and check-outoperations is performed, notably solar

array deployment. Phasingmanoeuvres then bring ATV to theStation’s altitude. About 90 minbefore it enters the approachellipsoid, integrated operations beginand mission authority is transferredto the Mission Control Center inMoscow. Beginning at about 30 km,ATV performs final approach anddocking manoeuvres automaticallyover a period of 5 h, with eitherautomatic or manual capability fromthe Station crew to trigger a collisionavoidance manoeuvre. On firstcontact, ATV thrusts to ensurecapture and to trigger the automaticdocking sequence.

After docking, the hatch remainsopen unless it is closed to minimisethe power required from the Station.The crew manually unloads cargofrom the pressurised compartmentwhile ATV is dormant. Dry cargo iscarried in a shirtsleeve environment.

Station refuelling is powered and

ATV Capacities

Launch mass: 20500 kgCargo: 7372 kg

dry cargo: 1500-5500 kgwater: 0-840 kggas (O2, N2, air): 0-100 kgRefuelling propellant: 0-860 kg

(306/554 kg MMH/MONReboost propellant: 0-4000 kg

Dry mass: 10720 kg (inc. 7% margin)Spacecraft dry: 5127 kgCargo Carrier dry: 3455 kg (cargo

hardware 1437 kg)Consumables (propellant/He): 2408 kg

Right: ATV’s launch on an Ariane-5. Bottomleft: ATV about to dock with Zvezda.

(ESA/D. Ducros)

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ATV’s four 490 N main engines fire to boost theStation’s altitude. (ESA/D. Ducros)

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controlled (integrity checks, lineventing, fluid transfer and linepurging) by the Station throughconnectors in the docking face. ATVis reactivated during the attitudecontrol and reboost.

After undocking, ATV automaticallymanoeuvres for deorbiting andcontrolled reentry in the Earth'satmosphere. Carrying up to 5.5 t ofwaste from the Station, ATV will besafely consumed during reentry.

Satellite configuration: flared cylinder,10.27 m long, max 4.51 m dia,22.315 m across solar wings.Propulsion Module, Avionics Moduleand Separation and DistancingModule, all of aluminium alloy, with aMeteoroid and Debris ProtectionSystem mounted on the primarystructure.

Attitude/orbit control: Propulsion &Reboost Subsystem uses four 490 Nmain engines (SI >310 s) and 20220 N attitude thrusters (minimumimpulse bit < 5 Ns). Mixed oxides ofnitrogen (MON, oxidiser) andmonomethyl hydrazine (MMH, fuel)stored in eight identical 1 m-diatitanium tanks, pressurised withhelium stored in two high-pressuretanks regulated to 20 bar. Tanks canhold up to 6760 kg of propellant formain navigation and reboostrequirements. GNC calculations arebased on two GPS receivers forposition, four gyros and two Earthsensors for attitude, and tworendezvous sensors for final approachand docking.

Power system: four wings (each offour 1.158x1.820 m panels) in X-configuration, using mix of GaAs andhigh-efficiency Si cells. 3860 W BOLin Sun-pointing mode; 3800 W EOL.Supported by NiCd batteries during

eclipse periods; non-rechargeablebatteries are used during some flightphases. Attached to ISS, ATV indormant mode requires up to about600 W from the Station.

Communications: via two redundantS-band systems – a TDRS link toground control and a proximity linkto the Station.

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Planned achievements: key Space Station laboratoryLaunch date: planned for October 2004Mission end: nominally after 10 yearsLaunch vehicle/site: NASA Space Shuttle from Kennedy Space Center, FloridaLaunch mass: 12700 kg (including 2500 kg payload)Orbit: as ISS (about 400 km, 51.6°)Principal contractors: Astrium GmbH (ex-DASA-RI: prime; ex-Dornier: ECLS),

Alenia Spazio (structure, ECLS, thermal control, pre-integration), Astrium SAS(ex-MMS-F, DMS), Aerospatiale (MDPS); Phase-C/D began January 1996; PDRDecember 1997; CDR October-December 2000

The Columbus laboratory is thecornerstone of Europe’s participationin the International Space Station(ISS). Columbus will provide Europewith experience of continuousexploitation of an orbital facility,operated from its own ground controlfacility in Oberpfaffenhofen,Germany. In this pressurisedlaboratory, European astronauts andtheir international counterparts willwork in a comfortable shirtsleeveenvironment. This state-of-the-artworkplace will support the mostsophisticated research inweightlessness for 10 years or more.Columbus is a general-purposelaboratory, accommodatingastronauts and experiments studyinglife sciences, materials processes,technology development, fluidsciences, fundamental physics andother disciplines.

In October 1995, the ESA MinisterialCouncil meeting in Toulouseapproved the programme ‘EuropeanParticipation in the InternationalSpace Station Alpha’ – 10 years afterthe authorisation of studies andPhase-B work by the MinisterialCouncil in Rome in 1985. During thatperiod, a variety of space elementswas studied in parallel with theAriane-5 launcher and Hermes mini-shuttle programmes, which all led tointegrated European scenarios that

were clearly unaffordable. In late1994, a series of dramatic cutbacksbegan, which underwent frequentiterations with ESA Member Statesand the ISS Partners. The processculminated in a package worth some€2.6 billion being approved inToulouse, including what was thencalled the Columbus Orbital Facility(COF).

Soon after that Toulouse conference,a contract was signed with primecontractor Daimler Benz Aerospace(which then became DaimlerChryslerAerospace or DASA, and nowAstrium GmbH), in March 1996, at afixed price of €658 million, the largestsingle contract ever awarded by theAgency at the time.

The completed primary structure at Alenia Spazio.The circular holes will carry pressure-relief valves.

ColumbusColumbus

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Columbus will be attached to Node-2 onthe Station’s starboardleading edge.(ESA/D. Ducros)Below: Columbus rackpositions. The finallocations of ESA’s multi-user facilities shown inA1-A4 have yet to befinalised.

ESA’s ISS Exploitation Programme for2000-2013 was presented at theMinisterial Council in Brussels inMay 1999, where the overallprogramme approach and the initialphase of activities were approved. Theprogramme is being carried out in5-year phases, each with a 3-yearfirm commitment and a 2-yearprovisional commitment. Theprogramme covers the SystemOperations costs in Europe, the ESAshare (8.3%) of the overall StationCommon Operations costs and theEuropean Utilisation-related costs.The average yearly cost over thewhole period will be about€280 million (1998 rates).

Under an agreement with the ItalianSpace Agency (ASI), ESA provided theColumbus-derived ECLS for ASI’sthree Multi-Purpose LogisticsModules (MPLMs, first flown March2001) in exchange for the Columbusprimary structure, derived from thatof MPLM. The estimated saving toeach partner was €25 million.

Although it is the Station’s smallestlaboratory module, Columbus offersthe same payload volume, power,data retrieval etc. as the others. Thisis achieved by careful use of theavailable volume and sometimes bycompromising crew access andmaintainability in favour of payload

accommodation. A significant benefitof this cost-saving design is thatColumbus can be launched alreadyoutfitted with 2500 kg of payloadracks.

Columbus will be delivered to the ISSby the Space Shuttle Orbiter, carriedin the spaceplane’s cargo bay via itstrunnions and keel fitting. A Station-common grapple fixture will allow theSpace Station Remote ManipulatorSystem (SSRMS) to lift it out of theOrbiter and transport it to its finaldestination on Node-2.

In exchange for NASA launchingColumbus aboard the Space Shuttle,ESA is providing two of the Station’sthree Nodes, the Cryogenic Freezerand the Crew Refrigerator/Freezer.

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ESA has entrusted responsibility fordeveloping Nodes-2 and -3, which usethe same structural concept asColumbus, to ASI. Once Columbus isoperational, an ESA astronaut willwork on average aboard the Stationfor 3 months every 8 months.

Inside Columbus, the InternationalStandard Payload Racks (ISPRs) arearranged around the circumference ofthe cylindrical section, in a 1 gconfiguration, to provide a workingenvironment for up to threeastronauts. A total of 16 racks can becarried in four segments of four rackseach. Three in the floor containsystems; D1 contains environmentalequipment (water pumps, condensingheat exchanger, water separator,sensors); D2/D3 are devoted toavionics. Ten racks are fully outfittedwith resources for payloads and threeare for stowage. NASA has the rights

to five of the payload racks. ISPRshave standardised interfaces thatallow operation in any non-RussianISS module.

Columbus will be launched with aninternal payload of up to 2500 kg infive racks: ESA’s Biolab, FluidScience Lab (FSL), EuropeanPhysiology Modules (EPM), EuropeanDrawer Rack (EDR) and a stowagerack. The other five will be deliveredby MPLM, the only carrier that candeliver and return whole racks.

European rack payloads areconnected to dedicated busses fordata to be routed, via the ISS datatransfer system, directly to theEuropean Control Centre and thenceto the individual users. For NASArack payloads, interfaces to the NASAData Management System areprovided.

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Columbus also offers an ExternalPayload Facility (EPF), added in 1997when it became clear that leasing USlocations would be too expensive andthat Japan’s sites were full. Externalpayloads will be installed on the fourpositions on-orbit using the SSRMS.EPF provides the same mechanicalinterfaces as NASA’s standardExpress Pallets for external payloadson the ISS Truss.

The central area of the starboardcone carries system equipment thatrequires undisturbed crew viewingand handling access, such as videomonitors and cameras, switchingpanels, audio terminals and fireextinguishers.

Although Columbus is notautonomous – so it has no powergeneration or attitude control, forexample – it is a manned laboratoryand therefore has an EnvironmentalControl and Life Support Subsystem(ECLSS), largely in system rack D1.

Configuration: 10.2 t without researchequipment, 12.7 t at launch, 20.1 tfully outfitted. 10.16 t payload(9.0/1.16 t internal/external). 4.2 m-dia cylinder closed with weldedendcones. 3004 kg primary structureis 2219-aluminium alloy, 3.8 mmthick, increasing to 7 mm for theendcones. Length 6.2 m. Total volume75 m3, reducing to 50 m3 with fullrack load. Passive NASA-providedCommon Berthing Mechanism (CBM)attaches it to active CBM of Node-2.Other endcone has 2.5 m-dia holeused on ground to install large itemssuch as system racks. Oncompletion, the hole is permanentlyclosed off by a bolted plate.

Power: Columbus is sized to receive20 kW total (13.5 kW for payloads).

The External Payload Facility (EPF) on thestarboard endcone.(ESA/D. Ducros)

The Station’s 120 Vdc system goesthrough the Columbus PowerDistribution Unit and then, as120 Vdc or 28 Vdc, to all payloadracks, EPF locations, centre aislestandard utility panels andsubsystems.

Thermal control: active control via awater loop serving all payload racklocations. Connected to the ISScentralised heat rejection system viainterloop heat exchangers. Inaddition, there is an air/water heatexchanger to remove condensationfrom cabin air. The module iswrapped in goldised Kapton MultiLayer Insulation to minimise overallheat leaks. A system of electricalheaters combats the extreme coldpossible at some Station attitudes.These heaters will be activated duringlaunch and the transfer from theShuttle bay, drawing on the SSRMSpower supply.

Atmosphere: the cabin is ventilated bya continuous airflow entering viaadjustable air diffusers on the upperstand-offs, sucked in from Node-2 bya fan centred below the hatch in theport cone. The air returns to the Nodefor refreshing and carbon dioxide

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removal. The crew can controltemperature (16-30ºC) and humidity,and air content is monitored forcontamination.

Communications: data rate up to43 Mbit/s available (ESA has accessright of 8.3%) through NASA Ku-bandTDRS/White Sands system; alsopossible via Japan’s JEM to Artemisto Redu ground station. System/payload control by Data ManagementSystem (DMS) using MIL-STD-1553bus and Ethernet LAN; crew accessvia laptops. Columbus Control Centreis at DLR Oberpfaffenhofen (D).

MDPS: Columbus is protected by the2 t Meteoroid and Debris ProtectionSystem of bumper panels. There aretwo general panel types: single

(1.6 mm-thick Al 6061-T6), double(2.5 mm-thick Al 6061-T6 with aseparate internal bumper of Nextel& Kevlar). The cylinder carries 48panels: double on the side along thevelocity vector (+Y), and single onthe anti-velocity face (–Y). The portcone has 16 singles; the starboardcone has 16 doubles plus a singleon the central disc.

Columbus internal payloadaccommodation: payloads carried in10 ISPRs supplied with services viaColumbus. ISPR (2013 mm ht,1046 mm width, 858 mm maxdepth, empty mass 99 kg) supports704 kg in 1.2 m3. 3 kW & 6 kWversions, located in six 6 kW & four3 kW slots, with water cooling loopsized to match. 13.5 kW total to

The communications infrastructure for Columbus. ATV-CC: Automated Transfer Vehicle Control Centre (Toulouse,F), EAC: European Astronaut Centre (ESA), FRC: Facility-Responsible Centre, IGS-CN: Interconnected GroundSubnet, JSC: Johnson Space Center (NASA), MCC-H: Mission Control Center-Houston, MSFC: Marshall SpaceFlight Center (NASA), PDSS: Payload Data Service System, POIC: Payload Operations Integration Center(NASA), SSCC: Space Station Control Center, SSIPC: Space Station Inter-Process Communication, USOC: UserSupport Operations Centre.

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payloads. GN2 supply, vacuum(except rack positions O2/O1),32 Mbit/s (Columbus max43 Mbit/s), video system (A4 only),MIL-STD-1553 Columbus & Destinypayload bus & LAN. Payloads alsocarried in centre aisle, whichprovides only data & power (500 W)links.

EPF Accommodation: 4 positionseach offer payload envelope of981x1194 mm, 1393 mm ht, 227 kg,2x1.25 kW @ 120 Vdc, 32 Mbit/s,interface to Columbus/Destinypayload bus & LAN. No thermalcontrol GN2, venting. Payloadcarries integrated standard ExpressPallet Adapter (EPA) on active Flight

Releasable Attachment Mechanism(FRAM). SSRMS positions payload onEPF’s passive FRAM.

Nodes-2 & -3Nodes-2 & 3 will provide importanton-orbit resources for operating otherISS elements. In particular, Node-3will provide water processing andoxygen generation for the USsegment, avoiding sole dependence onthe Russian segment. Node-2 will bedelivered to the Kennedy SpaceCenter in September 2002, andlaunched in November 2003. Node-3delivery is scheduled for July 2003,and its launch July 2005. The Nodeshave the same basic geometry of a

Columbus will belaunched with four multi-user units installed (fromleft): EPM, FSL, Biolaband EDR.(ESA/D. Ducros)

Columbus will be launched with four ESAresearch facilities, and a fifth (MSL) will beadded later.

Biolab: biological experiments on micro-organisms, animal cells, tissue cultures, smallplants and small invertebrates in zero gravity.This is an extension – as for so many otherpayloads – of pioneering work conducted onthe Spacelab missions.

European Physiology Modules (EPM): bodyfunctions such as bone loss, circulation,respiration, organ and immune systembehaviour, and their comparison with 1 gperformance to determine how the results canbe applied to Earth-bound atrophy and age-related problems.

Fluid Science Laboratory (FSL): the complexbehaviour of fluids, the coupling between heatand mass transfer in fluids, along withresearch into combustion phenomena thatshould lead to improvements in energyproduction, propulsion efficiency andenvironmental issues.

Material Science Laboratory (MSL, to bedelivered later): solidification physics, crystalgrowth with semiconductors, measurement ofthermophysical properties and the physics ofliquid states. For example, crystal growthprocesses aimed at improving ground-basedproduction methods can be studied. In metalphysics, the influences of magnetic fields onmicrostructure can be determined.Microstructure control during solidificationcould lead to new materials with industrialapplications.

The first set of external payloads is: AtomicClock Ensemble in Space (ACES), providing anultra-accurate global time-scale, supportingprecise evaluations of relativity; Expose,mounted on the Coarse Pointing Device (CPD),for long-term studies of microbes in artificialmeteorites and different ecosystems; Sport, tomeasure polarisation of sky diffuse backgroundat 20-90 GHz; Solar (also CPD), to measure theSun’s total and spectral irradiance; EuropeanTechnology Exposure Facility (EuTEF), a widerange of on-orbit technology investigations.

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Node-2’s welded primary structure at Alenia Spazio, Turin.Cupola will provide a clear view of robot arm operations.

Node-3 internal configuration.(Alenia Spazio)Node-2 internal configuration. (Alenia Spazio)

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cylindrical pressure shell capped bytwo end-cones with axial ports. Thecylinder is a 2-bay section forhousing eight racks plus a sectionwith four radial ports.

The initial NASA concept design forNodes-2/3 was the same as that ofNode-1. However, NASA wantedlonger Nodes from Europe. Stretchingprovided additional locations forstowage. Node-3 was a Node-2 copyfor future Station use. NASA thendecided to make the stowage areaconfigurable for Crew Quarters, sothat Node-2 could provide earlyStation habitation, and most of theformer US Habitation modulefunctions such as air revitalisationand water processing were movedinto Node-3. Eventually, Node-3 wasconfigured with resources for otherattached elements: Cupola, CRV anda future Habitation module, inaddition to providing redundant portswith growth utilities for docking ofMPLMs, Shuttle or another laboratorymodule.

Each Node is 7.19 m long overall and4.48 m in diameter. Node-2 carriesfour avionics racks and four racklocations for either stowage or crewquarters. Node-3 has two avionicsracks, four for environmental controland one Waste & HygieneCompartment packaged in two racklocations. Externally, the layoutincludes almost 100 MDPS panelswith thermal blankets underneath, tominimise heat flux across the shelland to protect against meteoroids anddebris. Heat exchangers between theexternal panels and the pressureshell reject heat from Node internalequipment and attached modules.

Node-2 will be attached in front of theUS Lab, with its longitudinal axisalong the Station’s velocity vector.The forward port supports Shuttledocking, via a Pressurized MatingAdapter (PMA). The starboard sideprovides resources for Columbus, andthe port side for JEM. At the zenithposition, Node-2 will initiallyaccommodate Japan’s Experimental

Logistics Module-Pressurized Section(ELM-PS), before JEM appears, andlater the Centrifuge AccommodationModule (CAM). Finally, the nadir portwill allow temporary docking of MPLMor Japan’s HII Transfer Vehicle (HTV).

When Node-2 is delivered to theStation, its aft port will be dockedfirst to Node-1’s port side so thatShuttles can continue docking withthe US Lab’s forward port. TheStation’s arm will then move it to thefinal position.

Node-3 will be attached to Node-1’snadir, with its radial ports closer toEarth. To starboard, Node-3accommodates the CRV, while theport side is reserved for a futureHabitation module. The forwardposition includes utilities for berthingthe Cupola and is also a backuplocation for the MPLM. The aft portcan be used for temporary parking ofCupola. Nadir offers a redundantlocation for Shuttle docking, via aPMA.

Node-2 Major Capabilities– regulation and distribution of power to elements

and Node equipment (sized for 56 kW);– active thermal control of coolant water for heat

rejection from internal Node equipment andfrom attached elements;

– temperature, humidity and revitalisation controlof cabin air and air exchanged with attachedelements;

– distribution lines for cabin air sampling,oxygen, nitrogen, waste water and fuel cellwater;

– data acquisition and processing to supportpower distribution, thermal control andenvironmental control functions inside theNode, as well as data exchange between the USLab and Node-attached elements;

– audio and video links.

Node-3 Major CapabilitiesFeaturing the same basic Node-2 capabilities,

Node-3 manages less power but adds:– on-orbit air pressure and composition control,

including carbon dioxide removal;– oxygen generation, a dedicated rack also

scarred for future water generation;– waste and hygiene compartment;– urine and water processing;– controlled venting of byproducts from

environmental control;– drinking water distribution;– audio and video recording;– on-orbit reconfiguration of utilities provided to

Cupola, MPLM and TransHab.

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GalileoGalileoPlanned achievements: first European global navigation satellite system; civil

position accuracy to 5 mLaunch date: first planned for 2004 Mission end: each satellite will have a nominal lifetime of 15 yearsLaunch vehicle/site: operationally in clusters of 8 on Ariane-5ECB from Kourou,

French Guiana, replacements by vehicles such as Soyuz-Fregat Launch mass: about 650 kgOrbit: 24 000 km circular, 56°, in 3 orbital planesPrincipal contractors: to be selected (expected early 2002). Development &

Validation Phase 2001-2005 (Preliminary System Design Review March 2002;Validation Review late 2005); Deployment Phase 2006-2007; Operations Phasebeginning 2008

For the first time, the Agency hastaken a lead role in the navigationsegment. ESA and the EuropeanCommission have joined forces todesign and develop Europe’s ownsatellite navigation system, Galileo.Whereas the US GPS and RussianGlonass systems were developed formilitary purposes, the civil Galileowill offer a guaranteed service,allowing Europe to develop its ownintegrated transport system. It willalso allow the European economy tobenefit from the enormous growthexpected in value-added services andequipment for navigation systems.

Europe’s first venture into satellitenavigation is EGNOS (EuropeanGeostationary Navigation OverlayService), a system to improve thereliability of GPS and Glonass to thepoint where they can be used forsafety-critical applications, such aslanding aircraft and navigating shipsthrough narrow channels. ESAengineers first developed plans for aGPS and Glonass augmentationsystem in the late 1980s. Working inclose cooperation with the EC and theEuropean Organisation for the Safetyof Air Navigation (Eurocontrol), theAgency later adopted the plans as theEGNOS programme. When EGNOSbecomes operational in early 2004using payloads on Artemis and theInmarsat-3 satellites, it will be

Europe’s contribution to the firststage of the Global NavigationSatellite System (GNSS-1),complementing similar enhancementsin the US and Japan.

In addition, the European Unionrecognised the need for its ownindependent global satellitenavigation system. Consultationsbegan in 1994 and 4 years later plansfor a fully European system were

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drawn up. Early in 1999, the ECannounced Galileo. It will beinteroperable with GPS and Glonass,so that a user can take a positionwith the same receiver from any ofthe satellites in any system. Byoffering dual frequencies as standard,however, Galileo will deliverpositioning accuracy down to 5 m,which is unprecedented for a publicsystem. It will also guaranteeavailability of the service under allbut the most extreme circumstancesand will inform users within 6 s of afailure of any satellite. This will makeit suitable for safety-criticalapplications.

The fully-fledged service will beoperating by 2008 when 30 Galileosatellites are in position in circularorbits 24 000 km above the Earth.The first will be launched in 2004and by 2006 sufficient should be inplace to begin an initial service. Theorbits will be inclined at 56º to theequator, giving good coverage at alllatitudes. 27 satellites will beoperational and three will be activespares, ensuring that the loss of onehas no discernible effect on the user.Ground stations spread around theglobe will monitor the satellites’positions and the accuracies of theironboard clocks. The stations will beconnected to central control facilitiesin Europe via a dedicated network.

The control facilities will monitor andcontrol the constellation and computenavigation messages to transmit tothe satellites via the ground stations.The control facilities will also keepservice providers, such as providersof traffic management services,informed about the operating statusof the satellites.

ESA approved the initial studies inMay 1999 at its Ministerial CouncilMeeting in Brussels; the EC decidedon its own go-ahead in June 1999.The EC is responsible for the politicaldimension: it is undertaking studiesinto the overall system architecture,the economic benefits and the needsof users. ESA defined the Galileospace and ground segments under itsGalileoSat programme. ESA placedcontracts with European industry todevelop critical technologies, such asnavigation signal generators, poweramplifiers, antennas and highlyaccurate atomic clocks, and toprovide tools to simulate theperformance of the wholeconstellation. The studies supportedby the EC are feeding into the designof the Galileo constellation andground segment as well as thepolitical decision-making process.

Europe’s traditional space industryundertook Galileo definition studiesfrom November 1999 to February

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2001 under two contractsfinanced separately by the ECand ESA. Industries not usuallyassociated with space, such asmobile phone manufacturers andservice providers, are, in additionto their involvement in the EC-funded definition study,undertaking their own studies ofservices and applications.

The EC, ESA and private industryare expected to meet the €3250million estimated cost of Galileothrough a public/privatepartnership (PPP). The nature ofthe PPP has yet to be worked outand will depend partly on theoutcome of studies and trials todetermine the nature and size ofthe revenue stream that Galileo isexpected to deliver to industry.

Early studies suggest that Galileowill repay its initial investmenthandsomely, estimating thatequipment sales and value-addedservices will earn an extra€90 billion over 20 years. Morerecent studies yielded higherestimates by consideringpotential earnings from value-added services that combineGalileo positioning with otherservices provided by the newgeneration of mobile phones,such as internet services. So far,the EC and ESA have agreed acommon action plan, but havefinanced the industrial definitionstudies and critical technologydevelopments separately. Fundingon the EC side has come from theEuropean Union’s 5th Frameworkprogramme of research anddevelopment and on the ESA sidefrom the Agency’s GalileoSatdefinition programme (November

1999 to February 2001, costing€40 million, including technologydevelopment). Differentarrangements, however, will be inplace for the development phasewhen it begins at the end of 2001.

The development phase can bedivided into three, each of whichcould operate under differentfinancial arrangements. The firststage (development and in-orbitvalidation) will cost about €1.1 billion,financed equally by ESA and the EUfrom public money. It aims to have inorbit by the end of 2005 a handful ofsatellites for validating the Galileosystem.

At its meeting on 30 January 2001,ESA’s Navigation Programme Boardapproved the release of funding forthe Galileo Phase-B2 study, themission consolidation studies and thesupport studies on the Galileoservices and PPP consolidation tobegin. ESA plans to release thesecond tranche (€497 million) beforethe end of 2001.

On the EU side, the TransportMinisters on 5 April 2001 approvedthe release of €100 million to startdevelopment. The ministers willdecide on the release of a further€450 million in December 2001, whenthey will also approve the setting up

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of an entity to manage theprogramme. Moreover, they agreed totake the formal decision by the end of2003 on the deployment of the fullconstellation.

The €2.1 billion cost of the second(deployment) phase will be met from amixture of public and privatesources. ESA and the EU areexpected to share the public sliceequally. During this phase, which willend in 2007 when all the satellitesare launched, industry and serviceproviders will be able to developcommercial opportunities.

The third (operations) phase, startingafter 2007, will be financed by theprivate sector. By this time, thesystem will be fully operational andavailable for a wide variety ofcommercial and public service users.

The total cost for the design,development, in-orbit validation anddeployment of the first fullconstellation, including the groundsegment, EGNOS integration andapplication developments is€3250 million. The annual operatingcosts for EGNOS and Galileo areestimated at €25 million and€220 million, respectively.

Navigation payload: the navigationsignals are modulated with 500-

1000 bit/s data messagesreceived from ground stationsthrough the TT&C subsystemand stored, formatted andencoded onboard. Galileo willuse up to four L-band carriers.A highly stable onboardreference frequency of10.23 MHz is generated fromrubidium atomic frequencystandards and passivehydrogen masers operating inhot redundant, parallelconfiguration. Solid-stateamplifiers generate some 50 Woutput per signal carrier. Thenavigation antenna uses twinbeam-forming networks (one foreach band) and an array ofradiating elements to provideglobal coverage with a singlebeam. High data ratesmaximise the potential forvalue-added services such asweather alerts, accidentwarnings, traffic informationand map updates. The use of‘pilot’ signals (ranging codeswith no data messages) is beingstudied for improvingnavigation signal acquisitionunder adverse receivingconditions.

Search & Rescue (SAR) payload:the SAR payload, being definedin cooperation with COSPAS-SARSAT, receives signals fromstandard 406 MHz distressbeacons through a dedicatedantenna. The signals areamplified and transmitted at1544 MHz to SAR ControlCentres using the navigationantenna. Acknowledgementmessages are relayed back tothe beacon by integrating themin the navigation data stream.

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ERAERAPlanned achievements: first European robot arm in spaceLaunch date: no earlier than 2005Mission end: 10 yearsLaunch vehicle/site: Space Shuttle from Kennedy Space Center, Florida, USLaunch mass: 630 kg Orbit: attached to International Space Station (altitude about 400 km)Principal contractors: Fokker Space (prime), SABCA (joint subsystem), Netherlands

National Aerospace Laboratory (MPTE)

ESA’s European Robotic Arm (ERA)will play an important role inassembling and servicing theInternational Space Station (ISS). It isa cooperative venture between ESAand Rosaviakosmos, the Russianspace agency. The project began asthe Hermes Robot Arm (HERA) for theHermes mini-shuttle. When Hermeswas discontinued, studies for the armto fly on Russia’s proposed Mir-2second generation space station wereconducted between Fokker and RSC-Energia. These studies highlightedthe value of a robotic manipulator inreducing the time needed forexpensive manned activities in ahazardous environment.

Following Russia joining the ISSprogramme in 1993, the arm wasformally incorporated into thestation’s Russian Segment in July1996. It will be mounted on Russia’sScience and Power Platform (SPP),launched together on the US Shuttle.Among its first tasks is installing theSPP’s solar arrays. ESA will be readyto deliver the arm in 2002, but

launch will not be before 2005because Russian funding problemshave postponed work on the SPP.

ERA is functionally symmetrical, witheach end sporting an ‘End-Effector’that works either as a hand or as abase from which the arm can operate.There are seven joints (in order: roll,yaw, pitch, pitch, pitch, yaw, roll), ofwhich six can operate at any onetime. This configuration allows ERAto relocate itself on to differentbasepoints on the SPP, using acamera on the End-Effector to locatea basepoint accurately.

Each End-Effector includes a specialfixture to grapple and carry payloadsof up to 8 t. Through this fixture, thearm can supply power and exchange

ERA Characteristics

Total length: 11.4 mReach: 10 mMass: 630kgPayload positioning accuracy: 5 mmPayload capability: 8000 kgMaximum tip speed: 20 cm/sStiffness (fully stretched): >0.4 N/mmOperating power: 475 W avgArm booms: carbon fibre, 25 cm dia

EQMthermaltesting atESTEC

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Cosmonauts will controlERA via IMMI’s

computer-generatedviews.

data and video signals. In addition,it features a built-in IntegratedService Tool that can activate smallmechanisms in the grappledpayload. Equipped with a footrestraint, ERA can carryspacewalking cosmonauts, providingthem with a platform as they workin weightlessness.

Unusually, ERA’s main computer ismounted on the arm itself, providinga simpler control interface. It can becontrolled from inside the Station,using the Internal Man-MachineInterface (IMMI) at the Zvezdamodule’s central control post.IMMI’s synoptic display providescomputer-generated detailed andoverview pictures of ERA and itssurroundings. In addition, monitorsdisplay video images from ERA andthe Station’s external cameras. ERAcan also be operated by aspacewalker via the External Man-Machine Interface (EMMI) controlpanel. Commands are entered viatoggle switches, while LEDs displayarm status and operations progress.Both approaches offer an automaticmode (using prepared missionplans), semi-automatic mode(standard autosequences) andmanual mode (controlling theindividual joints).

Cosmonauts will train on theMission Planning and TrainingEquipment (MPTE), a realisticsimulator of the arm and itsenvironment. At its core is a fullyflight-representative ERA onboard

computer using the full flightsoftware. The MPTE, including mock-ups of IMMI and EMMI and usingtheir flight software, will be located inRussia at RSC Energia and theGagarin Cosmonaut Training Centre,and at ESTEC in the ERA SupportCentre. The Russian MPTEs will beused to train ERA’s cosmonautoperators and generate ERA flightprocedures for transmission to theSpace Station. The MPTE can alsoprovide on-line support during ERAoperations, and play back andanalyse actual operations. ESTEC’sMPTE will support these activities.

Crews aboard the Station willmaintain their expertise via a special‘Refresher Trainer’. This is a reducedERA simulation built into astandalone laptop. They can practisean entire ERA operation before it isdone for real.

There are two main developmentmodels. The Engineering/Qualification Model (EQM) was testedin November 1999 in the Large SpaceSimulator at ESTEC to check itsthermal balance. The Flight Modelunderwent EMC and vibration testingat ESTEC at the end of 2000. Thefinal acceptance review is planned for1Q 2002. Under the July 1996agreement, Russia takes ownership ofthe flight hardware once it islaunched, in exchange for which ESAwill participate in robotics activitiesaboard the Station and Agencyastronauts will be trained at theGagarin centre.

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GOCEGOCEPlanned achievements: major improvement in measurement of Earth’s gravity field

and geoid; first use of gradiometry in spaceLaunch date: planned for October 2005Mission end: nominal 20 months, extended 30 monthsLaunch vehicle/site: to be selectedLaunch mass: 980 kg (194 kg payload)Orbit: planned 250 km circular, 96.5° Sun-synchronousPrincipal contractors: Alenia Spazio (prime), Astrium GmbH (bus), Alcatel Space

Industries (gradiometer), ONERA (gradiometer accelerometers); Phase-A June1998 - June 1999; Phase-B December 2000 - December 2001; Phase-C/DJanuary 2002 - July 2005

The Gravity Field and Steady-StateOcean Circulation Mission (GOCE)will measure the Earth’s gravity fieldand geoid with unprecedentedaccuracy and resolution using a3-axis gradiometer. This will improveour understanding of the Earth’sinternal structure and provide amuch better reference for ocean andclimate studies, including sea-levelchanges and ice-sheet dynamics.

GOCE was selected in 1999 as thefirst Earth Explorer Core mission inthe Living Planet programme. Fourcandidate Core missions wereselected in November 1996 for12-month Phase-A studies, completedin June 1999. GOCE and ADM wereselected in November 1999 in order ofpriority. Alenia Spazio was selected asGOCE prime contractor in January2001.

The Earth’s gravity field is thefundamental physical force for everydynamic process on the surface andbelow. Since the beginning of thespace age, mapping the global gravityfield has been seen as a high-prioritytask. The first era lasted for about 40years, with Europe playing a leadingrole. The research was characterisedby a combination of satellite geodetic(optical, laser, Doppler) and terrestrialgravity methods. It producedsignificantly improved measurementsat the scale of several thousand km.The new era will see dedicated

satellite missions determining theglobal gravity field with consistentaccuracy and higher resolution.GOCE puts European science andtechnology in a leading position. TheGOCE-derived gravity field andassociated geoid will be the dominantreference for the following decades ofgeophysical research.

GOCE’s gradiometer will for the firsttime measure gravity gradients in alldirections. It is specifically designedfor the stationary gravity field –measuring the geoid and gravityanomalies with high accuracy (1 cmand 1 mgal, respectively) and highspatial resolution (100 km). Inparticular, it will provide:

GOCE will map Earth’s gravity field with unprecedented accuracy. The gradiometer packageis in the centre. GOCE flies along the line of themain axis, towards top right. The paired ionthrusters for combating air drag are at bottom left(purple). The red spheres are propellant tanks.

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GOCE’s gradiometer uses six paired proof masses(red) to measure variations in the gravitational field.

MLI: Multi-Layer Insulation. TMDD: Thermal andMechanical Decoupling Devices.

– new understanding of the physicsof the Earth’s interior such asgeodynamics associated with thelithosphere, mantle compositionand flow, uplifting and subductionprocesses,

– a precise estimate of the marinegeoid for the first time, which isneeded in combination withsatellite altimetry for measuringocean circulation and transport ofmass,

– estimates of the thickness of thepolar ice sheets through thecombination of bedrock topographyderived from space gravity and icesurface height measured byaltimeter satellites.

– a high-accuracy global heightreference system for datum pointlinking, which can serve as areference surface for studyingtopographic processes, includingthe evolution of ice sheets and landtopography.

Space gradiometry is themeasurement of accelerationdifferences, ideally in all threedimensions, between separated proof

masses inside a satellite. Thedifferences reflect the variousattracting masses of the Earth,ranging from mountains and valleys,via ocean ridges, subduction zonesand mantle inhomogeneities down tothe topography of the core-mantle-boundary. The technique can resolveall these features imprinted in thegravity field. Non-gravitational forceson the satellite, such as air drag,affect all the accelerometers in thesame manner and so are cancelledout by looking at the differences inaccelerations. The satellite’s rotationdoes affect the measured differences,but can be separated from thegravitational signal in the analysis.The gravitational signal is strongercloser to Earth, so an orbit as low aspossible is chosen, although thatthen requires thrusters to combat theair drag. GOCE’s cross-section is assmall as possible in order to minimisethat drag.

The gradiometer measurements aresupplemented by a GPS/Glonassreceiver that can ‘see’ up to 12satellites simultaneously. The gravity

http://www.estec.esa.nl/explorer

Why Measure the Gravity Field?

The gravity field plays a dual role inEarth sciences. Gravity anomaliesreflect mass variations inside the Earth,offering a rare window on the interior.The geoid is the shape of an ideal globalocean at rest, and it is used as thereference surface for mapping alltopographic features, whether they areon land, ice or ocean. The geoid’s shapedepends solely on Earth’s gravity field,so its accuracy benefits from improvedgravity mapping. Measuring sea-levelchanges, ocean circulation and icemovements, for example, need anaccurate geoid as a starting point. Heatand mass transport by oceans areimportant elements of climate change,but they are still poorly known andawait measurement of ocean surfacecirculation.

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The Benefits from GOCE

GeodesyGeodesy is concerned with mappingthe Earth’s shape, to the benefit of allbranches of the Earth sciences.Whereas positions on the Earth’ssurface can be measured purelygeometrically, height determinationrequires knowledge of the gravityfield. Geodetic levelling provides mm-precision over short distances, buthas systematic distortions on acontinental scale that severely limitthe comparison and linking of heightsystems used in neighbouringcountries or, for example, of tidegauges on distant coasts. Separationof land areas by sea inevitably leadsto large discontinuities betweenheight systems. GOCE data will serveto control or even replace traditionallevelling methods, making feasiblelevelling with a global spacebornereference system such as GPS andGalileo. This will help us towardsworldwide unification of heightsystems, allowing, for example,comparison of the sea-level and itschanges in the North Sea with thosein the Mediterranean.

Absolute Ocean Circulation With ocean topography measured byaltimeter satellites using GOCE’sgeoid as the reference surface, almostall ocean current systems from thestrongest (Gulf Stream, Kuroshio,Antarctic Circumpolar Current)through to weaker deep-ocean andcoastal current systems will bedetermined in terms of location andamplitude. Uncertainties in mass andheat transport will be halved in theupper layers, with significantreductions throughout the oceandepths. Clear benefits are expected inhigh-resolution ocean forecasting.

Solid Earth Detailed 3-D mapping of densityvariations in the lithosphere andupper mantle, derived from acombination of gravity, seismictomography, lithospheric magnetic-anomaly information and topographicmodels, will allow accurate modellingof sedimentary basins, rifts, tectonicmovement and sea/land verticalchanges. It will contributesignificantly to understanding sub-

The accuracies necessary to resolve the notedphenomena and processes. The curves showcurrent accuracies and what can be expected fromGOCE. (INS: Inertial Navigation System –uncertainties in the geoid introduce errors in inertialnavigation.)

crustal earthquakes – a long-standingobjective.

Ice SheetsGOCE will improve our knowledge ofthe bedrock landscape under theGreenland and Antarctic ice sheets,especially undulations at scales of50-100 km. A precise geoid will be amajor benefit to modern geodeticsurveys of the ice sheets, whileimproved gravity maps in polarregions will help satellites to measuretheir orbits using altimeters.

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field’s long-wavelength effects on theorbit will be detected by thistechnique. Laser-ranging willindependently monitor GOCE’s orbitto cm-precision to validate theGPS/Glonass-based orbit.

Satellite configuration: bus is anoctagonal prism 4 m long, 90 cm dia,of sandwich side panels connected byCFRP longerons. Total length 5.02 m;span 2.20 m across fixed solar wings.Small cross-section of 0.8 m2 andsymmetry minimises atmosphericperturbations. Six transverseplatforms.

Attitude/orbit control: paired Xe ionthrusters operate continuously alongthe main axis to counteract

atmospheric drag and torques. Thrustcontrollable 1-22 mN; 22 kg Xe.Attitude control proportional cold-gasor electric thrusters; thrust selectable0-1 mN. GPS/Glonass receiver,gradiometer and star trackers provideattitude/orbit data.

Thermal: MLI, paint pattern, OSRs,heaters, thermistors. GOCE flies inSun-synchronous orbit with one sidealways facing the Sun. High-dissipation units such as the batteryand transponders are mounted onthe ‘cold’ side.

Power system: >1.3 kW EOL fromfixed 3-panel GaAs wings, supportedby 19 Ah NiCd battery.

Communications: controlled fromESOC via Kiruna. 5 kbit/s datastored on 1 Gbit SSR for S-banddownlinking to Kiruna at maximum1 Mbit/s.

GOCE Payload

Gradiometer

The Electrostatic Gravity Gradiometer (EGG)is a set of six 3-axis capacitiveaccelerometers mounted in a diamondconfiguration in an ultra-stable carbon/carbon structure. Each accelerometer pairforms a ‘gradiometer arm’ 50 cm long, withthe difference in gravitational pull measuredbetween the two ends. Each accelerometermeasures the voltage needed to hold theproof mass centred between electrodes,controlled by monitoring the capacitancebetween the mass and the cage. Three armsare mounted orthogonally: along-track,cross-track and vertically.

Laser Retroreflector (LRR)

As used on ERS and Envisat: 9 corner cubesin hemispherical configuration, Earth-facingfor orbit determination by laser ranging.

SREM

ESA Standard Radiation EnvironmentMonitor for the ionising radiationenvironment in the EGG’s sensitivity range.

Satellite-to-Satellite Tracking (SST)

Position, speed and time data using GPSC/A signal and, optionally, Glonass.

CDMU: command & datamanagement unit. FCU: flowcontrol unit (ion propulsion).

PCDE: power control &distribution electronics (ionpropulsion). PCDU: powercontrol & distribution unit.

SSTIE: satellite-to-satellitetracking instrument

electronics.

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VVegaegaAchievements: cost-effective small launcherLaunch dates: planned debut December 2005Launch site: Kourou, French GuianaLaunch mass: 132 tPerformance: optimised for LEO. 1500 kg into 700 km 90º polar, 800-1200 kg into

1200 km Sun-synchronousPrincipal contractors: Vega: FiatAvio, CASA, Fokker, TNO-StorkContraves, CRISA,

SABCA, Saab; P80: FiatAvio, Europropulsion, Snecma, TNO-Stork, Sabca

Vega’s primary role is to fill a gap inEurope’s line of launchers. While themarket segment initially targeted wasscience satellites of 1000-1200 kginto LEO, new forecasts prompted afocus on polar-orbiting Earthobservation satellites of 400-2500 kg.The launch service price will be below$20 million (15% lower thanequivalent US vehicles), assisted bysynergy with Ariane-5 production andoperations and based on only 2-4government and 1-2 commercialmissions annually.

Vega began as a national Italianconcept in the 1980s. BPD Difesa eSpazio in 1988 proposed a vehicle tothe Italian space agency (ASI) toreplace the retired US Scout launcherbased around the Zefiro (Zephyr)motor developed from the company’sAriane expertise. The design wassignificantly reworked in 1994towards the current configuration.Over the same period, CNES wasstudying a European Small Launcherdrawing on Ariane-5 technology andfacilities. Spain was also studying itsown smaller Capricornio to operatefrom the Canary Islands.

In February 1998, ASI proposed Vegaas a European project. In April 1998,ESA’s Council approved a Resolutionauthorising programme start in June;but this was only a limiteddeclaration, approving the first stepof pre-development activities, notablyon solid-booster design. Step-1,running from June 1998 to

September 1999, studied a Vega usinga P85 first stage derived fromAriane-5’s existing strap-on. At theOctober 1999 Council, Francedeclined to support that approach. InNovember 2000, after a further trade-off between different vehicleconfigurations and options for the firstand third stages, it was agreed toproceed with two strands: anadvanced booster that could serve asboth an improved Ariane-5 strap-onand as Vega’s first stage, and the Vegaprogramme itself.

The Vega Programme was approved byESA’s Ariane Programme Board on 27-28 November 2000, and the projectofficially started on 15 December 2000when seven countries subscribed tothe Declaration.

The separate P80 Solid PropulsionStage Demonstrator programme willcost a total of €123 million (2000

What’s in a Name?

Vega is the brightest star in theconstellation of Lyra, and one ofbrightest in the northern sky. Thename has never been used for a launchvehicle before - but it has come close. Itwas suggested in 1973 for the L3Sdesign, but at the time it evoked abrand of beer. L3S, of course, insteadtook the name Ariane. Even furtherback, NASA planned the Atlas-Vegalauncher for deep-space missions butthe upper stage was cancelled in 1959.

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conditions), with €54 million fromESA, €63 million from Italy via theprime contractor and €6.6 millionfrom Belgium via ESA’s GeneralSupport & Technology Programme.The overall contributions aretherefore (€ million): Belgium 6.6,France 45.1, Italy 68.6, Netherlands2.7, Spain t.b.d.

The ESA cost-to-completion for Vegais €335 million (1997 conditions),including the €44 million of Step-1.The national contributions (€ million)are: Belgium 5.63%, France 1.5%(Step 1 contribution), Italy 65%,Netherlands 2.75%, Spain 5%,Sweden 0.8%, Switzerland 1%.

Vega’s PDR was held in June-July2001, with the CDR planned formid-2003. The qualification flight inDecember 1995 will be offered tocustomers at a reduced price. P80’sPDR is planned for end-2001, thedevelopment firing test end-2004,CDR end-2004, Qualification Motor-1firing early 2005 and QM-2mid-2005.

Vega will be integrated in a newBâtiment d’Intégration Vega (BIV) atKourou; the final configuration of theground installations is under study.

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Vega is sized for 1000-1500 kg into 700 km polar orbits.

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Vega CharacteristicsTotal length: 30.22 mPrincipal diameter: 3.005 mLaunch mass: 132 tLaunch thrust: 2700 kN sea levelCapability: reference mission is 1500 kginto 700 km 90º orbit from Kourou.Injection accuracy: ±10 km altitude,±0.1º inclination

Reliability goal: 0.98Guidance: avionics mounted on AVUM,most derived from Ar-5. Stages-1/2 useAr-5 guidance approach, stage-3/AVUMAr-4.

Stage-1Principal contractor: FiatAvioLength: 12.18 mPrincipal diameter: 3.005 m Motor: P80, filament-wound graphite-epoxy casing, 84 t high-aluminium/lowbinder content, SI 280 s vac, 100 baroperating pressure, length 11.0 m, massfraction 0.92

Thrust: 2500 kN avg vacuumBurn time: 100 sSteering: TVC by ±8º deflection of flexiblenozzle joint by electromechanicalactuators provides vehicle pitch/yawcontrol

Stage-2Principal contractor: FiatAvioLength: 7.5 m Principal diameter: 1.90 m Motor: Zefiro 23, filament-wound carbon-epoxy casing, 23 t HTPB 1614 16%aluminium, finocyl grain, SI 287 s vac,95 bar maximum operating pressure,mass fraction 0.92, EPDM insulation.Scaled up from Zefiro 16 (length 7.3 m)tested 18 Jun 1998, 17 Jun 1999,15 Dec 2000 (QM1), 2 more planned

Thrust: 900 kN avg vacuumBurn time: 70 sSteering: TVC by ±6.5º deflection ofsubmerged flexible nozzle joint byelectromechanical actuators providesvehicle pitch/yaw control, 3D carbon-carbon throat

Stage-3Principal contractor: FiatAvioLength: 3.2 m (including 3.6 m interstage)Principal diameter: 1.915 m Motor: P9 derived from Zefiro 16, filament-wound carbon-epoxy casing, 9 t HTPB1614 16% aluminium, finocyl grain, SI294 s vac, 67 bar maximum operatingpressure. mass fraction 0.925, EPDMinsulation

Thrust: 220 kN avg vacuumBurn time: 116 sSteering: as stage-2

AVUM Attitude & Vernier Upper ModuleAVUM provides the final injectionaccuracy, orbit circularisation, deorbit,roll control during stage-3 burn and3-axis control during all coasts. Lowersection is APM (AVUM PropulsionModule), upper is AAM (AVUM AvionicsModule).Principal contractor: FiatAvioLength: 1.6 m Principal diameter: 1.90 m Propulsion: single 2000-2400 NNTO/UDMH (370 kg propellant) enginefor delta-V, SI 315 s; 50 N GN2 thrustersfor attitude control

Burn time: typically 200 s #1 + 170 s #2

Payload Fairing and AccommodationPayloads are protected by a 2-piecefairing until it is jettisoned after about200 s during the coast after stage-2 burn.Carbon skin on aluminium honeycomb.Total length 7.5 m, diameter 3.5 m;payload envelope 4.5 m high, 2.3 mdiameter. Prime contractor Contraves. Acceleration load (static): 5.5 g maxlongitudinal, 1 g lateralAcoustic: max 140 dB overal

AVUM carries Vega’scontrol system andprovides the finalinjection.

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MetopMetopPlanned achievements: first European meteorological satellite in polar orbitLaunch dates: Metop-1 planned for December 2005, Metop-2 2009, Metop-3 2013Mission end: nominally after 5 yearsLaunch vehicle/site: Soyuz-ST from Baikonur Cosmodrome, KazakhstanLaunch mass: 4175 kg Orbit: planned 825 km circular, 98.7° Sun-synchronous (09:30 local time

descending node), 5-day/71-rev repeat cyclePrincipal contractors: Astrium SAS (prime, service module), Astrium GmbH

(payload module, ASCAT, GRAS), Officine Galileo/Alenia Difesa (GOME-2),Fokker (solar generator); all other instruments are customer-provided

Metop will provide Europe’s firstmeteorological satellites in polar orbit.They were originally part of a muchlarger satellite concept, called POEM,which was to have been the successorto ERS, based on the Columbus PolarPlatform. This very large satellitewould have carried the payloads ofboth Envisat and Metop and wouldhave been serviceable in-orbit. At theESA Ministerial Council in Granada,Spain in 1992, this approach wasabandoned and Envisat and Metopwere born. Metop is a jointundertaking by ESA and Eumetsat aspart of the Eumetsat Polar System(EPS). In addition to the satellites, EPScomprises the ground segment, thelaunches and various infrastructureelements. ESA is funding 64% ofMetop-1, while Eumetsat covers therest and all of the follow-on satellites.The Metop-1 agreement between thetwo agencies was signed on9 December 1999, at the same time asthe industrial contract with MatraMarconi Space (now Astrium). EPS willprovide an operational service for 14years, which requires three satelliteswith nominal lifetimes of 5 years.

The US currently operates polarmeteorological satellites in four Sun-synchronous planes, for two services:an early morning and afternoon pair of DMSP military satellites and a mid-morning and afternoon NOAA civilpair. In future, the joint US/Europeansystem will maintain three orbitalplanes: early morning, mid-morning

and afternoon; EPS/Metop will providethe mid-morning service. During thetransitional phase, the older generationof instruments will continue to fly asthe newer instruments are introduced;Metop-1/2/3 will carry olderinstruments from NOAA as well asmore advanced, European, ones.

Metop stores its data for downlinkingeach orbit, but it also providescontinuous direct data-broadcastservices to users. The channels can beselectively encrypted for decoding by

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commercial customers. The High-Resolution Picture Transmission(HRPT) broadcasts the full data set atL-band for regional meteorologicalorganisations to receive relevant datain realtime. The Low-ResolutionPicture Transmission (LRPT) serviceat VHF provides a subset of the HRPTdata. It is comparable to the existingNOAA automatic picture transmission(APT) service, with the objective ofproviding inexpensive access to low-resolution AVHRR images by localusers.

Notable improvements for Metop are:

– new instruments: ASCAT, IASI,GOME-2 and GRAS;

– an innovative onboard compressionscheme that is a considerableimprovement over the existing APTservice and provides LRPT useraccess to three channels of AVHRRdata at full instrument spatial andradiometric resolution;

– continuous onboard recording ofthe global data set for dumpingevery orbit to a high-latitudeground station, with the globalprocessed data available to userswithin 2.25 h of the measurements;

– high pointing and orbital stabilityto ensure that data may be geo-located without reference toground-control points in imagery;

– a selective encryption system toensure the commercial and data-denial needs of Eumetsat and theUS Government, respectively.

The first three Metops will carrytransition instruments from thecurrent NOAA satellites:

– Advanced Very High ResolutionRadiometer (AVHRR), anoptical/infrared imager for global

http://earth.esa.int

Above: theengineering modelof Metop’s payload

module beingprepared for testing

in ESTEC’s LargeSpace Simulator,

June 2001

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Metop payload

Advanced Very High Resolution Radiometer (AVHRR/3) NOAA

6-channel visible/IR (0.6-12µm) imager, 2000 km swath, 1x1 km resolution. Global imagery ofclouds, ocean and land. 35 kg, 622/39.9 kbit/s (high/low rate), 27 W.

High-resolution Infra-Red Sounder (HIRS/4) NOAA

20-channel optical/IR filter-wheel radiometer, 2000 km swath, IFOV 17.4 km (nadir). 35 kg,2.9 kbit/s, 21 W. Replaced on Metop-3 by IASI.

Advanced Microwave Sounding Unit (AMSU-A1/A2) NOAA

Step-scan 15-channel microwave radiometers for 50 GHz oxygen absorption line, 2000 km swath,IFOV 30 km (nadir). A1/A2: 55/50 kg, 2.1/1.1 kbit/s, 78/30 W.

Microwave Humidity Sounder (MHS) Eumetsat

5-channel quasi-optical heterodyne radiometer, 190 GHz (water vapour absorption line), 89 GHz(surface emissivity), 2000 km swath, IFOV 30 km (nadir). 66 kg, 3.9 kbit/s, 89 W.

Infrared Atmospheric Sounding Interferometer (IASI) CNES/Eumetsat

Fourier-transform spectrometer, 3.62-15.5 µm in 3 bands. 4 IFOVs of 20 km at nadir in a square50x50 km, step-scanned across track (30 steps), synchronised with AMSU-A. 2000 km swath.Resolution 0.35 cm-1. Radiometric accuracy 0.25-0.58 K. Integrated near-IR imager for clouddiscrimination. Water vapour sounding, NO2 & CO2, temperature sounding, surface & cloudproperties. 251 kg, 1.5 Mbit/s, 240 W.

Advanced Scatterometer (ASCAT) ESA

5.255 GHz C-band radar scatterometer with 3 dual-swath (2x500 km width, offset 384 km left/rightof groundtrack) antennas (fore/mid/aft) for measurement of radar backscatter at three azimuthangles to provide surface wind vectors of 4-24 m/s with accuracy ±2 m/s & ±20°, spatial resolution50 km (25 km experimental), 0.57 dB radiometric accuracy. Incidence angle range 25-65°, 10 mspulses of 120 W peak power. Additional products such as sea-ice cover, snow cover and vegetationdensity. 270 kg, 60 kbit/s, 251 W.

Global Ozone Monitoring Experiment (GOME-2) ESA/Eumetsat

Scanning spectrometer, 250-790 nm, resolution 0.2-0.4 nm, 960 km or 1920 km swath, resolution80x40 km or 160x40 km. Double monochromator design: first stage of quartz prism with physicalseparation of 4 channels (240-315, 311-403, 401-600, 590-790 nm); second stage of blazed gratingin each channel. Detector: 1024-pixel random-access Si-diode arrays. Ozone total column & profilesin stratosphere & troposphere; NO2 BrO OClO ClO. Albedo and aerosol; cloud fraction, cloud-topaltitude, cloud phase. 73 kg, 40 kbit/s, 45 W.

GNSS Receiver for Atmospheric Sounding (GRAS) ESA/Eumetsat

GPS satellite occultation (up to 500/day) receiver for bending angle measurement better than1 µrad, fitting data to stratospheric model for temperature profile (vertical sounding of ±1 K withvertical resolution of 150 m in troposphere (5-30 km altitude) and 1.5 km in stratosphere), retrievalof refractive index vs. altitude profile. 30 kg, 60 kbit/s, 42 W.

Advanced Data Collection System (ADCS) CNES

Collection of oceanographic, atmospheric and/or meteorological 400 bit/s or 4800 bit/s data on401.65 MHz from platform transmitters (PTT) on buoys, ships, land sites and balloons worlwide forlater relay to ground via X-band and HRPT. Determines PTT location by Doppler. Can transmit toPTTs on 466 MHz at 200 bit/s or 400 bit/s. 24 kg, 7.5 kbit/s, 64 W.

Space Environment Monitor (SEM-2) NOAA

Multi-channel charged particle spectrometer as part of NOAA’s ‘Space Weather’ activities. Totalenergy of electron & proton fluxes 0.05-1 keV and 1-20 keV; directional & omni measurements in 6bands 30-6900 keV for protons and three bands 30-300 keV for electrons. 18 kg, 166 bit/s, 6 W.

Search & Rescue (S&R) CNES/NOAA

Relay of distress beacon signals; 121.5/243.0/406.05 MHz uplink from EPIRB (Emergency PositionIndicating Rescue Beacon), 1544.5 GHz 2.4 kbit/s downlink. 35 kg, 77 W.

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coverage of clouds, ocean and land;– High-resolution Infra-Red Sounder

(HIRS), a spectrometer with arelatively coarse spatial resolutionand a mechanical scan over a wideswath, from which height profilesof atmospheric pressure andtemperature may be derived(replaced on Metop-3 by IASI);

– Advanced Microwave SoundingUnit-A (AMSU-A), a mechanicallyscanned multi-channel microwaveradiometer for pressure andtemperature profiles;

– Microwave Humidity Sounder(MHS), a new instrument on thelast of the NOAA satellites: globalsounding, cloud and Earthradiation budget, sea ice;

– Space Environment Monitor (SEM),to measure the charged-particleradiation environment;

– Data Collection System (DCS/Argos), which receives brieftelemetry signals from a largeglobal network of remote stations.As well as delivering thesemessages to a central processingsite, this new version can also sendmessages to the terminals;

– Search & Rescue (S&R),immediately rebroadcasts signalsfrom emergency transmitterstypically carried on ships andaircraft.

The new generation of instrumentsoffers improved sensing capabilities:

– Infrared Atmospheric SoundingInterferometer (IASI), an importantdevelopment providing a significantimprovement in determining thevertical temperature and humidityprofiles in the atmosphere;

– Advanced Scatterometer (ASCAT),developed within the Metop-1contract, uses multiple radarbeams to measure the small-scaleroughness of the ocean surfacefrom three directions, over a wideswath on each side of the satellite,enabling the speed and direction ofthe wind to be determined;

– Global Ozone MeasurementExperiment-2 (GOME-2), an

improved successor to the ERS-2GOME-1, is a high-resolutionvisible/UV spectrometer formeasurements over a wide swathand wide spectral range todetermine ozone profiles and totalcolumn amounts of many othertrace gases;

– GNSS Receiver for AtmosphericSounding (GRAS), developed withinthe Metop-1 contract, is a geodetic-quality GPS receiver equipped withthree antennas to measure thesignals from GPS satellites inoccultation by Earth’s atmosphere,revealing temperature and pressureprofiles.

Satellite configuration: payloadmodule supported by box-shapedservice module derived fromEnvisat (mechanical design) andSpot-5 (electrical design) servicemodules. Launch configuration3.4x3.4x16.5 m; in-orbit5x5.2x17.6 m.

Attitude/orbit control: 3-axis by 30.45 Nm/40 Nms (at max2400 rpm) reaction wheels; orbitadjust by redundant 8x22.7 Nthrusters in blowdown mode(22 bar BOL), max. 316 kghydrazine in 4 tanks.

Power system: single 8-panel flatpacksolar array provides 2210 W orbitaverage EOL (1700 W for payload).Each panel 1x5 m. Total area40 m2 carrying 32.4x73.7 mm SiBSR cells, supported by five 40 Ahbatteries.

Communications: data downlinked at70 Mbit/s 7750-7900 MHz X-bandfrom 24 Gbit solid-state recorder to2 ground stations within one orbitof recording. Realtime broadcastingof data for HRPT (3.5 Mbit/s1701.3 MHz L-band) and LRPT(72 kbit/s 137.1 MHz VHF).Spacecraft controlled by Eumetsatvia Kiruna (S) S-band2053/2230 MHz up/down groundstation at 2.0/4.096 kbit/sup/down. Metop autonomy for36 h without ground contact.

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SMOSSMOSPlanned achievements: the first global measurements of soil moisture and ocean

salinity; first passive L-band interferometer in spaceLaunch date: planned for early 2006Mission end: after minimum 3 years (2-year extension expected)Launch vehicle/site: to be selected from Rockot, Dnepr and PSLVLaunch mass: about 600 kg (300 kg payload)Orbit: planned 755 km circular, Sun-synchronous (06:00 local time ascending

node), revisit time 3 days at the equator for ascending passesPrincipal contractors (Phase-A): Alcatel Space Industries platform prime, EADS-

CASA payload prime; Extended Phase-A September 2000 - November 2001;Phase-B/C/D to follow

The Soil Moisture and Ocean Salinity(SMOS) mission will use an L-band(1.4 GHz) passive interferometer tomeasure two crucial elements ofEarth’s climate: soil moisture andocean salinity. It will also monitor thevegetation water content, snow coverand ice structure. SMOS was selectedas the second Earth ExplorerOpportunity mission in ESA’s LivingPlanet programme. The AO wasreleased on 30 June 1998 and 27proposals were received by closure on2 December 1998. On 27 May 1999,ESA approved Cryosat as the first forPhase-A/B, with an extended Phase-Afor SMOS as the second.

Human activities appear to beinfluencing our climate. The mostpressing questions are: is the climateactually changing and, if so, how fastand what are the consequences,particularly for the frequency ofextreme events? Answering thesequestions requires reliable models topredict the climate’s evolution and toforecast extreme events. Significantprogress has been made in weatherforecasting, climate monitoring andextreme event forecasting in recentyears, using sophisticated models fedwith data from operational satellitessuch as Meteosat. However, furtherimprovements now depend to a largeextent on the global observation of anumber of key variables, including soilmoisture and sea-surface salinity. No

such long-duration space mission hasyet been attempted.

The RAMSES mission (RadiométrieAppliquée à la Mesure de la Salinitéet de l’Eau dans le Sol) was proposedin 1997 to CNES as a Frenchnational mission and studied toPhase-A. Further work produced theSMOS proposal to ESA by a team ofscientists from 10 Europeancountries and the USA, bringingtogether most of the availableexpertise in the related fields. As anExplorer Opportunity mission, it isdedicated to research and headed byLead Investigators (LIs): Yann H. Kerrof the Centre d’Etudes Spatiales de laBiosphère (CESBIO, F); and JordiFont, Institut de Ciències del Mar (E).The LIs are responsible for overallmission design, implementation anddata processing. ESA has overallmission responsibility, CNES ismanaging the satellite bus, and Spainwill manage the Payload Mission andData Centre at Villafranca. A jointteam will be in charge of technicaldefinition and development of themission components.

The microwave emission of soildepends on its moisture content to adepth of a few cm. At 1.4 GHz, thesignal is strong and has minimalcontributions from vegetation and theatmosphere. Water and energyexchange between land and the

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atmosphere strongly depend on soilmoisture. Evaporation, infiltration andrunoff are driven by it. It regulates therate of water uptake by vegetation inthe unsaturated zone above the watertable. Soil moisture is thus critical forunderstanding the water cycle,weather, climate and vegetation.

At sea, the 1.4 GHz microwaveemission depends on salinity, but isaffected by temperature and sea-state.Those factors must be accounted for.

The global distribution of salt in theoceans and its variability are crucialfactors in the oceans’ role in theclimate system. Ocean circulation isdriven mainly by the momentum andheat exchanges with the atmosphere,which can be traced by observing sea-surface salinity. In high-latitude oceanregions such as the Arctic, salinity isthe most important variable because itcontrols processes such as deep waterformation by controlling the density.This is a key process in the oceanthermohaline circulation ‘conveyorbelt’. Salinity is also important for thecarbon cycle in oceans, as itdetermines ocean circulation andplays a part in establishing thechemical equilibrium that, in turn,regulates the carbon dioxide uptakeand release – important for globalwarming.

Monitoring sea-surface salinity couldalso improve the quality of theEl Niño-Southern Oscillation

prediction by computer models. Thelack of salinity measurements createsmajor discrepancies with the observednear-surface currents.

Satellite configuration: CNES Proteusplatform, 1 m cube, hardware on fourside walls. Flight software inMA3-1750 microprocessor; dualredundant MIL-STD-1553B bus.

Attitude/orbit control: 3-axis control to32±3º in pitch by reaction wheels andmagnetotorquers, measured to 0.05ºusing star tracker, Sun sensors, 3gyros and 2 magnetometers. Orbitcontrol by four 1 N hydrazinethrusters (30 kg propellant, includingsafe mode and SMOS disposal EOL).GPS provides 100 m position accuracy.

Power system: 619 W EOL orbitaverage from two solar wings of four80x150 cm Si panels each; 220/300 Wrequired for payload/bus. Supportedby Spot-4 NiCd battery.

Communications: controlled fromSpacecraft Operations Control Centreat Toulouse (F), commands uplinkedtypically weekly at 4 kbit/s S-band.Science data downlink under study:either generic Proteus S-band toKiruna (plus possibly second station),or by X-band, both with onboardstorage. Payload Mission and DataCentre, Villafranca (E).

http://www.cesbio.ups-tlse.fr/indexsmos.htm

SMOS Payload

The accuracies required are: soil moisture 4%every 3 days at 50x50 km resolution; salinity0.1 psu every 10-30 days at 200x200 kmresolution (psu = practical salinity unit,equates to 0.1% mass; oceans typically 32-37 psu); vegetation moisture 0.3 kg/m2 every7 days at 50x50 km resolution.

L-band Interferometric Radiometer

A passive interferometer using three CFRParms in Y-shaped configuration. 72 receiverelements, including 21 on each arm, 19 cm-dia. 1.4 GHz L-band, H/V polarisation.Records emission within FOV instantan-eously and then at several incidences (0-55º)as SMOS moves along orbit. Swath width~1000 km, spatial resolution < 35 km incentre FOV, radiometric resolution 0.8-2.2 K.