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XMM-NewtonXMM-NewtonAchievements: high-throughput (2x1022 W/cm2 for 6 h exposure), broadband (0.1-

10 keV), medium resolution (20-30 arcsec) X-ray spectrophotometry andimaging of sources, ranging from nearby stars to quasars

Launch date: 10 December 1999, routine science operations began 1 July 2000Mission end: nominally after 2 years; 10-year extended mission possibleLaunch vehicle/site: Ariane-504 from Kourou, French GuianaLaunch mass: 3764 kg Orbit: Ariane delivered it into 838x112 473 km, 40.0°; six apogee burns 10-16

December produced 7365x113 774 km, 38.9°, 48 hPrincipal contractors: Dornier Satellitensysteme (prime), Carl Zeiss (mirror

mandrels), Media Lario (mirror manufacture/assembly), MMS Bristol (AOCS),BPD Difesa e Spazio (RCS), Fokker Space BV (solar array)

The X-ray Multi-Mirror (XMM)observatory (renamed XMM-Newton inFebruary 2000) forms the second,High-Throughput X-ray Spectroscopy,cornerstone mission of ESA’sHorizons 2000 space science plan,concentrating on the soft X-ray portionof the electromagnetic spectrum(100 eV to 12 keV). By virtue of itslarge collecting area and highlyeccentric orbit, it is making longobservations of X-ray sources withunprecedented sensitivity. Most of the50-200 sources in every image arebeing seen for the first time. Whereasthe German/US/UK Rosat mission,launched in 1990, pushed the numberof known X-ray sources to 120 000,XMM-Newton will see millions.

The heart of the mission is the X-raytelescope. It consists of three largemirror modules and associated focalplane instruments held together by thetelescope’s central tube. Each modulecarries 58 nested gold-coated nickelmirrors using their shallow incidenceangles to guide the incoming X-rays toa common focus for imaging by thescientific instruments.

One module uses an advanced PN-CCDcamera (EPIC; see box) capable ofdetecting rapid intensity variationsdown to a thousand of a second or less– important for tracking down blackholes. All three EPIC cameras measurethe proportions of different X-ray

wavelengths to give a broad impressionof each source’s spectrum. For a morethorough analysis, two telescopesdivert about 40% of their beams withgrating stacks to diffract the X-rays,fanning out the various wavelengths onto a CCD strip (RGS; see box). Spikesstand out at specific wavelengths,corresponding to individual chemicalelements.

Besides its X-ray telescopes, XMMcarries a sensitive conventionaltelescope (OM; see box) to observe thesame sections of sky by UV and visible

XMM’s Flight Modellower module in the

Large Space Simulatorat ESTEC. The three

mirror modules areclearly visible.

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One of the first scienceimages: an EPIC-PNview of the 30 Doradusregion of the LargeMagellanic Cloud. In thisregion, supernovae areseeding the creation ofnew stars. At rightcentre is the blue-whitearc of a supernovashockwave heating theinterstellar gas. At lowerright is the remains ofthe star that explodedas Supernova 1987A –the first naked-eyesupernova since 1604.The brightest object (leftcentre) is also asupernova remnant star.(EPIC Consortium)

Far left: combined 30 hexposures from theEPIC-PN and EPIC-MOS cameras of theLockman Hole, where a‘window’ in our Galaxyprovides a view into thedistant Universe. Red(low-energy) objectswere previouslyobserved by Rosat, butgreen and blue objectsare much moreenergetic and are beingseen clearly for the firsttime. It is evident thatXMM is detecting vastnumber of previouslyunknown X-ray sources.

light so that astronomers know exactlywhat the satellite is observing. Anexciting prospect is that XMM cantrack simultaneously the X-ray andoptical afterglow of gamma-ray bursts.

XMM confirms Europe’s position at theforefront of X-ray astronomy byproviding unprecedented observationsof, for example, star coronas;accretion-driven binary systemscontaining compact objects; supernovaremnants (SNRs); normal galaxiescontaining hot interstellar medium aswell as point-like accretion-drivensources and SNRs; Active GalacticNuclei, which may well be accretion-driven; clusters of galaxies with hotinter-cluster medium.

Activation of the science instrumentsbegan on 4 January 2000, after themirror module and OM doors had beenopened on 17/18 December 1999.After engineering images werereturned, the first science images fromall three EPIC cameras and OM weretaken 19-24 January 2000. The firstRGS spectrum was produced25 January. The Calibration &Performance Validation phase began3 March 2000; routine scienceoperations began 1 July 2000.

Satellite configuration: 10.05 m totallength in launch configuration;10.80 m deployed in orbit, 16.16 mspan across solar wings. XMM’sshape is dominated by the telescope’s7.5 m focal length and the need to fitinside Ariane-5’s fairing. The mirrormodules are grouped at one end inthe Mirror Support Platform (MSP)and connected rigidly by the 6.7 m-long CFRP telescope tube to thescientific instruments in the FocalPlane Array (FPA). Launch was withthe mirror end attached to Ariane’sadapter because of centre of mass

http://sci.esa.int/xmm/

XMM in operationalconfiguration

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The lower module Flight Model in theLarge Space Simulator at ESTEC.

XMM principal features. 1: X-ray mirrormodules (x3). 2: Focal Plane Assembly.3: EPIC MOS camera. 4: EPIC PNcamera. 5: RGS (x2). 6: Antenna (x2).7: OM. 8: Star tracker (x2). 9: ServiceModule. 10: Telescope sunshield.11: Solar array wing (x2).

Left: installing XMM-Newton on itslauncher. (CSG/Arianespace/ESA)

MirrorModule 3

ApertureStops

EPICCamera 3

(MOS)

EPICCamera 2

(MOS)

EPICCamera 1

(p-n)

RGSCamera 2

RGSCamera 1

MirrorModule 2

Optical MonitorTelescope

TelescopeTube

MirrorModule 4

GratingAssembly 2

GratingAssembly 1

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XMM-Newton Scientific Instruments

Three Wolter 1-type mirror modules provide the photon-collecting area(1475 cm2 at 1.5 keV; 580 cm2 at 8 keV EOL) for three CCD imaging arraycameras and two reflection grating spectrometers; complemented by aseparate optical monitor. Each 420 kg mirror module contains 58paraboloid/hyperboloid mirrors 60 cm-long with a gold coating providingthe reflective surface at 0.5° grazing incidence. Outermost mirror 70.0 cmdiameter, innermost 30.6 cm; mirror thickness reduces from 1.07 mm to0.47 mm and separation from 4 mm to 1.5 mm. Focal length 7500 mm.

European Photon Imaging Camera (EPIC)

A CCD camera at the prime focus of each mirror module performsbroadband (0.1-10 keV) imaging/spectrophotometry. Two cameras basedon MOS-CCD technology share the mirrors with the RGS; one camerausing PN-CCD has its own dedicated mirror module. EPIC consortium ledby University of Leicester, UK (PI: M. Turner)

Reflection Grating Spectrometer (RGS)

Half of the exit beams from two mirror modules are intercepted by a20 cm-deep gold-coated reflection grating stack for dispersion to a CCDstrip at the secondary focus. E/∆E of 300-700 (1st order) for 0.35-2.5 keV; resolving power >500 at 0.5 keV. RGS consortium led by SRON,The Netherlands (PI: A.C. Brinkman)

Optical Monitor (OM)

A separate 30 cm-aperture Cassegrain telescope provides simultaneouscoverage at 160-550 nm for coordination with X-ray observations andpointing calibration. Photon-counting detectors provide 17x17 arcminFOV at 1 arcsec resolution down to magnitude +24.5. A 10-positionfilter/prism wheel permits spectroscopy. The OM group is led by theMullard Space Science Laboratory, UK (PI: K.O. Mason)

Right: each of XMM’s mirror modules houses 58nested mirrors to reflect the incoming X-rays at

grazing angles to a focus. Two modules also have agrating stack for spectroscopy.

considerations. Subsystems arepositioned around the externalcircumference at this end in theseparate Service Module (SVM),including solar arrays, data handlingand AOCS. ESA’s Integral gamma-rayobservatory uses the same ServiceModule to reduce cost. The FPA isprotected by a Sun shieldincorporating passive radiators forinstruments’ detector cooling.

Attitude/orbit control: four ReactionWheels, two Star Trackers, fourInertial Measurements Units, threeFine Sun Sensors and three SunAcquisition Sensors provide 3-axiscontrol. Pointing accuracy better than1 arcmin, with a drift of 5 arcsec/hand 45 arcsec/16 h. Hydrazinethrusters provide orbit maintenance,attitude control and wheeldesaturation: four sets of two thrusters(primary + redundant) 20 N blowdown(24-5.5 bar) thrusters sit on each offour inter-connected 177 litre titaniumpropellant tanks. Hydrazine loading530 kg.

Power system: two fixed solar wings,each of three 1.81x1.94 m rigid panelswith Si cells totalling 21 m2, sized toprovide 1600 W after 10 years. 28 Vmain bus. Eclipse power from two24 Ah 41 kg NiCd batteries.

Communications: data rate 70 kbit/s inrealtime (no onboard storage) to

stations at Perth (Australia), Kourou(French Guiana) and – added in Feb 2001– Santiago (Chile). Controlled from ESOC;science data returned to Villafranca,Spain. Observations possible for up to40 h out of each 48 h orbit, when aboveEarth’s radiation belts.

Left: the gold-plated mandrel for one of XMM’s mirrors. Mirrormanufacturing was based on a replication process, which transferreda gold layer deposited on the highly polished master mandrel to theelectrolytic nickel shell that was electroformed on to the gold.

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ESA’s Advanced Relay andTechnology Mission Satellite aims todemonstrate new telecommunicationstechniques, principally for data relayand mobile services. At present, usersof Earth observation satellites suchas ERS in low orbit must rely onglobal networks of ground stations toreceive their vital data. But, asinformation requirements and thenumber of missions grow, thisapproach is becoming too slow andexpensive and sometimes technicallyunfeasible. Two payloads aboardArtemis will explore data relaydirectly between satellites, receivingdata from low Earth-orbiting satellitesand relaying them to Europe: theSILEX laser terminal and the SKDRS/Ka-band terminal.

Europe’s Spot-4 and Japan’s OICETSsatellites will use SILEX; Japan’sADEOS-2 Earth-observation satelliteand JEM Space Station module mayuse SKDR. ESA’s own Envisat will bean important user of the Ka-bandelement.

The L-band payload provides 2-waylinks between fixed Earth stationsand land mobiles in Europe, NorthAfrica and the Near East. This‘L-band Land Mobile’ (LLM) payload isfully compatible with the EuropeanMobile System (EMS) payload alreadydeveloped by ESA and flying aboardItalsat-2. It will be used by Eutelsat.

Artemis also carries a transponderproviding part of the EuropeanGeostationary Navigation OverlayService (EGNOS) for enhancing theGPS/Glonass navigation satelliteconstellations.

Artemis is the first ESA satellite to flyelectric propulsion technologyoperationally: UK and German xenonion thrusters will control the north-south drift in geostationary orbit. Theattractions of this technology are itshigh specific impulse (3000 s) and

ArtemisArtemisAchievements: first European data relay mission, including first optical

intersatellite link; first European operational use of ion propulsionLaunch: 21:58 UT 12 July 2001Mission end: design life 10 yearsLaunch vehicle/site: Ariane-510 from ELA-3, Kourou, French GuianaLaunch mass: 3105 kg (550 kg payload, 1538 kg propellant)Orbit: planned geostationary, above 21.5ºEPrincipal contractors: Alenia Spazio (prime, thermal control, RF data relay

payload), FiatAvio (propulsion), CASA (structure), Officine Galileo (power),Fokker (solar array), Bosch/Alcatel (RD data relay payload elements), AstriumSAS (SILEX), Astrium UK (Electro-bombardment thruster), Astrium GmbH (RFIon Thruster), Telespazio (ground segment)

Artemis was released into a low transfer orbit because of a malfunction in Ariane-5’supper stage: instead of the planned858x35 853 km, 2°, the under-burn resultedin 590x17 487 km, 2.94°. As scheduled, thesolar wings were partially deployed some 2 hafter launch and began delivering power whilecontrollers formulated a 4-step recoverystrategy:

1. 18-20 July: the Liquid Apogee Engine (LAE)fired during five perigee passes to raiseapogee to about 31 000 km.

2. 22-24 July: the LAE raised perigee in threeburns to produce a circular orbit at about31 000 km, 0.8°, 20 h. The solar arrays werethen fully deployed, as were the two antennareflectors.

3. software will be created and loaded fororbit-raising by the ion engines.

4. from late September 2001 for severalmonths, the ion engines will raise the orbit toachieve GEO. Spacecraft commissioning thenrequires 2 months.

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Artemis is dominated bythe two 2.85 m-diameterreflectors for inter-orbitlinks and the LLMpayload.

http://telecom.esa.int

Above left: the Flight Model bus atESTEC, showing the Large ApogeeEngine. Above it are two ion thrusternozzles.

Above: unpacking the Flight Modelafter its arrival at ESTEC to begin

systems-level testing in July 1998. TheSILEX terminal is at bottom right; the

telescope is pointing down. Above it isthe Ku-band feeder link for the LLM

land mobile package. At top left is theKa-band feeder for the SKDR and

SILEX packages.

Left: vibration testing of the ArtemisStructural Model in launch configurationat ESTEC in December 1996. One ofthe ion thruster packages can be seenat bottom left.

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low thrust (about 20 mN) in contrastwith chemical propulsion. ForArtemis, the result is a mass savingof about 60 kg. Launch problemsmean that the ion thrusters are beingused unexpectedly to achieve GEO(see box p.166).

Artemis also differs from thetraditional approach by combiningonboard data handling and AOCSinto a centralised processingarchitecture.

Satellite configuration: classical box-shaped 3-axis bus derived fromItalsat and other Europeanpredecessors. 25 m span across solarwings. Primary structure of central

cylinder (aluminium honeycombskinned by carbon fibre), mainplatform, propulsion platform andfour shear panels. The secondarystructure is principally the N/Sradiators, E/W panels and Earth-facing panel. The central propulsionmodule houses the propellant tanks,LAE, pressurant tanks and RCSpipes. E panel: L-band antenna/feed.W panel: IOL antenna. N/S panels:host most of the electroniccomponents requiring heatdissipation. Earth panel: otherantennas and AOCS sensors.

Attitude/orbit control: Earth/Sunsensors & gyros for attitudedetermination, reaction wheels for

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attitude control (RCS thrusters forwheel offloading). UnifiedPropulsion System: conventionalbipropellant (MMH/N2O4) system ofa single 400 N Liquid ApogeeEngine (LAE, for insertion into GEO)and two redundant branches ofeight 10 N RCS thrusters each.Propellants in two Cassini-type 700-litre tanks; helium pressurant inthree spheres. LAE operates atregulated 15.7 bar, then isolatedonce in GEO for RCS to operate inblowdown mode. E/W positioningmaintained by RCS; N/S by ionthrusters. Ion PropulsionSubsystem (IPS) comprises twothruster assemblies, one each onN/S faces: a 15 mN RF Ion Thruster

(RIT) and an 18 mN Electro-bombardment Thruster (EIT). Eachis powered and monitoredseparately, but with commonpropellant supply (40 kg Xe). 600 Wrequired in operation.

Power: twin 4-panel solar wingsprovide 2.8 kW at equinox after10 years to 42.5 Vdc bus; eclipseprotection by two 60 Ah nickel-hydrogen batteries.

Communications: satellite controlfrom Control Centre and TT&Cstation in Fucino Mission Control;In-Orbit Testing from ESA Redu. Seeseparate box for Artemiscommunications payload.

Artemis Communication Payload

Semiconductor Laser Intersatellite Link Experiment (SILEX)

SILEX is the world’s first civil laser-based intersatellite data relay system. The transmitter terminalon Spot-4 in LEO will beam data at 50 Mbit/s (bit error rate <10-6) to the receiver on Artemis forrelay via the feeder link to Spot’s Earth station near Toulouse, France. Japan’s OICETS satellite willalso be used in experiments (including 2 Mbit/s from ground via Artemis to OICETS). Each terminalhas a 25 cm-diameter telescope mounted on a coarse pointing mechanism; pointing accuracy0.8 mrad. Optical power source: 830 nm GaAlAs semiconductor laser diode, peak output 160 mW(60 mW continuous), beamwidth 0.0004º (400 m-diameter circle at the distance of receiver).Receiver: silicon avalanche photodiode (SI-APD), followed by a low-noise trans-impedance amplifier;1.5 nW useful received power. CCD acquisition/tracking sensors direct fine-pointing mechanism oforthogonal mirrors. A 1 m telescope at the Teide Observatory on Tenerife will also act as a teststation.

S/Ka-band Data Relay (SKDR)

The 2.85 m antenna tracks a LEO user satellite via either loaded pointing table values and/or errorsignals to receive up to 450 Mbit/s Ka-band or up to 3 Mbit/s S-band for relay via the feeder link toEarth. Up to 10 Mbit/s Ka-band and 300 kbit/s S-band can be transmitted by Artemis to LEOsatellite. Single Ka-band transponder (plus one backup) 25.25-27.5/23.2-23.5 GHz rx/tx, adjustableEIRP 45-61 dBW, G/T 22.3 dB/K, up to 150 Mbit/s each of three channels LEO to Artemis and upto 10 Mbit/s Artemis to LEO. RH/LHCP on command. One S-band transponder (plus one backup)2.200-2.290/2.025-2.110 GHz rx/tx, adjustable EIRP 25-45 dBW, G/T 6.8 dB/K, 15 MHzbandwidth, up to 3 Mbit/s single channel LEO to Artemis and up to 300 kbit/s Artemis to LEO.RH/LHCP on command. Artemis broadcasts 23.540 GHz beacon to help the LEO satellite track it.

SILEX and SKDR feeder link

Three transponders (plus one backup, 4-for-3) act as Artemis-ground links for SILEX and SKDR.27.5-30/18.1-20.2 GHz rx/tx, EIRP 43 dBW, G/T 0 dB/K, 234 MHz bandwidth, linear verticalpolarisation.

L-band Land Mobile (LLM)

Designed principally for mobile users such as trains and trucks. Artemis carries 2.85 m antenna andmultiple element feed for pan-European coverage and three European spot beams. Three 1 MHz plusthree 4 MHz SSPA channels, providing up to 650 2-way circuits with EIRP >19 dBW 1550 MHzL-band transmitting to terminals (1650 MHz receiving) and 14.2/12.75 GHz Ku-band rx/tx for thefeeder links to the home stations. All channels fully tunable and most commandable LH/RHCP. LLMprovides an operational service in conjunction with Italsat-2’s European Mobile System package (alsofunded through ESA).

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ProbaProbaAchievements: ESA’s first small satellite for technology demonstration, highest-

performance computing system yet flown by ESA, first flight of new type ofEarth-observation instrument

Launch date: planned for September 2001Mission end: 2-year nominal mission durationLaunch vehicle/site: auxiliary passenger on India’s PSLV-C3 (main payload

Technology Experiment Satellite) from Sriharikota, IndiaLaunch mass: 95 kg (payload 25 kg)Orbit: planned 570x640 km, 97.8° Sun-synchronousPrincipal contractors: Verhaert Design & Development (B; prime), Spacebel (B;

software), Verhaert/SIL/Netronics (B/UK/NL; RF, power & avionics), SAS (B;ground segment), Université de Sherbrooke/NGC Aerospace (Canada; attitudecontrol and navigation), Space Systems Finland (FIN; software validation tool),Officine Galileo (I; solar array).

ESA’s small, low-cost Proba (Projectfor On-Board Autonomy) satellite isdesigned to validate new spacecraftautonomy and 3-axis control anddata system technology as part of theAgency’s In-orbit TechnologyDemonstration Programme. India’sPolar Satellite Launch Vehicle willcarry it as a passenger into a Sun-synchronous orbit, which is requiredfor Proba’s main payload. Provided asan ESA Announcement ofOpportunity instrument, this is thenovel CHRIS imaging spectrometer.

Proba is ESA’s first fully autonomousspacecraft, specifically aimed atreducing the cost of space-missionoperations and including some of themost advanced technology yet flown.It carries the highest-performancecomputing system yet flown by ESA,with 50 times the processing power ofSoho. It relies on radiation-hardSPARC (ERC-32) and DSP processors,the latter resulting from an ESA-European Union co-fundedtechnology development effort.

Proba’s high-performance attitudecontrol and pointing system supportsthe complex directional spectralreflectivity measurements of CHRIS.The GPS-based position and attitudedetermination and autonomous starsensor technology make it one of the

best-performing small satellites inproduction.

Proba offers the opportunity for rapidflight-testing of technologies such asminiature digital cameras anddistributed sensing, all of which areof strategic importance for future ESAmissions and European industrialcompetitivity.

In particular, the following autonomyfunctions have been implemented:

– commanding for management ofonboard resource andhousekeeping functions;

– scheduling, preparation andexecution of scientific observations(eg slew, attitude pointing,instrument settings);

– scientific data collection, storage,processing and distribution;

– data communications managementbetween Proba, scientific users andthe ground station;

– performance evaluation andestimation of drifts, trends;

– failure detection, reconfigurationsand software exchanges.

A core of technologies to demonstrateautonomy is accommodated in theattitude control and avionicssubsystems as an integral part ofProba’s design:

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– autonomous double-head startracker, all-sky coverage, high-accuracy for Earth observation andastronomy. Can autonomouslyreconstruct inertial attitude from‘lost in space’ attitude (SpaceInstrumentation Group, TechnicalUniv. Denmark);

– GPS L1 C/A receiver and fourantennas for position and medium-accuracy attitude determination. Itprovides all essential AOCSmeasurements without groundintervention, plus all onboardtiming for autonomous operations(Surrey Satellite Technology Ltd,UK);

– ERC-32 onboard computer performs

guidance, navigation, control,housekeeping, onboard schedulingand resource management. Itsupports all processing normallyperformed on ground. It is a spaceversion of a standard commercialprocessor, intended by ESA tovalidate a computing core forfuture spacecraft. >80 Krad SPARCV7 processor, 10 MIPS &2 MFLOPS with floating-point unit;

– Digital Signal Processor (DSP),TCS21020, for payload dataprocessing;

– Memory Management Unit, 1 Gbit,as part of Payload Processor Unit,to support autonomous datatransmissions (Astrium UK).

Proba is ESA’sfirst small satellite for

technology demonstration.(The apertures reflect an

earlier configuration.)Inset: the flight model

during integration

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S-band patch antenna (x4)

Earth-viewing:CHRIS, HRC, WACA: Astrium/Sira star tracker

B: DTU star tracker

velocity

Sun

B

B

A

DEBIE

DEBIE

GPS antenna (x4)

SREM

Top right and centre right: the flight model at prime contractorVerhaert D&D. Centre left: the CHRIS imaging spectrometer will providenew perspectives for land surface studies. Lower right: the two DEBIEdetectors and their processing unit. Bottom left: vibration testing of thelauncher final stage and piggyback payloads (all structural models) atSriharikota, February 2001. Proba is at right; DLR’s BIRD at left. Bottomright: the SREM radiation detector.

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Proba Scientific Payload

Compact High Resolution Imaging Spectrometer (CHRIS)

Targeted at directional spectral reflectance of land areas; Proba slewingallows multiple imaging (typically 6) of same Earth scene under differentviewing & illumination geometries. This permits new land surfacebiophysical & biochemical information to be derived. For example, itshould be possible for the first time to distinguish between variations inleaf biochemistry and leaf/canopy structure (such as wilting due to waterstress), which both affect the spectral reflectance of vegetation canopies.Range 415-1050 nm, spectral resolution 5-12 nm, spatial resolution 25 m at nadir, swath width 19 km, area CCD. The spectral bands arefully programmable over whole range; up to 19 can be acquiredsimultaneously at full resolution. Contractor: Sira Electro-optics (UK).

Space Radiation Environment Monitoring (SREM)

Measures electron (0.3-6 MeV) & proton (8-300 MeV) fluxes using threedetectors (two stacked) and total dose (RadFET). Box 10x12x22 cm;2.5 kg. Contractor: Contraves (CH).

Debris In-orbit Evaluator (DEBIE)

Measures mass (>10-14 g), speed and penetration power using impactionisation, momentum & foil penetration detection of Proba’s sub-mmmeteoroid & debris environment. Two 10x10 cm impact detectors, on ram side and deep space face, plus processing unit. Standard detectordesigned for different missions with little modification; also planned forInternational Space Station in 2004. Contractor: Patria Finnavitec (FIN).

Cameras

High-level imaging requests handled by Proba’s planning capability topredict ground target visibility. High Resolution Camera (HRC) is a black& white camera to demonstrate high-res (10 m) imaging using aminiaturised telescope (Cassegrain, 115 mm aperture dia, focal length2296 mm) and steerable spacecraft; also complements CHRISmultispectral imaging. 1024x1024 CCD detector using 3D packagingtechnology. FOV 0.504º across diagonal. Wide Angle Camera (WAC) is aminiaturised (7x7x6 cm) black & white camera using a 640x480 CMOSActive Panel Sensor. FOV 40x31º. WAC (already carried by XMM-Newtonand Cluster) is for public relations and educational purposes; 40 Belgianschools in the EduProba project will perform WAC & HRC experiments via the Internet. Contractor: OIP (B).

Smart Instrumentation Point (SIP)

SIP sensors provide measurements of total radiation dose andtemperature around the spacecraft, using smart sensor and 3Dtechnology. Contractors: Xensor (NL), 3D plus (F).

Satellite configuration:600x600x800 mm box-shapedstructure of conventional aluminiumhoneycomb design. Load-carryingstructure is of three panels in ‘H’configuration. Top panel: 4 GPSantennas, 2 S-band patch antennas,solar cells. Ram panel: DEBIE sensor,solar cells. Anti-ram panel: SREMsensor hole, solar cells. Deep-spacepanel: star sensor (2 holes), solar cells.Earth panel: CHRIS aperture, 3 S-bandpatch antennas, launcher interface.

Attitude/orbit control: 3-axisstabilisation by four 5 mN-m Teldixreaction wheels (unloaded by three5 Am2 Fokker magnetotorquers);attitude determination by autonomoushigh-accuracy (<10 arcsec) 2-head startracker, GPS sensor & 3-axismagnetometer. Nadir pointing to150 arcsec accuracy; off-nadir inertialpointing to 100 arcsec. A second startracker (Astrium UK/Sira) is validatinga new star-pattern recognitionalgorithm. Autonomous navigation viaGPS and orbit propagation. No onboardpropulsion.

Power/thermal system: 4x4 cm200 µm-thick GaAs cells on a Gesubstrate body-mounted on five facesprovide 90 W peak (72 W max.required; 17 W in safe mode),supported by 9 Ah Li-ion battery (AEATechnology, UK). 28 Vdc bus. Passivethermal control.

Communications: S-band link to ESARedu (B) control centre (2.4 m dish);4 kbit/s packet TC uplink, 1 Mbit/spacket TM down (2 W redundanttransmitter). 1 Gbit MemoryManagement Unit, orbit allowscomplete dump at least every 12 h ifrequired.

Operations: controlled from a dedicatedground station at ESA Redu (B; 2.4 mdish). Scientific data distributed fromRedu via a webserver. Contractors: SAS(B), Enertec (F), Gigacomp (CH).

Proba-2

Proba-2 will demonstrate technologyminiaturisation and improvements in thecapabilities of small satellites:

– miniaturised attitude sensors;– attitude control for high-res sensors;– electric propulsion, such as FEEP;– deorbit devices, possibly a 30 m tether;– communication using commercial

telecommunications satellite services;– integrated data handling and power;– 100 Mbit/s Ka-band downlink.

It also provides a flight opportunity fortechnological and scientific payloads; an AOwill be released in 2001. Proba-2’s 6-monthPhase-B is expected to begin in 2001;Phase-C/D requires 2 years. Launch isprojected for end-2003. ESA’s financialenvelope is €8 million, as part of the GeneralSupport Technology Programme.

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ESA’s second generation remote-sensing satellite not only providescontinuity of many ERS observations– notably the ice and ocean elements– but adds important new capabilitiesfor understanding and monitoring ourenvironment, particularly in the areasof atmospheric chemistry and oceanbiological processes.

Envisat is the largest and mostcomplex satellite ever built in Europe.Its package of 10 instruments isdesigned to make major contributionsto the global study and monitoring ofthe Earth and its environment, suchas global warming, climate change,ozone depletion and ocean and icemonitoring. Secondary objectives aremore effective monitoring andmanagement of the Earth’s resources,

and a better understanding of thesolid Earth processes.

As a total package, Envisat’scapabilities exceed those of anyprevious or planned Earthobservation satellite. The payloadincludes three new atmosphericsounding instruments designedprimarily for atmospheric chemistry,including measurement of ozone inthe stratosphere. The advancedsynthetic aperture radar can collecthigh-resolution images with avariable viewing geometry, togetherwith new wide swath and selectabledual-polarisation capabilities. A newimaging spectrometer is included forocean colour and vegetationmonitoring, and there are improvedversions of the ERS radar altimeter,

EnvisatEnvisatAchievements: Europe’s largest and

most sophisticated satellite;Europe’s most ambitious Earth-observation mission and a key toolin understanding the Earth system

Launch date: planned November 2001Mission end: 5-year nominal mission

durationLaunch vehicle/site: Ariane-5 from

Kourou, French GuianaLaunch mass: 8140 kg (2145 kg

payload; 300 kg hydrazine)Orbit: planned 800 km circular,

98.54º Sun-synchronous, 35-dayrepeat cycle with same groundtrack as ERS-2, 10:00amdescending node mean local time

Principal contractors: Dornier (missionprime), Matra Marconi Space-Bristol (satellite & Polar Platformprime; MMS-Toulouse: ServiceModule; Dornier: PayloadEquipment Bay), Dornier (payload;MMS-Portsmouth: ASAR; MMS-Toulouse: GOMOS; Alcatel EspaceCannes: MERIS & LRR; Dornier:MIPAS; Alenia Aerospazio: MWR &RA-2), Alcatel Espace (Payload DataSegment)

Envisat in operationalconfiguration.

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microwave radiometer andvisible/near-IR radiometers, togetherwith a new very precise orbitmeasurement system.

Envisat will observe many of thefactors related to changes inatmospheric composition. The resultsof these changes include theenhanced greenhouse effect,increases in levels of UV-B radiationreaching the ground and changes inatmospheric composition.Understanding the processes involvedand the ability to observe the keyparameters are both currentlylacking.

The oceans exert a major influence onthe Earth’s meteorology and climatethrough their interaction with the

atmosphere. Understanding thetransfer of moisture and energybetween ocean and atmosphere, aswell as the transfers of energy by theoceans themselves, are matters ofscientific priority. Envisat willcontribute to this area by providing

Envisat in launch configuration inside the Ariane-5 fairing.

(Dornier)

http://earth.esa.int

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The wavelength rangesof Envisat’s suite of

instruments.

GOMOS will enablesimultaneous monitoring of ozoneand other trace gases in Earth’satmosphere, as well astemperature distributions in thestratosphere.

The range of ASARoperating modes

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information on ocean topography andcirculation, winds and waves, oceanwaves and internal waves,atmospheric effects on the seasurface, sea-surface temperatures,coastal bathymetry and sedimentmovements, as well as thebiophysical properties of oceans.

The Earth’s land surface is a criticalcomponent of the Earth system as itcarries more than 90% of thebiosphere. It is the location of mosthuman activity and it is therefore onland that human impact on the Earthis most visible. Within the biosphere,vegetation is of fundamentalimportance as it supports the bulk ofhuman and animal life and largelycontrols the exchanges of water andcarbon between the land and theatmosphere. Yet our understanding ofthe many processes involved islimited. Envisat observations willcharacterise and measure vegetationparameters, surface water and soilwetness, surface temperature,elevation and topography. These arecritical data sets for improvingclimate models.

Last, but not least, the cryosphere isa key component of the climatesystem. It includes the ice sheets aswell as sea-ice and snow cover. Here,Envisat’s all-weather capabilities willbe exploited to the full as thehostility, remoteness, winterdarkness, inclement weatherconditions and frequent cloud coverof high-latitude ice/snow-coveredregions make the use of remotesensing mandatory. Envisat willprovide important information onseasonal (and long-term) variations insea-ice extent and thickness,evolutions in the ice sheets and snowcover. All affect the climate system;several are very sensitive indicators ofclimate change. Here again, our

knowledge of many of the processesinvolved is lacking.

The satellite comprises the payloadcomplement mounted on the PolarPlatform. Some of the instrumentsfocus on ensuring data continuitywith the ERS satellite: ASAR, AATSRand RA-2 with the MWR, DORISand LRR supporting instruments(see the box for instrument details).Observation of the ocean andcoastal waters – with the retrieval ofmarine biology constituentinformation – is the primaryobjective of MERIS. The ability toobserve the atmosphere, followingon from the GOME instrument onERS-2, is significantly enhanced bythree complementary instruments:SCIAMACHY, GOMOS and MIPAS.They can detect a large number ofatmosphere trace constituents byanalysing absorption lines, andcharacterise atmospheric layers bycomplementary limb and nadirobservations.

Envisat flight model during integration andalignment tests at ESTEC, April 2000. (ESA)

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Satellite configuration: 10.5 m long,4.57 m diameter envelope in launchconfiguration; 26x10x5 m deployed inorbit. The satellite comprises the bus(Polar Platform; PPF) and the payload.The PPF is divided into the ServiceModule (derived from Spot-4’s SMand providing power, AOCS andS-band communications) and thePayload Module providing datahandling, power and communicationsfor the payload.

Attitude/orbit control: primary 3-axisattitude control by five 40 Nmsreaction wheels, magnetorquers forfine control (0.1º 3σ). Hydrazinethrusters provide orbit adjust andattitude control. Attitudedetermination by Earth, Sun and starsensors, with gyros.

Power system: single 5x14 m14-panel Si solar wing generates6.5 kW after 5 years (1.9/4.1 kWaverage/peak to payload); eclipsepower provided by eight 40 Ah NiCdbatteries.

Communications: Flight OperationsSegment including Flight OperationsControl Centre at ESOC working viathe primary S-band TT&C station atKiruna-Salmijärvi (S). Payload DataSegment including Payload DataControl Centre at ESRIN, PayloadData Handling Stations at Kiruna-Salmijärvi (X-band data) and ESRIN(Ka-band data via Artemis), PayloadData Acquisition Station at Fucino (I,X-band data) and 6 ProcessingArchiving Centres in F, UK, D, I, E &S. Data transmission by 100 Mbit/schannel for ASAR and one0-32 Mbit/s and nine 0-10 Mbit/schannels for others. Three 30 Gbittape recorders with Low Bit Rate(LBR) recording at 4.6 Mbit/s;60 Gbit solid-state recorder for LBRplus ASAR high-rate & MERIS full-resolution; playback 50 Mbit/s.Realtime and recorded data viaX-band direct links and Ka-band viaArtemis (steerable 90 cm-diameterantenna on 2 m-long mast); eachwith two channels of 50 &100 Mbit/s.

Dornier

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Envisat Payload

Advanced Synthetic Aperture Radar (ASAR) ESA

Improved version of ERS SAR. 5.331 GHz, 1.3x10 m array of twenty 66.4x99.5 cm radiating panels(each 16 rows of 24 microstrip patches). Imaging mode: HH or VV polarisation, 29x30 m/2.5 dBresolution, 7 selectable swaths 100-56 km wide at 15-45º incidence angles, 96.3 Mbit/s, 1365 Wpower consumption. Alternating polarisation mode: HH + VV, 29x30 m/3.5 dB, 7 selectable swaths100-56 km at 15-45º, 96.3 Mbit/s, 1395 W. Wide swath mode: HH or VV, 150x150 m/2.5 dB,405 km swath width in 5 subswaths at 17-42º, 96.8 Mbit/s, 1200 W. Global monitoring mode: HH orVV, 1000x1000 m/1.5 dB, 405 km swath width in 5 subswaths at 17-42º, 0.9 Mbit/s (allowingonboard storage), 713 W. Wave mode: HH or VV, 30 m/2.0 dB, two vignettes of 5x5 km every100 km in any swath at 20-45º, 0.9 Mbit/s, 647 W. Up to 30 min of high-resolution imagery can bereturned on each orbit. 830 kg.

Radar Altimeter (RA-2) ESA

Fully-redundant nadir-pointing pulse-limited radar using 1.2 m-diameter dish at 13.575 GHz &3.3 GHz. Derived from ERS RA; 3.3 GHz channel added to correct for ionosphere propagationeffects. Fixed pulse repetition frequencies of 1800/450 Hz are used respectively by the twochannels. Onboard autonomous selection of transmitted bandwidth (20, 80 or 320 MHz) makespossible continuous operation over ocean, ice and land. Altitude accuracy after ionosphericcorrection improved to <4.5 cm for Significant Wave Height up to 8 m. 110 kg, 100 kbit/s, 161 W.

Advanced Along-Track Scanning Radiometer (AATSR) UK/Australia

Continues data from ERS-1/2 ATSR on sea-surface temperatures (accurate to 0.5 K) for climateresearch and operational users. AATSR adds land and cloud measurements for vegetation biomass,moisture, health and growth stage, and cloud parameters such as water/ice discrimination andparticle size distribution. Seven channels: 0.555, 0.67, 0.865, 1.6, 3.7, 10.85 & 12 µm, spatialresolution 1x1 km, swath width 500 km. 101 kg, 625 kbit/s, 100 W.

Scanning Imaging Absorption Spectrometer for Atmospheric Chartography (SCIAMACHY) D/NL

240-2380 nm grating spectrometer (limb/nadir viewing) for detrimental trace gas measurement introposphere/stratosphere. Resolution 2.4 Å UV and 2.2-14.8 Å visible/IR. Swath 1000 km wide innadir mode. 198 kg, 400 kbit/s (1867 kbit/s realtime), 122 W.

Medium Resolution Imaging Spectrometer (MERIS) ESA

MERIS is the first programmable imaging spectrometer. 400-1050 nm, 250 m spatial/12.5 nmspectral resolution (adjustable as required), swath width 1130 km. Water quality measurements,such as phytoplankton content, depth and bottom-type classification and monitoring of extendedpollution. Secondary goals: atmospheric monitoring and land surfaces processes. 207 kg,24/1.6 Mbit/s (full/reduced resolution), 148 W average.

Michelson Interferometer for Passive Atmospheric Sounding (MIPAS) ESA

A Fourier transform spectrometer observing mid-IR 4.15-14.6 µm limb emissions with high spectralresolution (<0.03 cm-1), allowing day/night measurement of trace gases (including the completenitrogen-oxygen family and several chlorofluorocarbons) in stratosphere and cloud-free troposphere.Global coverage, including poles. 320 kg, 533 kbit/s (8 Mbit/s raw), 195 W.

Global Ozone Monitor by Occultation of Stars (GOMOS) ESA

Two UV to near-IR spectrometers observe setting stars through the atmosphere for 50 m verticalresolution and 0.1% annual variation sensitivity of ozone/related gases. Occultation method is self-calibrating and avoids the long-term instrumental drift problems of previous sensors.Spectrometer A: 2500-6750 Å, resolution 0.3 nm/pixel; B: 9260-9520 Å (H2O) and 7560-7730 Å(O2), resolution 0.05 nm/pixel. The Fast Photometer Detection Module provides 1 kHz 2-bandscintillation monitoring of the star image. 163 kg, 222 kbit/s, 146 W.

Microwave Radiometer (MWR) ESA

A nadir-viewing 23.8/36.5 GHz Dicke radiometer with 600 MHz bandwidth. ERS MWR designmodified mainly in the mechanical layout and antenna configuration. Radiometric stability <0.5 Kover 1 year. Periodic onboard calibration by switching receiver input between two references: a hornpointing at the cold sky and a hot radiator at ambient. MWR determines tropospheric column watervapour content by measuring the radiation received from Earth’s surface to correct RA-2 altitudemeasurements. 25 kg, 16.7 kbit/s, 23 W.

Doppler Orbitography & Radio-positioning Integrated by Satellite (DORIS) France

DORIS determines the orbit with accuracy of cm. It receives 2.03625 GHz & 401.25 MHz signalsfrom ground beacons and measures the Doppler shift every 7-10 s. 91 kg, 16.7 kbit/s, 42 W.

Laser Retroreflector (LRR) ESA

Precise orbit determination using cluster of laser reflectors mounted close to RA-2.

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ESRO approved development ofEurope’s first applications-satelliteproject in 1972, creating the Meteosatsystem that is now an integral andindispensable part of the world’snetwork of meteorological satellites.The success of the first three pre-operational satellites paved the wayfor the Meteosat OperationalProgramme in 1983 (Meteosat-4/5/6)and the current Meteosat TransitionProgramme (Meteosat-7).

ESA was responsible for developingand operating the system on behalf ofthe newly-created EuropeanOrganisation for the Exploitation ofMeteorological Satellites (Eumetsat),which took direct operational controlon 1 December 1995. The MeteosatSecond Generation (MSG) is ready forintroduction in 2002 to provideservices until at least 2014. ESAcontinues responsibility fordeveloping and procuring thesesatellites; MSG-1 funding is sharedby ESA (two-thirds) and Eumetsat,while MSG-2/3 are fully funded byEumetsat.

MSG is a significantly enhancedfollow-on system to Meteosat,designed in response to userrequirements to serve the needs of‘nowcasting’ applications andNumerical Weather Prediction (NWP),in addition to providing important

data for climate monitoring andresearch.

MSG carries the new SpinningEnhanced Visible and Infrared Imager(SEVIRI), improving on its Meteosatpredecessor:

• 12 spectral channels, instead ofthree, to provide more precise dataabout the atmosphere, improvingthe quality of the startingconditions for NWP models,

• 15-min imaging cycle, instead of30-min, to provide more timelydata for nowcasting, helping in theforecasting of severe weather suchas thunderstorms, snow and fog,

• better horizontal image resolutionin visible light (1 km vs 2.5 km) togreatly help forecasters in detectingsmall-scale weather phenomena,

• the all-digital data transmissionincreases performance and datarates.

Digital image data and meteorologicalproducts will be disseminated via twodistinct channels: High-Rate ImageTransmission for the full volume ofprocessed image data in compressedform; Low-Rate Image Transmissionfor a reduced set of processed imagedata and other data in compressedform. Different levels of access toHRIT and LRIT data will be providedthrough an encryption system.

MSGMSGAchievements: continues and extends Europe’s geostationary meteorological

satellite systemLaunch dates: planned MSG-1 July 2002; MSG-2 December 2003; MSG-3 2007Mission end: 2014 (7-year design lives)Launch vehicle/site: Ariane-4 or -5 from Kourou, French Guiana Launch mass: about 2 t (MSG-1 dry mass 1063 kg)Orbit: geostationary over 0°Principal contractors: Alcatel Space Industries (Cannes, prime), Astrium SAS

(radiometer), Astrium GmbH (power system, AOCS, propulsion system), AleniaSpazio (Mission Communication Package), Saab-Ericsson Space (Data HandlingSubsystem)

http://www.esa.int/msg/

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The Eumetsat central processingfacility in Darmstadt will generate arange of meteorological products,including:

• atmospheric wind vectors atvarious altitudes,

• cloud analysis providingidentification of cloud layers withcoverage, height and type,

• tropospheric humidity at mediumand upper levels,

• high-resolution precipitation index,• cloud top height images for aviation

meteorology,• clear sky radiances• global airmass instability,• total ozone product.

The Global Earth Radiation Budget(GERB) instrument was selected byESA as an Announcement ofOpportunity payload for MSG-1.Eumetsat then decided to fundfurther instruments for MSG-2 andMSG-3. GERB will provide criticaldata on the Earth’s reflected solarand thermal radiation for climateresearch.

MSG will provideenhancedmeteorological servicesuntil at least 2014. (ESA)

The MSG-1 Structural and Thermal Model (STM)during assembly at Alcatel’s facility in Cannes,

France. The SEVIRI imager is in the centre, with its(covered) viewing aperture pointing down and left.

(ESA/Alcatel Space Industries)

http://www.eumetsat.de

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The principal features of the Meteosat SecondGeneration spacecraft. (Alcatel Space Industries)

Integration of the SEVIRI Engineering Model at AstriumSAS. The telescope stands 1.3 m high, plus the 1.2 m-

high passive cooler for cooling the infrared detectors.

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MSG Earth Observation Payload

Spinning Enhanced Visible and Infrared Imager (SEVIRI)

SEVIRI returns 12 visible/IR full-disc Earth images every 12 min,followed by a 3 min reset period. The 5367 mm-focal length telescopeuses three lightweight Zerodur mirrors. The 51 cm diameter circularmain mirror is combined with a 512x812 mm elliptical scanning mirror,stepped with MSG’s rotation to scan Earth’s disc at 9 km intervals southto north. There are 42 detector elements: 9 for High-Resolution Visible(HRV, 0.5-0.9 µm) and 3 each for the other 11 channels: 0.56-0.71, 0.74-0.88, 1.50-1.78, 3.48-4.36, 8.30-9.10, 9.80-11.80 & 11.00-13.00 µm(enhanced imaging); 5.35-7.15, 6.85-7.85, 9.38-9.94 & 12.40-14.40 µm(air-mass, pseudo-sounding). Data sampling intervals 3 km, except 1 kmfor HRV. 270 kg, 153 W nominal power consumption.

Global Earth Radiation Budget (GERB)

Together with SEVIRI, GERB enables study of water vapour and cloudforcing feedback, two of the most important (and poorly understood)processes in climate prediction. It measures Earth’s long- and shortwaveradiation to within 1%; a full scan takes 5 min. Bands: 0.35-4.0 and0.35-30 µm (4.0-30 µm by subtraction). 45x50 km nadir pixel size.Instrument 26 kg, 36 W power consumption, 55 kbit/s data rate.Cooperative project by UK, Italy and Belgium; developed by consortiumled by Rutherford Appleton Laboratory, UK.

In addition to these instruments,MSG will receive 100 bit/s DataCollection Platform information fordistribution in near-realtime, and theSearch & Rescue transponder willrelay 406 MHz distress signals fromships, aircraft and other vehicles tothe COSPAS-SARSAT system.

Satellite configuration: 3.218 m-dia,3.742 m-high (2.4 m body) steppedcylinder, with SEVIRI field of view at90° to axis for scanning Earth disc.

Attitude/orbit control: operationallyheld within ±1° at 0° longitude bythrusters. Spin-stabilised (bythrusters) at 100 rpm anti-clockwisearound main axis parallel to Earth’saxis. Attitude information from Earthhorizon and Sun sensors.Bipropellant (MMH/MON-1) UnifiedPropulsion System (94 kg dry) of two400 N apogee engines for 3-burninsertion into GEO and six 10 Nthrusters. Four 75 cm-dia sphereshold up to 976 kg propellant.

Power system: eight panels 2.4 mhigh, 1.25 m wide on cylindrical bodycarry 7854 32x60 mm high-ε Si cells,supported by two 29 Ah 27.5 kg NiCdbatteries, deliver 700 W at equinoxafter 7 years. Solar array mass 76 kg.

Communications payload: raw datadownlinked at 3.27 Mbit/s1686.8 MHz L-band using one ofthree 10 W solid-state transmitters,together with relayed Data CollectionPlatform data, to Eumetsat inDarmstadt, Germany. L-bandretransmits processed imageryreceived at S-band and other data tousers in high/low rate data streams(up to 1 Mbit/s without compression).MSG’s data volume is a magnitudegreater than the current Meteosat.Telecommand/telemetry at S-band.

The MSG-1 STM being prepared for shipping fromAlcatel in Cannes to ESTEC for testing in 1998. The

black & white panels simulated the thermalcharacteristics of the real solar array. SEVIRI’s

covered aperture is at left. (ESA/Alcatel Space Industries)

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ESA’s International Gamma-RayAstrophysics Laboratory (Integral) isdedicated to the fine spectroscopy(E/∆E=500) and fine imaging (angularresolution: 12 arcmin FWHM) ofcelestial gamma-ray sources in theenergy range 0.015-10 MeV. Integralwas selected by the Agency’s ScienceProgramme Committee in 1993 as theM2 medium-size scientific mission. Itis conceived as an observatory withcontributions from Russia (Protonlauncher) and NASA (Deep SpaceNetwork ground stations).

Integral will provide the sciencecommunity at large with anunprecedented combination ofimaging and spectroscopy over a widerange of X-ray and gamma-rayenergies, including opticalmonitoring. For the first time,astronomers will be able to makesimultaneous observations over sevenorders of magnitude in photon energy(from visible light to gamma-rays) ofsome of the most energetic objects inthe Universe.

Gamma-ray astronomy exploresnature’s most energetic phenomenaand addresses some of the mostfundamental problems in physics andastrophysics. It embraces a greatvariety of processes: nuclearexcitation, radioactivity, positronannihilation, Compton scattering, andan even greater diversity ofastrophysical objects andphenomena: nucleosynthesis, novaand supernova explosions, the

interstellar medium, cosmic rayinteractions and sources, neutronstars, black holes, gamma-ray bursts,active galactic nuclei and the cosmicgamma-ray background. Not only dogamma-rays allow us to see deeperinto these objects, but the bulk of thepower radiated by them is often atgamma-ray energies.

The two main instruments are SPIand IBIS (see table). The finespectroscopy by SPI over the entireenergy range will permit spectralfeatures to be uniquely identified andline profiles to be determined for

IntegralIntegralAchievements: detailed spectroscopy and imaging of celestial gamma-ray sources;

largest ESA science satelliteLaunch date: planned for September 2002Mission end: 2-year nominal duration; spacecraft designed for minimum 5 yearsLaunch vehicle/site: planned on Proton from Baikonur Cosmodrome, KazakhstanLaunch mass: about 4100 kgOrbit: planned 10 000x153 000 km, 51.6°, 72 hPrincipal contractor: Alenia Spazio

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physical studies of the source region.The fine imaging capability of IBISwithin a large field of view willaccurately locate and hence identifythe gamma-ray emitting objects withcounterparts at other wavelengths,enable extended regions to bedistinguished from point sources andprovide considerable serendipitousscience, which is very important foran observatory-class mission.

These instruments are supported bytwo monitors for complementaryobservations in the X-ray (JEM-X)and optical (OMC) bands. SPI, IBISand JEM-X have a common principleof operation: they are all coded-aperture mask telescopes. Thistechnique is the key that allowsimaging at these high energies, whichis all-important in separating andlocating sources.

One of Integral’s most importantobjectives is to study compactobjects, e.g. neutron stars and blackholes. Almost all are significant

sources of high-energy radiation.Integral will image them inunprecedented detail and SPI willprovide the first detailed physicalanalysis at gamma-ray energies.

star trackers

SPI

OMC

service module

IBIS codedmask

JEM-X codedmasks

Left: the Integral Flight Model ready for payload integration at Alenia Spazio.

Integral’s Structural and Thermal Model (STM) at ESTEC in 1998. The patternedcoded-aperture mask for IBIS is visible at top right. SPI is the cylindrical unit atleft. Facing page: SPI’s 2.75 MeV image of a simulated star using a sodium-24source is the first-ever high-resolution (better than 2°) gamma-ray image beyond1 MeV produced by the coded mask technique. It was generated during the SPIcalibration campaign at CESR Toulouse in April 2001.

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Satellite configuration: 5 m high, bodydiameter 3.7 m, 16 m span acrosssolar wings. The Service Module(SVM) design, including power, datahandling and attitude/RCS, is reusedfrom the XMM mission, with minormodifications. The scienceinstruments are accommodated inthe Payload Module (PLM), designedto be tested separately and thenattached to the SVM via simpleinterfaces.

Attitude/orbit control: four ReactionWheels, two Star Trackers, four InertialMeasurements Units, a RateMeasurement Unit, three Fine SunSensors and three Sun AcquisitionSensors provide 3-axis control. Pointingaccuracy better than 15 arcmin, with adrift of 4 arcsec/h. Hydrazine thrustersprovide orbit adjust, attitude controland wheel desaturation: four pairs(primary + redundant) of 20 Nblowdown (24-5.5 bar) thrusters sit onthe SVM base, supplied by four inter-connected 177-litre titanium

Integral’s STM under test in ESTEC’s Large SpaceSimulator in 1998. The Service Module, at left, is the

XMM STM. Adopting a common SVM design andRussia’s provision of the Proton launch in exchange

for observing time have allowed this cornerstone-class mission to be executed at the cost of a

medium-class project.

propellant tanks. Hydrazine loading520 kg.

Power system: two fixed solar wings,each of three 1.81x1.94 m rigidpanels of Si cells totalling 21 m2,sized to provide 1600 W after10 years. 28 V main bus. Eclipsepower from two 24 Ah 41 kg NiCdbatteries.

Communications: science data rate85 kbit/s in realtime (no onboardstorage) to ESA Redu & NASAGoldstone ground stations. Controlledfrom ESOC; science data routed byESOC to the Integral Science DataCentre (provided by the usercommunity) in Geneva, Switzerland.ESTEC’s Integral Science OperationsCentre will plan the observation foruplinking by ESOC. Observations willbe possible above Earth’s radiationbelts, from about 40 000 km.

Further information on Integral can be found at

http://astro.estec.esa.nl/Integral/

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Integral Scientific Instruments

SPI IBIS JEM-X OMCSpectrometer Imager X-ray Monitor Optical Monitor

Energy range 0.020-8 MeV 0.015-10 MeV 3-35 keV 500-850 nm

Detector 19 6x7 cm Ge 16384 CdTe Microstrip CCDcooled to 85 K (4x4x2 mm) Xe-gas (1.5 bar) + V-filter

4096 CsI(9x9x30 mm)

Detector area (cm2) 500 2600 CdTe; 3100 CsI 2x500 2048x1024 pixel

Spectral resolution 2 keV at 1.3 MeV 7 keV at100 keV 1.5 keV at 10 keV –

Field of View 16 9x9 4.8 5x5(fully coded, degrees)

Angular res. (FWHM) 2.5º 12 arcmin 3 arcmin 17.6 arcsec/pixel

10σ source location < 60 arcmin < 1 arcmin < 20 arcsec < 8 arcsec

Continuum sensitivity* 7x10-8 at 1 MeV 4x10-7 at 100 keV 9x10-6 at 6 keV 19.2m (103 s)

Line sensitivity* 5x10-6 at 1 MeV 1x10-5 at 100 keV 2x10-5 at 6 keV –

Timing accuracy (3σ) 129 µs 92 µs 122 µs vary in 1 s units

Mass (kg) 1309 746 65 17

Power (W) 385 240 50 15

*sensitivities are 3σ in 106 s, units photons/(cm2 s keV) continuum, photons/(cm2 s) line

SPI

The spectrometer performs spectral analysis of gamma-ray sources and regions with unprecedentedenergy resolution using 19 hexagonal high-purity germanium detectors cooled by active Stirlingcoolers to 85 K. A hexagonal coded-aperture mask 1.7 m above the detection plane images largeregions of the sky. To reduce background radiation, the detector assembly is shielded by an activescintillator veto system around the bottom and side of the detectors almost up to the coded mask. Co-PIs G. Vedrenne (CESR Toulouse, France) & V. Schoenfelder (MPE Garching, Germany).

IBIS

The imager provides fine imaging and spectral sensitivity to continuum and broad lines over a wideenergy range, achieved by two layers of detector elements: a front layer of CdTe backed by CsIelements. A tungsten coded-aperture mask 3.2 m above the detection plane is optimised for highangular resolution. The two layers of detectors allow the photons to be tracked in 3D as they scatterand interact with elements. The aperture is restricted by a lead tube system and shielded in all otherdirections by an active scintillator veto system. PI: P. Ubertini (IAS Frascati, Italy).

JEM-X

The X-ray monitor is crucial for identifying gamma sources by making observations simultaneouslywith the main gamma-ray instruments. Its prime 3-35 keV energy band can be extended to 100 keV.Two identical imaging microstrip gas chambers each view the sky through a coded-aperture maskpositioned 3.2 m above the detection plane. PI: N. Lund (Danish Space Research Institute, Denmark).

OMC

The optical monitor consists of a passively cooled CCD in the focal plane of a 50 mm lens. It offersthe first opportunity for making long-duration optical observations simultaneously with those atX/gamma-rays. Variability patterns from tens of seconds up to years are monitored.PI: A. Gimenez (INTA Madrid, Spain).

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SMARTSMART-1-1Planned achievements: test new technologies, first European lunar orbiter, first

global map of lunar elementsLaunch date: nominally November 2002 (window October 2002 - October 2003)Mission end: nominally after 6 months of lunar-orbit operationsLaunch vehicle/site: Ariane-5 from Kourou, French GuianaLaunch mass: 350 kg (15 kg science payload)Orbit: planned GTO, then lunar orbit of 300-2000x10000 km, 90°, with pericentre

over South Pole (±30º)Principal contractors: Swedish Space Corp (prime), SNECMA (plasma thruster),

APCO (bus structure), Saab Ericsson Space (integration, test, thermal, RTUs,harness), Fokker Space (solar array), CRISA (battery management electronics),Primex (hydrazine system); Phase-A May - September 1997, Phase-B April 1998- June 1999, Phase-C/D October 1999 - October 2002

SMART-1 is the first of the SmallMissions for Advanced Research inTechnology of ESA’s Horizons 2000science plan. Its principal mission isto demonstrate innovative and keytechnologies for deep-space sciencemissions. Its primary objective is toflight test Solar Electric PrimaryPropulsion (SEPP) for future largemissions; the BepiColombo Mercurymission will be the first to benefit,followed by Solar Orbiter and,probably, LISA, GAIA and Darwin.

ESA’s Science Programme Committee(SPC) in November 1998 approved alunar-orbiting mission as thebaseline, with the possibility ofextending it to a flyby of a near-Earthasteroid. The SPC approved the€84 million (1999 rates) funding inSeptember 1999; the prime andlaunch contracts were signed inNovember 1999. The low budgetmeant that a low-cost launch and anew procurement and managementapproach were adopted. SMART-1will therefore fly as an auxiliarypayload on a commercial Ariane-5launch into GTO. The SEPP will beused to spiral out from GTO over 15-18 months, followed by lunarswingby, lunar capture and thenspiralling in to a near-polar lunarorbit with a perilune of 300-2000 kmand an apolune of 10000 km.SMART-1’s SEPP will use a 70 mN

Hall-effect xenon plasma thrusterfrom SNECMA,

The SPC confirmed the selectedscience payload in November 1999following AOs in March 1998(science) and April 1998 (technology).It includes a 5° field-of-viewmulticolour micro-camera (AMIE)with high resolution and sensitivityeven for lunar polar areas. A verycompact IR spectrometer (SIR) willmap lunar minerals and look forwater and carbon dioxide ice in

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SMART-1 will use itsPPS-1350 plasmathruster (bottom left) tospiral out from GTO tothe Moon.(ESA/J. Huart)

eternally shadowed craters. An X-raymapping spectrometer (D-CIXS) willprovide the first global map of themajor rock-forming elements and alsoX-ray monitoring of very brightcosmic sources during the cruise. Theabsence so far of global maps ofmagnesium, aluminium and siliconabundances is a serious hurdle tounderstanding the Moon. The XSMsolar X-ray monitor will also performspectrometric observations of the Sunduring the cruise. The lightweightSPEDE will characterise the naturaland induced plasma environmentaround the spacecraft. Thecomplementary EPDP technologypackage will monitor the plasma andcontamination created by the SEPPthruster. The RSIS radio scienceinvestigation makes use of the KATEX-Ka radio transponder technologypayload to perform severalexperiments. One, in conjunctionwith AMIE, will measure lunarlibration as a demonstration of thecrucial BepiColombo investigation ofMercury’s internal structure.

AMIE will help to validate deep-spaceoptical communications (LaserLinkExperiment) using ESA’s OpticalGround Station at the TeideObservatory in Tenerife. The camerawill also validate the OBANautonomous navigation experimentbased on image processing. For the

spacecraft bus, a combination of off-the-shelf hardware and innovativeapplication has been used. Forexample, the communications bus isinherited from the automotiveindustry.

In synergy with its technologyobjectives, SMART-1 provides anopportunity for lunar scienceinvestigations. These include studiesof the chemical composition andevolution of the Moon, of geophysicalprocesses (volcanism, tectonics,cratering, erosion, deposition of icesand volatiles) for comparativeplanetology, and high-resolutionstudies in preparation for futuresteps in lunar exploration. Themission could address several topics

http://sci.esa.int/smart/

SMART-1 Technology and Science Goals

– test SEPP and characterise the inducedenvironment

– test new spacecraft and payloadtechnology for Cornerstone missions(Li-ion modular battery package, X-Kadeep-space transponder with turbo-codes,deep-space laser link, Swept ChargeDevice X-ray detector, onboard softwareauto-code generation & autonomy)

– Moon elemental geochemistry andmineralogy

– Moon geology, morphology & topographyat medium- & high-resolution

– Moon exospheric and polar environment– cruise observations of X-ray cosmic

sources

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SMART-1 payload experiments.

Expt. Investigation Main TeamCode Type Investigator Co-Is Description of Experiment

AMIE Principal J.L. Josset F, NL, Asteroid Moon Imaging Experiment. Investigator (CSEM, CH) FIN. miniaturised CCD (1024x1024-pixel) camera,

I, ESA 27 m res from 300 km, 4 fixed filters (750, 847, 900, 950 nm), 5.3° FOV, 16.5 mm-dia aperture, 154 mm f.l., micro-Data Processing Unit. Also supports LaserLink, OBAN & RSIS. 1.8 kg (camera 0.45 kg), 9 W

LaserLink Guest Z. Sodnik Demonstration of a deep-space laser Technology (ESA) link with ESA Optical Ground Station:Investigator OGS aims 6 W 847 nm laser with

< 10 µrad accuracy for detection by AMIE.

OBAN Guest F. Ankersen Validation of On-Board Autonomous Technology (ESA) Navigation algorithm. AMIE stares at planet/Investigator asteroid, ground software uses star tracker

datato remove spacecraft attitude motions to reveal relative velocity vector. Ground demonstration only.

SPEDE Principal H. Laakso FIN, S, Spacecraft Potential, Electron & Dust Investigator (FMI, FIN) ESA, Experiment. 2 Langmuir Probes on 60 cm-

USA long CFRP booms measure spacecraft potential and plasma environment created by EP. 0-40 eV, 0.7 kg,2 W

EPDP Technology G. Noci I, ESA, Electric Propulsion Diagnostics Package Investigator (Laben FIN, A for monitoring the EP; plasma environment

Proel, I) characterisation. Plasma (0-400 eV):Langmuir Probe and Retarding Potential Analyser. Quartz-Crystal Microbalance measures mass of deposited contamination; dedicated solar cell monitors neutral ion deposition. 2.3 kg, 18 W

RSIS Guest L. Iess USA, D, Radio-Science Investigation System monitors Science (Univ. Rome, UK, F, the Electric Propulsion, using KATE and Investigator I) ESA, S AMIE.

SIR Technology U. Keller D, UK, SMART-1 IR Spectrometer. Miniaturised Investigator (MPAe, D) CH, I, 256-channel near-IR (0.9-2.4 µm) grating

IRL spectrometer for lunar surface mineralogy studies. 1.1 mrad FOV, 6 nm spectral res, 330 m spot size at 300 km. Passively cooled InGaAs array. Good discrimination between pyroxenes, olivines & feldspar; possible detection of H20, CO2 & CO ices/frosts. 1.7 kg, 2.0 W

D-CIXS/ Technology M. Grande S, E, Demonstration Compact Imaging X-ray XSM Investigator (RAL, UK) I, F, Spectrometer for mapping main lunar rock-

J. Huovelin ESA, forming minerals (Si, Mg, Fe, Na, O, C; (Univ. USA 30 km res) via X-ray fluorescence. 24 Swept Helsinki, FIN) Charge Detectors plus micro-collimators.

0.5-10 keV. X-ray Solar Monitor using Peltier-cooled Si diode measures the solar X-rays that excite the Moon’s fluorescence. 0.8-20 keV. Total D-CIXS/XSM: 3.3 kg, 10 W

KATE Technology R. Kohl ESA, Ka-band TT&C Experiment, demonstrates Investigator (Dornier, D) UK, I X-band (8 GHz) + Ka-band (32-34 GHz)

telecommunications (up to 500 kbit/s from lunar orbit) & tracking (X/Ka-band Doppler improves accuracy) and tests turbo-codes (2-3 dB increase) and VLBI operation. 5.2 kg, 18 W

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such as the accretional processesthat led to the formation of planets,and the origin of the Earth-Moonsystem. SMART-1 will also preparethe scientific community for theBepiColombo mission.

Satellite configuration: box-shaped115x115x94.5 cm 45 kg bus ofconventional aluminiumconstruction, with twin solar wingsspanning 14 m. Thin-walled rivetedthrust cone carries main loads (top,equipment and bottom platforms) andhouses Xe tank.

Primary propulsion: PPS-1350 Hall-effect Stationary Plasma Thruster,70 mN at 1350 W inlet power, 10 cm-dia chamber, SI 1500 s, up to 80 kgXe propellant, mounted in 2-degree-of-freedom gimbals, pointing accuracy0.02º.

Attitude/orbit control: 3-axis zeromomentum by 4x2.5 Nms Teldixreaction wheels, 10 arcsec accuracy

for 10 s required; 8x1 N hydrazinethrusters (4 kg hydrazine) for RWunloading and high-rate recovery.Attitude determination to 4 arcsec bytwo autonomous star trackers.

Power system: twin 3-panel (each800x1778 mm) solar wings, adaptedfrom Globalstar design, GaAs/InPmulti-junction cells, 1850 W @ 1 AUBOL. Supported by 5 Li-ion batteriestotalling 600 Wh capacity.

Communications: no realtime scienceor operations, ground contact for< 8 h every 4 d via 15 m ESA stationnetwork. Experimental mobile XKaTlaboratory based at ESTEC as backup(S-band 62 kbit/s, X-band 2 kbit/sfrom lunar orbit, Ka-band120 kbit/s). Redundant 4 Gbit solid-state mass memory. S1MOC MissionOperations Centre at ESOC, STOCScience & Technology OperationsCentre at ESTEC.

Inset: the glow of ionised Xefrom the PPS-1350 thruster

(SNECMA)

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RosettaRosettaPlanned achievements: first comet orbiter, first comet lander, first solar-powered

deep space probe, first European probe beyond MarsLaunch date: planned for 02:12 UT 13 January 2003 (19-day window)Mission end: 10 July 2013 (perihelion)Launch vehicle/site: Ariane-5 from Kourou, French GuianaLaunch mass: 3011 kg (165 kg Orbiter science payload, 90 kg Lander)Orbit: planned 200x4000 km, Earth parking orbit; heliocentric; cometocentricPrincipal contractors: Astrium GmbH (prime), Astrium UK (bus), Astrium SAS

(avionics), Alenia Spazio (AIV), DLR (Lander). Mission Definition 1993-1996,Phase-B March 1997 - September 1998, Phase-C/D 12 April 1999 - May 2002

The International Rosetta Missionwas approved in November 1993 byESA’s Science Programme Committee(SPC) as the Planetary CornerstoneMission in the Horizon 2000 scienceprogramme. Rosetta’s main goal ismankind’s first rendezvous with acomet: 46P/Wirtanen. Comets are themost primitive objects in the SolarSystem so they hold many clues tothe evolution of the Sun and planets.On its way to Wirtanen, Rosetta willinspect two asteroids, Otawara andSiwa, at close quarters.

The mission has a number of uniquefeatures and challenges, including afixed launch date and a flight of10.5 years in deep space while relyingon electrical power from solar arrays.The considerable variations in thedistances from the Sun (1.05-5.25 AU) and the Earth have majoreffects on thermal control, solar arraydesign and telecommunications.Onboard autonomous operations areparticularly important because theround-trip light time for radio signalsexceeds 90 min for a significant partof the mission.

Orbiting the comet for more than ayear, Rosetta will observe changes insurface activity as the nucleus iswarmed by the Sun. Instruments willanalyse the effusions of dust and gasand determine the chemical,

mineralogical and isotopiccomposition of the volatiles. TheLander will provide ground truth databy analysing in situ samples.Cometary material is representative ofthe early solar nebula 4600 millionyears ago; there has been littleevolution since then.

Comet Wirtanen was discovered in1948 by Carl Wirtanen at LickObservatory in the US. It belongs tothe family of short-period comets,which have aphelions at Jupiter’sorbit. As a consequence, the orbitscan be perturbed when they pass

Comet 46P/Wirtanen

Nucleus diameter (km) 1.2Orbital period (yr) 5.45Aphelion (million km/AU) 768/5.25Perihelion (million km/AU) 159/1.05Orbital eccentricity 0.657Orbital inclination (deg) 11.72

Asteroids 4979 Otawara & 140 Siwa

Otawara SiwaAvg distance from Sun 324 409

(million km)Orbital period (yr) 3.19 4.52Size (km) 2.6-4 110Rotation period (min) 162 278Orbital inc (deg) 0.91 3.19Orbital eccentricity 0.1449 0.2157Asteroid type V or SV CDiscovered Aug 1949 Oct 1874

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close to Jupiter – as happened toWirtanen in 1972 (0.28 AU) and 1984(0.46 AU), changing the orbital periodby almost 20% to 5.5 years. AlthoughWirtanen’s orbit is now fairlypredictable, continuous observationsare underway to determine its precisepath and to choose the optimalrendezvous trajectory for Rosetta.Wirtanen was selected in 1994because, for the required launchperiod, it was the only candidateattainable with current technology.

When Rosetta closes in to about100 000 km, the navigation cameraswill image the comet to optimise theapproach trajectory. Rosetta willrendezvous with Wirtanen on theinward leg to the Sun so that it canstudy the nucleus and itsenvironment as solar radiationincreases.

Three planetary gravity-assists plus adeep-space plane-change will set upthe rendezvous (see box). Rosetta will

twice fly through the main asteroidbelt, fulfilling the secondary missionobjective, and fly close to asteroidsOtawara and Siwa (see box). Theseprimordial rocks could hardly bemore different. Siwa will be thelargest asteroid yet encountered by aspacecraft, while Otawara will be thesecond smallest. Otawara issuspected to be a chunk of once-molten basalt, and Siwa is carbon-rich and blacker than coal. Otawara’srapid rotation should allow Rosetta toimage most of its surface during the10 km/s flyby 283 million km fromEarth. Rosetta will fly past Siwa at17 km/s, approaching on the sunlitside and then looking at a crescentphase as it moves away. At this time,they will be about 470 million kmfrom Earth, so signals will take26 min to reach ground stations.

On arrival at Wirtanen, Rosetta willmanoeuvre into an orbit with analtitude of 5-25 nucleus radii,depending on the comet’s actual size,

http://sci.esa.int/rosetta/

Rosetta Scientific Goals

The mission’s primary scientificobjectives are to study the origin of

comets, the relationship betweencometary and interstellar material and the

implications for theories on the origin ofthe Solar System. The measurements in

support of these objectives are:

• global characterisation of the nucleus,determination of dynamic properties,

surface morphology and composition;• determination of the chemical,

mineralogical and isotopiccompositions of volatiles and

refractories in a cometary nucleus;• determination of the physical

properties and interrelation of volatilesand refractories in a cometary

nucleus;• study of the development of cometary

activity and the processes in thesurface layer of the nucleus and the

inner coma (dust/gas interaction);• global characterisation of an asteroid,

including determination of dynamicproperties, surface morphology and

composition.

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Launch (13 Jan 2003): into 200x4000 km Earthparking orbit. After about 2 h, Ariane’s upper stagereignites to inject Rosetta into the interplanetarytrajectory.

Commissioning (Jan-Apr 2003): solar arrays aredeployed, Sun is acquired, all systems are checkedand the payload is commissioned. Rosetta thenhibernates for the cruise to Mars.

Mars flyby (26 Aug 2005): Rosetta flies past Mars at200 km, making some science observations, adding1.53 km/s.

Earth flyby #1 (28 Nov 2005): Rosetta is activeduring the cruise to Earth, flying past at 4500 km,adding 3.5 km/s. Operations mainly involve trackingand orbit determination.

Otawara flyby (11 Jul 2006): Rosetta hibernatesduring the cruise to asteroid Otawara. Flypast at2200 km; science data downlinked after the flyby.

Earth flyby#2 (28 Nov 2007): Rosetta hibernatesduring the cruise to Earth. Flyby at 1370 km altitudeadds 3.7 km/s. Operations mainly involve trackingand orbit determination.

Siwa flyby (24 Jul 2008): Rosetta hibernates duringthe cruise to Siwa. Flypast at 3500 km; science datadownlinked after the flyby.

29 Jun 2009: 430 m/s deep-space manoeuvre tomatch Wirtanen’s orbital plane. Rosetta then entershibernation; during this period, it records itsmaximum distances from the Sun (780 million km)and Earth (1000 million km).

Comet rendezvous (29 Nov 2011): Rosetta isreactivated before the rendezvous manoeuvre, whenthe thrusters fire for several hours to slow therelative drift rate to the comet to about 25 m/s.

Comet approach (Nov 2011 - May 2012): as Rosettadrifts towards the nucleus, controllers avoid dust andachieve good nucleus illumination conditions. Thefirst images dramatically improve calculations of thecomet’s position and orbit, as well as its size, shapeand rotation. The relative velocity is graduallyreduced, to 2 m/s after about 90 days.

Comet mapping/characterisation (May-Jun 2012):less than 200 km from the nucleus, images showthe comet’s attitude, angular velocity, majorlandmarks and other basic characteristics.Eventually, Rosetta is inserted into orbit at 35 kmaltitude. Relative velocity is a few cm/s. The Orbiterstarts to map the nucleus in great detail. Fivepotential landing sites are selected for closeobservation.

Landing (Jul 2012): once a landing site is selected,the Lander is released from a height of 1 km.Touchdown speed is < 1 m/s. Once it is anchored tothe nucleus, the Lander transmits high-resolutionimages and data on the ices and organic crust. TheOrbiter downlinks them to Earth at the next groundstation contact.

Escorting the nucleus (Jul 2012 - Jul 2013): theOrbiter observes events as perihelion approaches.The mission ends in Jul 2013, at perihelion, after3800 days.

Left: a simulated image of Rosetta’s flybyof asteroid Otawara, generated with theplanning software of the Rosetta ScienceOperations Centre.

Right: the Rosetta Structural & ThermalModel during thermal balance tests in the

Large Space Simulator at ESTEC.

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shape and mass. The comet’s radiusis expected to be about 600 m, whilethe Orbiter’s relative velocity will be< 1 km/h much of the time. Adedicated 1-month global mappingprogramme will lead to the selectionof a landing site, while otherinstruments simultaneously measurethe comet’s environment.

The Lander ejection, descent andlanding operation is a complexautonomous sequence of events. TheOrbiter navigation must be precise towithin 10 cm and 1 mm/s at the timeof ejection, at an altitude of 1 km fora touchdown speed of 1 m/s. Localgravity is only 10-4 g, so the Landerwill anchor itself to the dusty

snowball by firing a harpoon oncontact. The minimum goal forsurface operations is a week, but it ishoped that several months will beachieved.

As the Lander carries out itsexperiments, the data will be relayedto Earth via the Orbiter. Thereafter,the Orbiter will follow the comet foranother year. Throughout thecometary phase, there are severetechnical, operational andnavigational challenges becauseRosetta will fly at low altitude aroundan irregular celestial body with aweak, asymmetric and rotatinggravity field, enveloped by dust andgas jets.

MIRO

LGAMIDAS

RPC-MIP

RPC-LAP

RPC-LAP2RPC-MAG(boom notdeployed)

RPC-ICA

RPC-IES

navigationcameras

louvred radiator

Sunsensor

CONSERT

Sunsensor

ALICE

Lander leg (x3)

Berenice (deleted)

Lander

VIRTIS radiatorVIRTIS-M

VIRTIS-HROSINA COPS

ROSINADFMS

star tracker

star tracker

OSIRIS-WAC

OSIRIS-NAC

GIADA

COSIMA

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Rosetta Orbiter Scientific Instruments

ALICE. UV (70-205 nm) imaging spectrometer. Coma/tail gas composition, production rates of H2O& CO2/CO, nucleus surface composition. Spectral res 3-13 Å; spatial res 0.05x0.6º. PI: S.A. Stern,SouthWest Research Inst., USA. 3.1 kg, 2.9 W, Participating: F.

CONSERT (Comet Nucleus Sounding Expt by Radiowave Transmission). Nucleus deep structure.Lander transponds 90 MHz radio waves back to Orbiter after propagation through nucleus.PI: W. Kofman, CEPHAG, F. 3.1 kg, 2.5 W, 24 Mbit/orbit. Participating: D, I, NL, UK, USA.

COSIMA (Cometary Secondary Ion Mass Analyser). Characteristics of dust grains. Time-of-flightmass spectrometer, m/∆m 2000. PI: J. Kissel, MPI für Kernphysik, D. 19.1 kg, 19.5 W. Participating: A, CH, F, FIN, I, NL, USA.

GIADA (Grain Impact Analyser & Dust Accumulator). Number, mass, momentum & speeddistribution of dust grains. PI: L. Colangeli, Obs. di Capodimonte, I. 6.2 kg, 3.9 W. Participating: D,E, F, UK, USA.

MIDAS (Micro-Imaging Dust Analysis System). Dust environment: particle population, size, volume &shape. Atomic Force Microscope with nm-res. PI: W. Riedler, Space Research Inst., A. 8.0 kg, 7.4 W.Participating: D, F, N, NL, UK, USA.

MIRO (Microwave Instrument for the Rosetta Orbiter). Abundances of major gases, surfaceoutgassing rate, nucleus subsurface T. 30 cm-dia dish radiometer & spectrometer, 1.6 & 0.5 mmwavelengths. Spatial res 15 m & 5 m at 2 km, respectively. PI: S. Gulkis, JPL, USA. 19.5 kg, 43 W,2.53 kbit/s. Participating: D, F.

OSIRIS (Optical, Spectroscopic & IR Remote Imaging System). High-resolution imaging. WAC wide-angle camera (10.1 m res at 100 km, 12x12º, 245-800 nm, f.l. 767 mm) & NAC narrow-angle camera(1.9 m res at 100 km, 2.18x2.18º, 250-1000 nm, f.l. 132 mm). PI: H.U. Keller, MPI für Aeronomie, D.30.9 kg, 22 W. Participating: E, F, I, NL, S, TWN, UK, USA.

ROSINA (Rosetta Orbiter Spectrometer for Ion & Neutral Analysis). Atmosphere/ionospherecomposition, velocities of electrified gas particles and their reactions. Double-focusing spectrometer:12-200 amu, m/∆m 3000; time-of-flight spectrometer 12-350 amu, m/∆m 2900. PI: H. Balsiger,Univ. Bern, CH. 34.8 kg, 27.5 W. Participating: B, D, F, USA.

RPC (Rosetta Plasma Consortium). Nucleus physical properties, inner coma structure, cometaryactivity, cometary interaction with solar wind. Five particle/field sensors: Langmuir probe, ion &electron sensor, fluxgate magnetometer, ion composition analyser, mutual impedance probe.PI: R. Boström & R. Lundin, Swedish Inst. Space Physics, S; J. Burch, Southwest Research Inst.,USA; K-H. Glassmeier, TU Braunschweig, D; J.G. Trotignon, LPCE/CNRS, F. 8.0 kg, 10.6 W.Participating: HUN, S, UK.

RSI (Radio Science Investigation). Nucleus mass, density, gravity, orbit, inner coma by Dopplertracking of X-band signal. PI: M. Pätzold, Univ. Köln, D. Uses spacecraft telemetry. Participating:CHL, F, N, S, UK, USA.

VIRTIS (Visible/IR Thermal Imaging Spectrometer). Maps nature & T of nucleus solids, identifiesgases, characterises coma conditions, identifies landing sites. 0.25-5 µm. PI: A. Coradini, IAS-CNR, I.30.0 kg, 28 W, 3 Mbit/s. Participating: F, G.

with the solar wind (GIADA, RPC).CONSERT, on both craft, investigatesthe large-scale structure of the nucleus.RSI uses the telecommunicationssystem to study the mass distributionin the nucleus. The Lander focuses onthe in situ composition and physicalproperties of nucleus material. Itcarries cameras (CIVA and ROLIS) andinstruments for compositional analysis(APXS, COSAC, Ptolemy) and for thestudy of physical properties (SESAME,MUPUS, ROMAP, CONSERT).

The Orbiter payload comprises 11investigations. Four instruments(ALICE, OSIRIS, VIRTIS, MIRO) provideremote sensing of the nucleus, coveringthe wavelength range from UV to sub-mm. Three (ROSINA, COSIMA, MIDAS)provide compositional andmorphological analysis of the volatileand refractory components of thenucleus. They are complemented by asuite of instruments that describe thenear-nucleus gas and dustenvironments and the coma interaction

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Rosetta Lander Scientific Instruments

APXS (Alpha Proton X-ray Spectrometer). Surface elemental composition (C-Ni) by α-particlebackscattering and X-ray fluorescence. PI: R. Rieder, MPI für Chemie, D. 1.2 kg, 1.5 W.

ÇIVA (Comet IR & Visible Analyser). 6 identical ÇIVA-P micro-cameras return panoramic surfaceimages (1 mm res under Lander); 7th camera adds stereo. ÇIVA-M 1-4 µm spectrometer studiescomposition, texture & albedo of samples collected by SD2. 40 µm res. ÇIVA-M microscope 7 µm res,3-colour images. PI: J.P. Bibring, IAS, F. 4.4 kg (with ROLIS). Participating: D. Imaging MainElectronics shared with ROLIS.

ROLIS (Rosetta Lander Imaging System). CCD camera obtains high-res images during descent andstereo panoramic images of areas sampled by SD2. PI: S. Mottola, DLR Berlin, D. 0.94 kg, 4 W.Participating: F.

CONSERT See Orbiter entry. 1.9 kg, 2 W, 1.5 Mbit/orbit.

COSAC (Cometary Sampling & Composition Expt). Evolved gas analyser identifies complex organicmolecules from their elemental & molecular composition. PI: H. Rosenbauer, MPI für Aeronomie, D.4.85 kg, 8 W, typically 3 Mbyte from 1 sample. Participating: F.

MUPUS (Multi-Purpose Sensors for Surface & Subsurface Science). Sensors on anchor, probe &exterior measure density, thermal & mechanical properties of upper 32 cm of surface. PI: T. Spohn Univ. Münster, D. 2.0 kg, 1.3 W, 3 Mbit/first science sequence. Participating: A, F, POL,UK, USA.

Ptolemy. Evolved gas analyser focuses on isotopic ratios of light elements. Gas chromatograph &mass spectrometer. PI: I. Wright, Open Univ., UK. 4.3 kg.

ROMAP (Rosetta Lander Magnetometer & Plasma Monitor). Local magnetic field (fluxgatemagnetometer) and comet/solar wind interaction (Simple Plasma Monitor). Ions to 8 MeV, electronsto 4.2 MeV. PI: U. Auster, TU Braunschweig, D. 0.7 kg, 0.9 W, 4.4 kbit/s. Participating: A, HUN, RU,USA.

SD2 (Sampling & Distribution Device). Drills up to 23 cm deep, collects samples and delivers toÇIVA-M microscope and COSAC/Ptolemy ovens. PI: A. Finzi, Politecnico di Milano, I. 4.6 kg, 6 W.

SESAME (Surface Electrical, Seismic & Acoustic Monitoring Expts). 3 instruments measureproperties of nucleus outer layers to 2 m depth: Cometary Acoustic Sounding Surface Expt (CASSE,propagation of sound); Permittivity Probe (PP, electrical properties); Dust Impact Monitor (DIM, dustfalling back to surface). PI: D. Möhlmann, DLR Köln, D; H. Laakso, FMI, FIN; I. Apathy, KFKI, HUN.1.9 kg. Participating: F, NL.

merged their efforts into the RosettaLander. The prime contract wasawarded to what was then DornierSatellitensysteme in February 1997.The launch contract with Arianespacewas signed 19 June 2001.

Spacecraft configuration:2.0x2.1x2.8 m box-shaped bus ofconventional aluminiumconstruction, with solar arrayspanning 32 m. Central thrust

The AO for Orbiter investigations andInterdisciplinary Scientists was issuedin March 1995, and the selection wasendorsed by the SPC in February1996. It included two Surface SciencePackages: Champollion byNASA/JPL/CNES and RoLand byDLR/MPAe. During the 1-year scienceverification phase, the payload wasconsolidated and, for programmaticand budget reasons, NASA withdrewfrom Champollion in September 1996.CNES and the RoLand team then

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cylinder of corrugated Al honeycombwith shear panels connecting the sidepanels. Solar wings on Y panels, HGA+X, Lander –X, science instruments+Z.

Control: 90-min round-trip light-timemeans that realtime control is notpossible, so autonomousmanagement system executs pre-loaded sequences and ensuresimmediate corrective actions in caseof anomalies. Implemented on 4 MA-31750 processors (any 2 for datamanagement & AOCS).

Attitude/orbit control: 2 sets of12x10 N thrusters using 1670 kgNTO/MMH in 2x1108-litre tanks;2200 m/s delta-V. Blowdown mode(can be repressurised twice; 4x35-litre tanks). Attitude determination to40 arcsec by 2 star trackers(16.8x16.8º FOV), 3 laser gyros, Sunsensors.

Power system: twin 5-panel steerablesolar wings provide 850/395 W at3.4/5.25 AU. 249 W required during

hibernation, 401 W active cruise,660 W at comet. 62 m2 of Si LILTcells optimised for low-intensity(40 W/m2) & low-T (-130ºC). 4x10 AhNiCd batteries. In hibernation,totalling, 2.5 yr, almost all electricalsystems are off: Rosetta spins at Sun-pointing 1 rpm, with only radioreceivers, command decoders andpower supply active (these units arehot redundant, so autonomous fault-management system not requiredactive).

Thermal control: to cope with x25variation in solar heating. 132 W ofheaters for cold operations. MLI of 2sets of 10-layers; external foil is 1-milcarbon-filled black Kapton. Coolingby 14x0.17 m2 louvre radiators.

Communications: 2.2 m-dia 2-axisHGA S/X-band up/down (28 WX-band), 1º beam, up to 64 kbit/sdown (5 kbit/s min during main

Rosetta: a History

ESA’s Horizon 2000 long-term programmewas established in 1984 with ‘A Mission toPrimordial Bodies including Return ofPristine Materials’ as one of the fourCornerstones. A returned drill core from acomet was of the highest scientific interest.The Rosetta Comet Nucleus Sample Returnmission was studied in partnership withNASA for launch in 2002 to deliver 10 kg ofsamples in 2010 to Earth. The main craftwas based on NASA’s Cassini design, withESA providing the lander and returncapsule. NASA decided in 1991 that theproject could not begin until about 2000,departing in 2005 at the earliest. ESA’s newbaseline thus became a Europe-only missionfocusing on comet rendezvous and asteroidflyby. The reference mission was originallyComet Schwassmann-Wachmann 3 andasteroid Brita but studies showed there to beinsufficient margins. Comet 46P/Wirtanenand asteroids 3840 Mimistrobell &2530 Shipka were adopted in 1994.2703 Rodari replaced Shipka in 1996.Otawara and Siwa were baselined in 1998after further studies showed them to bemore interesting objects.

Rosetta’s Structural & Thermal Model duringvibration testing at ESTEC. The 2.2 m-dia HGA is

prominent.

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‘balcony’ area is exposed, with SD2sampling system and MUPUS & APXdeployable sensors. Nine experimentstotal 21 kg, plus SD2. On command,Lander self-ejects 1 km above surfaceby 3 rotating lead screws, providingejection speed selectable 0.05-0.5 m/s (±1%). Descent stabilised by5 Nms flywheel. Single 17.5 N GN2thruster provides up to 1 m/sincrease to reduce 1 h descent timeand minimise atmospheric effects.Tripod legs deploy and damp outimpact energy. Microswitches onfootpads trigger harpoon to anchorLander to surface (GN2 thruster alsofires). Legs can rotate, tilt and lift tolevel Lander. Orbiter provides powerbefore release; at release, 700 Whprimary non-rechargeable battery+68 Wh rechargeable. At 3 AU, SiLILT cells covering 1.4 m2 provide 9-10 W during comet’s ‘day’; average5 W projected. Primary battery allowscomplete measurement cycle in 60 hirrespective of solar power. 16 kbit/sdata rate to Orbiter, allowing 13 Mbitin 15 min after touchdown. At 3-2 AU, electronics are kept warminside two 20-layer MLI tents aidedby two absorbers on hood. At 2 AU,overheating becomes the problem.

science phases). Coding & modulationoptimised for power-limited system(Turbo Code will be demonstrated bySMART-1). Fixed 80 cm-dia MGA, 9ºX-band, 30º S-band. 2 LGAs foremergency S-band. 25 Gbit massmemory. Up to 12 h/day coverage by35 m-dia ground station at New Norcia,Perth, Australia. Mission ControlCentre at ESOC, Science OperationsCentre at ESTEC & ESOC; LanderControl Centre at DLR Cologne,supported by Lander Science Centre atCNES Toulouse.

Lander: carbon-fibre structure consistsof a baseplate and a 5-panelinstrument platform under a polygonalhood that is covered by solar cells. The

DL

R

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locations, as well as the temperature,pressure and fluid velocity at 17locations. The influence of thesloshing water on the spacecraft’sdynamics will be measured bygyroscopes and accelerometers.

Total experiment time will be about24 h, spread over the Shuttle’smission duration. Although operatedfrom the ground, Sloshsat’s data willbe routed via the Shuttle. Quiescentperiods between experiment runs willallow the water to settle and thebattery to charge. Checkout andcalibration will be performed beforeejection, concluding with a few runsusing Shuttle-induced motions.

The data will validate a newgeneration of fluid motion simulationmodels. These models will betterpredict sloshing behaviour in

Achievements: first satellite dedicated to studying fluid behaviour in weightlessnessLaunch date: planned January 2003 (ready by end-2001 for flight but depends on

Space Shuttle manifest)Mission end: about 10 days after deployment (depends on Shuttle mission

duration); reentry projected after about 6 monthsLaunch vehicle/site: Space Shuttle from Kennedy Space Center, USALaunch mass: satellite 129 kgOrbit: planned 225x225 km, 51.6° (to be confirmed)Contractors: National Aerospace Laboratory (NL; prime), Fokker Space (NL;

structure, power), Verhaert (B; ejection system, ground support equipment),Newtec (B; Hitchhiker Communication System, Sloshsat radio subsystem), Rafael(IS; reaction control system), NASA (launch), Kvant (RU; solar cells)

Sloshsat-FLEVO is a small satellitedesigned to investigate the dynamicsof fluids in microgravity. Thebehaviour of water in aninstrumented tank will be monitoredto help understand how sloshingaffects the control of launchers andspace vehicles. As a joint programmebetween ESA and the NetherlandsAgency for Aerospace Programmes(NIVR), satellite development is beingperformed within ESA’s TechnologyDevelopment Programme Phase 2 andNIVR’s Research & Technologyprogramme. FLEVO is the acronymFacility for Liquid Experimentationand Verification in Orbit; it is also theregion of The Netherlands where NLRis located. Indeed, the word means‘water’.

Sloshsat will be ejected from a NASAHitchhiker Bridge in the ShuttleOrbiter cargo bay. It is part of theSlosh Test Orbital Facility (STOF),which also includes ESA’s spring-loaded ‘ESAJECT’ ejectionmechanism (designed for 50-150 kgsatellites) and the HitchhikerCommunication System (HHCS) radiocommunications package interfacingwith the HH avionics.

For 10 days after ejection, Sloshsatwill transmit data on the behaviour ofthe water in its experiment tank. Thethickness of the water near the tankwall will be measured at 270

Sloshsat-FLEVOSloshsat-FLEVO

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Above: model of Sloshsat attached to its Hitchhiker interface platform. The ‘spots’ on the tank representthe 270 platinum capacitors for measuring thethickness of water on the wall. Right: Sloshsat beingprepared for thermal-vacuum testing. Bottom:mounting sensor electronics on the tank.

microgravity, yielding more accuratecontrol of satellites carrying largequantitites of fluids. For example,docking supply vessels to theInternational Space Station,controlling launch vehicles andaccurately pointing astronomicalsatellites will all benefit from the newmodels.

The mission’s main objectives are to:• design and develop the STOF and

perform the experiment;• obtain experiment data to

verify/validate existing fluiddynamic models;

• develop and qualify a low-cost,small spacecraft bus that complieswith the Shuttle safetyrequirements.

The data obtained will be used for thefollowing main scientific objectives:• to verify the adequacy of existing

analytical fluid dynamics models;• to verify the existing Computational

Fluid Dynamics (CFD) software;• to allow the development of new

CFD numerical models;• to provide information for designing

microgravity liquid managementsystems.

The effect of sloshing on spacecraftcontrol has so far been difficult topredict for real situations.

Configuration: box-shaped bus, 91.6x74.8x96.5 cm. 86.9-litre experimenttank, cylindrical with hemispherical ends,contains 33.5-litres of deionised water andnitrogen at 1 bar. The aluminium outer tank,736 mm long, 498 mm diameter, protects theinner tank of an aramid fibre reinforcedepoxy and 2.3 mm-thick polyethylene liner.Sensors: 270 capacitors (coarse waterthickness on wall); fine liquid thickness (3locations); liquid velocity (10); liquid pressure(1); liquid temperature (3). The tank ismounted on the equipment platform, whichhosts power, data handling, control, etcsystems. Sloshsat’s structure is aluminium.

Power/thermal control: body-mounted Sisolar cells on five sides provide 35-55 W,supported by 5.1 Ah/143 Wh NiCd batteryfor eclipses. 28 Vdc provided throughESAJECT eject system when attached to HH.Internal 0-70ºC maintained by multi-layersinsulation blankets, heaters & coatings.

Motion control: 12x1.1 N blowdown thrustersfed by four nitrogen tanks (1.6 kg N2 at600 bar) provide linear/rotational movementto excite fluid motion. (Corresponds tomaximum linear acceleration capability of0.0186 m/s2 for 432 s.) Three Litton-Liteffibre-optic gyros (±98º/s) and six AlliedSignalQA-3000-10 accelerometers (±0.030 &±1.5 g) monitor Sloshsat movements. Noattitude or orbit control. 20-95 kmseparation maintained by Shuttle.

Communications: when available, Shuttle Ku-band 16.3 kbit/s downlink used for realtimerelay of bulk data. A dedicated Payload &General Support Computer (a standard IBMlaptop in Shuttle cabin) stores data fordumping via Shuttle Ku-band after runs arecompleted. An S-band 1.2 kbit/sup/downlink is used for non-realtimeexperiment control and realtime experimentobservation. Spacecraft control by PayloadOperations Control Center at NASA Goddard.

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X-38 & CRX-38 & CRVVPlanned achievements: first ESA participation in development of

manned spacecraftLaunch date: X-38 space test February 2003 (STS-116); CRV #1 2007Mission duration: 3 years (CRV)Launch vehicle/site: NASA Space Shuttle from Kennedy Space Center, FloridaLaunch mass: CRV about 10 000 kg Orbit: as Space Station (about 400 km, 51.6°)Principal ESA contractors: MAN Technologie for X-38; with Alenia Spazio

(co-primes) for CRV

In a highly successful partnership,NASA, ESA and European industryare building the X-38 spacecraft, theprototype of the Crew Return Vehicle(CRV) for the International SpaceStation. Once attached to the Stationin 2007, the CRV will provide a routehome for the 7-strong crew if anyonerequires specialist medical attention,the Station becomes uninhabitable orit cannot be resupplied.

ESA is responsible for 15 X-38subsystems or major elements, plusengineering expertise. Subscriptionsby ESA Member States to the X-38programme is allowing a significantrole in the development andproduction of reusable spacetransportation systems. Atmosphericreentry technologies must bemastered for future cost-cuttingreusable space transportation – thesignificant technical knowhow ofEurope used in and gained throughthis programme can be applied toESA’s Future Launcher TechnologyProgramme and beyond. The basicX-38/CRV configuration offersinherent capabilities for future roles,such as a mini-shuttle or an in-orbitferry between vehicles.

The X-38 programme is using a rapiddevelopment approach ofincrementally designing, building andtesting essential systems andtechnologies. It is a new way ofreducing the cost of developingmanned spacecraft by an order ofmagnitude.

The X-38 is a lifting body with adisposable deorbit stage. This liftingbody adopts the heritage of the USX-23/X-24 vehicles developed andflight tested in the 1960-70s, but it isenlarged and significantly modified byESA to improve the flyingcharacteristics and volume efficiency.This design’s entry crossrangecapability allows a short orbital flightto assure a landing at specific groundsites irrespective of the time ofdeparture from the ISS. The briefflight results in simpler, more reliablevehicle systems.

The X-38 programme is using fiveprototypes: four atmospheric drop-test vehicles (V131, V132, V131R,V133) and a space test vehicle (V201).V131 had the X-24 shape, 7.32 mlong, with the primary goal of

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demonstrating the transition fromlifting body to parafoil flight. Thecontrol surfaces were fixed. It flewtwice in 1998-99 before it wasrefurbished as V131R to reflect themodified CRV shape, including theberthing/docking mechanism on thetop and the 680 m2 full-size parafoil.V132 has the original shape andscale, with the primary goal ofdemonstrating the flight controlsystems using Electro MechanicalActuators and advanced controlsoftware technology during its threeflights, in 1999 and 2000.

V133 will be the same size and shapeas the CRV, with the primaryobjective of verifying the aerodynamicshape modifications as well as thecontrol laws.

V201 is the space version, to bedeployed from Space ShuttleColumbia in 2003 for a full-up entrytest. It has the modified shape scaledto CRV’s 9.1 m. In order to fit into theOrbiter’s cargo bay, the upper partsof the fins are foldable. V201 is underassembly at the NASA Johnson SpaceCenter (JSC).

Initially a NASA in-house project, theX-38 programme developed into a full

partnership between ESA and NASA:ESA is responsible for the design anddevelopment of 15 major subsystemsor elements of the V201 spacecraft.ESA’s work is being performed undera firm fixed-price contract with MAN-Technologie (D) leading a team of 22industrial companies in eightcountries. ESA’s contributionsinclude:

– vehicle shape validation and overallaerodynamic and aerothermo-dynamic database;

– crew cabin design and layout;– aft fuselage design and manufacture

of major aft structure elements;– rudders, including accompanying

sensors;– the metal nose structure;– the front and main landing gear;– the cabin equipment pallets;– hot structure (Ceramic Matrix

Composite, CMC) leading edgesegments of the fixed fin, includingaccompanying sensors;

– the Thermal Protection System (TPS)blankets for the leeward vehiclesurfaces, including fins and aftfuselage frame;

– guidance, navigation and control(GNC) software, includingman/machine interfaces, for theparafoil flight phase;

– Fault Tolerant Computers withreentry GNC software;

– Vehicle Analysis and Data RecordingSystem (VADRS), including front-end electronics for overall vehicleinstrumentation;

– predevelopment of the CRV/ISSberthing/docking mechanism;

Preparing X-38 V131R for its maiden flight.

ESA’s GNC softwareguides X-38 V132 toits second landing,9 July 1999.

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– active thermal control waterpump;

– crew seat concept for the CRV andprovision of a representative crewseat with instrumented dummy forX-38 flight testing.

The ESA/NASA partnership iscomplemented by DLR’s TETRA(TEchnologie für zukünftige Raum-TRAnsportsysteme) programme,which is developing essential V201elements, specifically the CMC bodyflaps and the CMC nose capassembly.

The X-38 cooperation isunprecedented. It is the first timethat NASA, ESA and ESA contractorstogether are developing the prototypefor a reusable operationalspaceplane. The development ofessential systems and technologiesfor a reusable, reentry vehicle is afirst for Europe, and sharing thedevelopment of an advanced reentryspacecraft with foreign partners is a

first for NASA. It is remarkable thatNASA selected ESA’s aerodynamicshape over its own.

The fast pace of the X-38 programmewould not have been possible withouteffective decision-making processesand the collocation of industrialpartners in Europe and at JSC. ESA’sX-38 team has included engineersfrom 11 industrial firms working inan integrated team at JSC.

The V201 space test vehicle will flythe entire operational entry mission(speeds, altitudes, attitudes) of theoperational CRV. Much of this flightregime was covered by the three X-23space tests (1966-67), and once belowMach 2 and 25 km altitude it isoperating in the X-24 database (28flights 1969-75). When V201 entersthe subsonic region, it will be flyingthe same trajectory that atmosphericvehicles have followed many times.Once the drogue parachute isdeployed and has decelerated the

Left: X-38 V131R descends on itssecond flight, 10 July 2001.Bottom left: X-38 V201 in assemblyat JSC. (NASA)

CRV and European Industry

Austria MAGNA: foldable finBelgium SONACA/SABCA: aft structure

SAS/Spacbel: software independent validation, displays & controls/Man-Machine Interface

Verhaert: trunnion mechanism, IBDMFrance AML: aerodynamics

Dassault: aerodynamics/aerothermodynamics

ONERA: aerodynamicsGermany Astrium: parafoil GNC, software

independent validationDLR: CMC nose thermal protection,

aerodynamics/aerothermodynamicsMAN Technologie: industrial lead,

body flap assemblyItaly Aermacchi: aerodynamics

Alenia Spazio: industrial lead, nose primary structure

CIRA: aerothermodynamics, Scirocco plasma facility

SICAMB: crew seatsNetherlands Fokker: rudders

NLR: aerodynamicsSpain Sener: landing gear, IBDMSweden FFA: aerodynamicsSwitzerland Contraves: fin folding mechanism

(status February 2001)

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ESA’s participation inthe CRV programme (status February 2001).

An ESA and Belgian industry team is developing CRV’scockpit control and display systems. CRV’s flight willnormally be automatic but two astronaut ‘operators’will monitor progress and can intervene via keypadsand handcontrollers. The current concept provides sixmulti-functional screens – two per operator and twoshared. The CRV’s orbital position, deorbit burnopportunities and landing site positions are displayedon a continuously updated moving world map. Detailedinformation about each landing site can be called up.

vehicle, the rest of the flight will beunder the same conditions as alllarge parafoil deployments in theatmospheric vehicle and pallet drop-test programmes (approximately 90tests). These tests initially used a485 m2 parafoil and later the full-size680 m2 parafoil.

The CRV has three main elements:the lifting body entry vehicle, theCRV/ISS International Berthing/Docking Module (IBDM) and theDeorbit Propulsion Stage (DPS). ACRV will be delivered by the SpaceShuttle and remain attached to theISS for up to 3 years. It willaccommodate 0-7 crew members in ashirtsleeve environment, and provideautonomous flight and landing. Themaximum mission duration (for anemergency departure) is 9 h, andabout 3 h in the case of a medicalreturn. The CRV must be able toseparate from the Station at anyattitude with a tumbling rate of up to2º/s. The maximum sustained entry

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g-load will not exceed 4 g along thelongitudinal axis. The DPS thrustersare fired to initiate the descent, andthe module is jettisoned. The vehicleenters the atmosphere at an altitudeof about 120 km, travelling at27 000 km/h. Attitude is controlledinitially with nitrogen cold-gasthrusters but, as air pressureincreases, the rudders and body flapstake over. A drogue parachutedeploys at 8 km altitude, stabilisingthe vehicle in a 1 g sustained sinkrate. This is followed by the 5-stagedeployment of the 680 m2 parafoil.Automatic GNC software steers theCRV through its final descent andlanding. The landing accuracy is≤ 9 km radius and the landing speed≤ 9.1 m/s horizontally and ≤ 4.6 m/svertically. As an emergency vehicle,CRV must be able to land on a widerange of unprepared surfaces. Itskids to a halt within a shortdistance on three gravity-deployedtitanium legs using crushablecartridges to cushion the impact. Thelanding mass will not exceed10 000 kg.

ESA and NASA agreed in 1997 toextend the X-38 partnership to theCRV. These early agreements atprogramme-management level werefollowed up by the NASA/ESAProtocol on X-38/CRV Cooperation,signed in November 1998. At the May1999 ESA Ministerial Conference inBrussels, it was decided to link aEuropean contribution to the CRVprogramme with ESA’s ISSexploitation phase committments:

ESA is negotiating a BarterAgreement to receive NASA-providedStation services such astransportation and high data-rateservices. The scope of ESA’sparticipation goes beyond the X-38partnership, and includes additionalsubsystems or elements, such as thefoldable fins, the fin-foldingmechanisms, the trunnionretraction/extension mechanisms,the crew seats and the hot structurebody flaps and nose TPS.

CRV development is in two phases:Phase-1 includes the design activitiesup to the Critical Design Review;Phase-2 will include the production offour operational CRVs, including theDPS and the CRV Berthing Adapter(CBA) with the IBDM.

With the full start of Phase-1projected for late 2001, the firstoperational CRV is expected to be atthe Station before mid-2007. Untilthen, astronauts have to rely onRussian Soyuz capsules inemergencies and the crew size islimited to three.

ESA began early Phase-1 activities inDecember 1999: aerodynamics andaerothermodynamics; qualificationactivities for the CMC material, forthe hot structure body flaps, ruddersand nosecap TPS; display techniquedevelopment and man-machineinterfaces; design activities for theIBDM; and system and subsystemengineering as part of the integratedESA/NASA team at JSC.

The CRV departing from Node-3 of the Space Station. The DPS fires to begin the descent.

CRV’s 11.8 m3 cabin can carry seven passengers. CRV reentry.

Following the approval of theconsolidated programme by ESA’sManned Space Programme Board, theRequest for Quotation was issued toindustry in early October 2000.Phase-1 will be synchronised withNASA’s industrial Phase-1. Europe’sindustrial team has evolved from theX-38 team, with MAN Technologieand Alenia Spazio sharing the role ofprime contractor, and leading a teamof 19 subcontractors in the ninecontributing nations (see table). Bymid-2001, Member Statecontributions were approaching acomfortable €150 million. Countriesvowed to increase funding no laterthan the end of Phase-1 in order tosecure a European role throughoutCRV’s operational phase.

Full utilisation of the ISS requires acrew of 7 and hence the CRV. ESAand NASA, in March 2001, agreed inprinciple on a substantially expandedEuropean role in the CRVprogramme. This agreement has to beformalised and approved by the nineparticipating states. Financing wouldbe through ESA’s commitments inthe ISS exploitation programme,together with additional funding,motivated by the high technologycontent of the CRV programme andits synergies with future reusablespace transportation systems.

The potential expanded role and leadresponsibilities for Europe in CRVdevelopment offer a uniqueopportunity for ESA and Europeanindustry to enlarge its technology and

system engineering base significantlyfor atmospheric reentry and thereforefor reusable space transportationsystems – manned or unmanned.The ESA/NASA agreed potentialscope includes the system andsubsystem engineeringresponsibilities, including interfacemanagement and ground and flightoperations support for all ESA leadfunctions. These include thecomplete mechanical assembly of thevehicles, including: all the primarystructure, all control surfaces (CMC),hot structure (CMC) nose and leadingedges, TPS blankets, thermal controland life support, equipmentaccommodation in the cabin, crewseats, the IBDM system, CRV/ISSBerthing Adapter, trunnionmechanism and fin-foldingmechanism. It includes the cold-gasattitude control subsystem (ACS) andthe complete DPS. It increases thescope of the aerodynamics workthrough additional CFD, wind tunneltesting and the transonic flighttesting of V131R in Europe. Avionicscould include the architecture, fault-tolerant design, all data busses,MDM equivalents, power controlequipment, reaction control systemfor ACS and DPS, integrated healthmanagement, avionics qualification,independent software evaluation,GNC for the parafoil phase, displays& controls, instrumentation, solid-state recorders, audio, video &lighting, and system engineering inmission planning, operations conceptdevelopment, and ground supportsoftware and processing.

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