BepiColombo, ESA's Mercury Cornerstone mission

Planetary and Space Science 49 (2001) 1409–1420www.elsevier.com/locate/planspasci

BepiColombo, ESA’s Mercury Cornerstone mission

Alberto Anselmia ;∗, George E.N. Scoonb

aAdvanced Studies Department, Alenia Spazio S.p.A., Strada Antica di Collegno 253, 10146 Torino, ItalybESA Scienti#c Projects Department, European Space Research and Technology Centre, P.O. Box 299, 2200 AG Noordwijk ZH, The Netherlands

Received 18 July 2000; accepted 12 January 2001

Abstract

The paper presents the results of the de0nition studies performed for the European Space Agency (ESA) on system architectures andenabling technologies for “BepiColombo”, a Cornerstone class mission to be launched in the 2007–2009 time frame. The scienti0c missioncomprises 1-year observations by a Mercury Planetary Orbiter (MPO), dedicated to remote sensing, and a Mercury Magnetospheric Orbiter(MMO), dedicated to particles and 0elds, plus short-duration in situ analysis by a Mercury surface element (MSE). A 6exible approachto the programme has been developed, comprising two alternative launch scenarios. In the 0rst option (2009), the 2500-kg class satellitecomposite, including two propulsion modules and three scienti0c modules, is launched by an Ariane-5. The trajectory design is basedon Venus and Mercury gravity assists plus the thrust provided by a Solar Electric Propulsion Module (SEPM), that is jettisoned beforebeing captured into Mercury orbit. Capture and orbit insertion, executed by successive manoeuvres of a Chemical Propulsion Module(CPM), occur less than 2:5 yr after launch. In the second scenario, the mission is split into two launches of a small launch vehicle. Two1200-kg class composites are launched either in the same one-month window or at an interval of 1:6 yr. One composite comprises theSEPM, CPM, MMO and MSE and the other comprises duplicate SEPM+CPM and the MPO. The trajectory design follows the sameprinciples as the Ariane-5 mission, with the SEPM thrust reduced by half and cruise duration ranging between 2.3 and 3:5 yr. Whateverbe the implementation, the mission is expected to return about 1700 Gbit of scienti0c data during the one-year observation phase. Thecrucial aspects of the spacecraft design are associated with, and constrained by, the high-temperature and high-radiation environment.Basic feasibility has been demonstrated by an extensive design and analysis exercise, and the focus of the programme has now moved toa 3-year preparatory programme dedicated for developing the enabling technologies. c© 2001 Elsevier Science Ltd. All rights reserved.

1. Introduction

The BepiColombo mission to planet Mercury is understudy by ESA as a “Cornerstone” of its “Horizons 2000”scienti0c programme plan. 1 It is designed to improve ourunderstanding of the planet and its environment through awide variety of techniques such as remote sensing, particleand 0eld measurements, and radio science (Grard et al.,2000).

The 0rst system and technology study for ESA’s Corner-stone mission to Mercury was performed by the industryin 1998–1999 (Scoon et al., 1999). The system architecturefeatures three spacecraft elements used for planetary

∗ Corresponding author. Tel.: +39-011-7180-718; fax: +39-011-7180-319.

E-mail address: [email protected] (A. Anselmi).1 After submission of this manuscript, in October 2000, ESA’s Science

Programme Committee selected BepiColombo as the 5th Cornerstone ofHorizons 2000, to be launched in 2009 in collaboration with Japan.

exploration:

• three-axis stabilised, nadir-pointing Mercury PlanetaryOrbiter (MPO), dedicated to remote sensing observationsand radio science,

• spin-stabilised Mercury Magnetospheric Orbiter (MMO),dedicated to particles and 0elds investigations,

• short-lived Mercury Surface Element (MSE), dedicatedto in situ geochemical analysis.

The mission scenario elaborated in that study is basedon a direct launch by Ariane-5. The trajectory design in-cludes Venus and Mercury gravity assists in addition to thethrust provided by a large Solar Electric Propulsion Mod-ule (SEPM) that is jettisoned before being captured intoMercury orbit. Capture into Mercury orbit is executed bysuccessive manoeuvres of a Chemical Propulsion Module(CPM). The cruise to Mercury takes less than 2:5 yr.

More recently, a new mission design was elaborated inwhich the 0ve elements are accommodated in two separate

0032-0633/01/$ - see front matter c© 2001 Elsevier Science Ltd. All rights reserved.PII: S 0 0 3 2 -0 6 3 3 (0 1 )00082 -4

1410 A. Anselmi, G.E.N. Scoon / Planetary and Space Science 49 (2001) 1409–1420

Nomenclature

The mission scenario in which the three scienti0cmodules (MPO, MMO, MSE) are launched togetherby an Ariane-5 is henceforth referred to as the“single-launch scenario”. The mission option usingtwo separate launches of Soyuz–Fregat is called the“split-launch scenario”. For the latter option, we willlabel the MMO+MSE launch as “Launch 1” and theMPO launch as “Launch 2”, but the MPO compos-ite can be launched 0rst provided that the technologydevelopment (more demanding for the MPO) is com-pleted in time.

launches of Soyuz–Fregat. The 0rst launch can take placeas early as 2007, with other opportunities in 2009 and 2010.The second launch can occur in 2009 with another oppor-tunity in 2010. The launches can occur in either the samewindow, within one month of each other, or in two succes-sive windows. One launch carries a composite of the SEPM,CPM, MMO and MSE. The other launch delivers duplicateSEPM and CPM modules, and the MPO. The SEPM andCPM are nearly identical in the two launches, and approx-imately half the size of the corresponding modules in thesingle-launch scenario. As a result of the new study, the sci-enti0c modules are identical in all mission options (Arianeas well as Soyuz–Fregat based). The interplanetary trajec-tory design was tailored to the new mission scenario. A lu-nar gravity assist was included in the early mission phase,improving the mass margins available to the Soyuz–Fregatlaunch at the cost of 1 yr longer cruise (3:5 yr). The origi-nal short cruise of less than 2:5 yr will become possible ifplanned improvements in the Soyuz–Fregat vehicle and fa-cilities are implemented before 2007=2009.

The Ariane-5 based mission design also remains a viableoption. Launch mass margins are optimal, much larger than20%. Furthermore, the “split” composites were demon-strated to be fully compatible with a shared launch byAriane-5 with the Speltra adapter with mass margins aslarge as in the single-composite launch option. The studyresults demonstrate the ample 6exibility of launch optionscompatible with the mission requirements, and the capabil-ity of the mission to be adapted to varying programmaticconstraints.

2. Scienti�c payload

The scienti0c programme is performed by three satelliteelements sharing the same carrier to Mercury, where theyare deployed to their speci0ed orbits (or destination in thecase of the Mercury Surface Element).

The MPO element is dedicated to remote sensing obser-vations. Table 1 shows the reference payload. The MPO

requires a polar orbit for complete coverage of the planet;an orbit with 400 km pericentre and 1500 km apocentre, se-lected partly as a compromise with engineering constraints,provides proper ground track shifting with an orbital periodof 2:3 hr. Low-resolution observations from the apocentrewill identify areas of interest, while observations around thepericentre will serve for the higher-resolution studies. Allviewing instruments share a requirement for nadir pointing.Radio Science experiments dedicated to resolve Mercury’slibration state and to map its gravity 0eld are also includedin the mission concept (Iafolla et al., 2000; Iess, 2000; Mi-lani et al., 2000). They require concurrent operation of theMPO’s systems and payload (communications subsystem,star sensors, cameras, and accelerometer).

The MMO element is devoted to 0elds and particlescience (Table 2). The orbit for these goals is highly el-liptical and polar, with a pericentre/apocentre altitude of400 km=11; 800 km (∼ 5 Mercury radii), period of 9:3 h,and the line of apsides in the equatorial plane. The orbitis eJectively inertial, and thus not only covers the wholemagnetosphere but also scans the distant tail once per Mer-cury year. A spinning satellite (15 rpm) is best suited to theexperiments envisaged.

The MSE is conceived as a short-lived probe to studyin situ the physical properties of the surface. It is deployedfrom the MPO (in the single-launch scenario) or the MMO(in the split-launch scenario) into an impact trajectory toa high-latitude landing site where the environmental condi-tions are less severe than in lower latitude regions. A heat6ow and physical properties package is placed 2 to 3 m be-low the surface by either a penetrator driven into the soil bythe impact speed on landing or a self-propelled penetrationdevice (“Mole”). An X-ray spectrometer is mounted on amicro-rover allowing samples to be analysed at diJerent lo-cations in the landing zone. Two micro-cameras are mountedon the MSE body to provide images during the descent andon-site. A magnetometer and a seismometer complete thereference payload complement (Table 3).

3. Mission design

The approach selected for the mission requires the deliv-ery of masses of the order of 1000–1200 kg to the orbit ofthe MMO. Three scenarios have been considered (Langevin,1999):

• single launch (Ariane-5) with a solar electric propulsionmodule,

• single launch (Ariane-5) with a chemical propulsionmodule,

• split launch (Soyuz–Fregat) with a solar electric propul-sion module (with or without lunar swing-by).

Both the chemical and the solar-electric-propulsion (SEP)missions use gravity assists from Venus. The launch win-dows to this planet are spaced at intervals of 1:6 yr. Windowsin January 2009, August 2010 and March 2012 have been

A. Anselmi, G.E.N. Scoon / Planetary and Space Science 49 (2001) 1409–1420 1411

Table 1MPO reference payload complement

Instrument Range Mass Average power Average telemetry rate(kg) (W) (kbit=s)

Narrow angle camera 350–1050 nm 4 — 20Wide angle camera 350–1050 nm 4 — 20Common camera electronics — 4 16 —IR spectrometer 0.8–2:8 �m 6 10 1.5UV spectrometer 70–330 nm 3.5 3 2X-ray spectrometer 0.5–10 keV 4 5 0.1MXS Peltier cooler — 0.5 3 —Gamma-ray spectrometer 0.1–8 MeV 3 2 0.1MGS Stirling and Peltier coolers — 4.5 3. —Neutron spectrometer 0.01–5 MeV 5 3 0.05Radio science — — — —

• transponder 32–34 GHz 3.5 9 —• accelerometer 10−4–10−1 Hz 8 6.3 0.1

Table 2MMO reference payload complement

Instrument Range Mass Average power Average telemetry rate(kg) (W) (kbit=s)

Magnetometer ±4096 nT 0.88 0.35 0.8Ion spectrometer 50 eV–35 keV 4.4 4 0.5Electron analyser 0 eV–30 keV 1.1 1.2 0.1Cold plasma detector 0–50 eV 1.3 1.9 0.2Energetic particle detector 30–300 keV 1.2 0.7 0.04Search coil 0.1–1 MHz 1 — —Electric antenna 0.1–16 MHz 2.9 — —Common electronics — 1.2 4 1Central interface unit — 2.1 1 0.02Positive ion emitter 1–100 �A 2.7 3.8 0.02Camera 350–1000 nm 8 12 4

Table 3MSE reference payload complement

Instrument Deployment Mass Average power Typical data volume(kg) (W) (Mbit)

Heat 6ow and physical properties Mole or penetrator 1 0.3 1packageAlpha X-ray spectrometer Micro-rover 0.8 1 1Camera None 0.2 3 50Magnetometer Articulated arm 0.51 0.6 7

(optional)Seismometer None 0.875 0.6 7Mole — 0.4 5 —Micro-rover — 2 3 —

considered for the single launch options. The split launchescan take place either in the same window or in two succes-sive windows. In the latter case, launches in May 2007 andJanuary 2009 (or about 6 months later, with lunar swing-by)have been considered. A signi0cant drawback of the chem-

ical option is the long cruise time (6 yr or more) comparedwith the SEP options (2.3–2:5 yr without lunar swing-by,3:5 yr with lunar swing-by). The combination of SEP withgravity assists provides a set of mission opportunities withshort cruise times and outstanding programmatic 6exibility.


Table 4Mission opportunities with a single launch by Ariane-5 (with restartable upper stage). Both ion and SPT propulsions are considered

Launch epoch, propulsion type 2009, Ion stage 2009, SPT stage 2010, Ion stage 2010, SPT stage

Launch date 2008=12–2009=01 2008=12–2009=01 2010=08 2010=08Launch velocity and declination 2:8 km=s, 7◦S 2:8 km=s, 7◦S 2:8 km=s, 7◦S 2:8 km=s, 7◦SSEP OV 6:95 km=s 7:00 km=s 6:82 km=s 6:87 km=sThrust time at 0:6 N 6700 h 7200 h 6600 h 7100 hCruise time 2:64 yr 2:64 yr 2:22 yr 2:22 yrApproach velocity 370 m=s 368 m=s 385 m=s 380 m=s

Table 5Mission opportunities with Soyuz–Fregat (split launch). Ion propulsion is assumed

Launch date 2007=05 2008=12–2009=01 2010=08 2009=07 (Lunar swing-by)

Launch velocity and declination 2:3 km=s, 37◦N 2:3 km=s, 4◦S 2:3 km=s, 6◦S Not applicable (50 RE)SEP OV 7:42 km=s 7:52 km=s 7:41 km=s 7:24 km=sThrust time at 0:17 N 1100 h 1560 h 1290 h 1390 hThrust time at 0:34 N 6370 h 6230 h 6270 h 6730 hCruise time 2:31 yr 2:65 yr 2:28 yr 3:30 yrApproach velocity 290 m=s 360 m=s 360 m=s 365 m=s

Fig. 1. Ecliptic projection of Earth-to-Mercury trajectory—split launch2009 with lunar swing-by. Thicker arcs identify thrust phases; stars areplanetary encounters.

There is one launch opportunity every 1:6 yr, either with asingle Ariane-5 launch or with two Soyuz–Fregat launchesfrom Baikonur, with or without lunar swing-by. The massmargin is largest with Ariane-5 and an ion stage, but ade-quate mass margins can be obtained in all con0gurations.The SEP option has therefore been selected as the baselinefor the BepiColombo mission. Table 4 summarises the mainparameters of the single-launch option with solar electricpropulsion. Table 5 shows the corresponding data for thesplit-launch option. Fig. 1 shows an example of the trajec-tory. The SEP technology is mature for the Mercury appli-cation. DS1, a NASA mission, has successfully tested SEP

in space on a modest scale. SMART-1, approved by ESA in1999 and to be launched in late 2002=early 2003, will val-idate all system aspects of a mission associating SEP withgravity assists (Racca and Marini, 2000).

The Ariane-based missions (Table 4) are compatible withany of the electric thruster designs currently being developedin Europe (i.e., ion thrusters as well as stationary plasmathrusters, SPT). The thrust level is constant, 0:6 N, realisedby simultaneously operating (at 75% of maximum thrust)four 200 mN thrusters placed on the corners of a square,symmetric with respect to the centre of mass. A 0fth cen-tral thruster provides cold redundancy. In case of failure ofany of the outer thrusters, three diagonal thrusters are usedat maximum thrust (3×0:2 N), without any variation of themission pro0le. The thrust times are atleast 30% smaller thanthe demonstrated lifetime. For double failures, the same re-covery strategies are applied as in nominal or single-failurecases of the split-launch con0guration, discussed below.Since the thrusters are 0rst turned on at 0:6 AU distancefrom the Sun, the installed power demand at 1 AU is rela-tively low (6:5 kW for SPT, 10:5 kW for ion thrusters).

In the split-launch options (Table 5), two thrust levels,0.17 and 0:34 N, are used, and they are achieved by turningon either a single central thruster or two symmetric lateralthrusters. The lower thrust level is employed in the earlypart of the cruise, when the solar array power is not yet suf-0cient to feed both thrusters. Due to the limited capabilityof Soyuz–Fregat, the ion thrusters are the sole candidates.The required power is 5:5 kW at 1 AU. Failure of the cen-tral thruster does not aJect the mission pro0le. Failure ofone of the outer thrusters (or both of them) is met by us-ing the central thruster at 0:2 N, with a penalty on thrusthours and larger the cruise duration sooner the failure occurs.The worst case is 11,000 thrust hours (marginally above the


demonstrated lifetime) and 4.3 cruise years, when the failureoccurs at the very start of thruster operation. Double failuresinvolving the central thruster can be recovered by using theCPM propulsion, provided they occur at a later stage of thecruise. In that case, however, the nominal destinations ofthe scienti0c modules can no longer be reached.

As the spacecraft approaches the Sun, the power deliveredby the array increases, and so does the array temperature.Consistent with the design requirement that the SEP mod-ule must be built from commercial components, the max-imum operational temperature is 0xed at 150◦C. After thespacecraft has reached a distance at which T = 150◦C, thearray is progressively tilted away from the Sun. At Mercurydistance, 0:32 AU, the array is tilted about 65◦. Using thisstrategy, the power delivered is relatively constant, aroundtwice the 1 AU power, from 0:6 AU inwards.

The 0nal stages of the mission from Mercury orbit inser-tion are performed with chemical propulsion. Manoeuvringin Mercury orbit would expose the solar array to high in-frared 6uxes from the planet as well as to solar radiation.This would require a complex and long (several months)approach strategy to keep the array temperature within atolerable envelope. Hence, the SEPM is jettisoned close toMercury and a more eQcient 400 N bipropellant engine per-forms the insertion.

In the single-launch scenario, the MMO orbit is reached0rst (400 × 11800 km, polar, line of apsides in Mercury’sequatorial plane). Then the MMO is released and anotherburn of the 400 N engine lowers the apocentre to 1500 km,as required by the MPO. In the split-launch scenarios, a sim-ilar approach is used to provide the MMO orbit (Launch 1)and the MPO orbit (Launch 2). In all scenarios, the Chem-ical Propulsion Module, too, is jettisoned after the acquisi-tion manoeuvres have been completed.

The arrival conditions are constrained such that the op-erational orbits have their pericentre in the anti-solar direc-tion at perihelion, which creates a more benign environmentfor the thermal control. Moreover, the MPO and MMO or-bits are resonant (4 : 1), so that there is an opportunity forbackup inter-orbiter communications should one of the twohigh gain antenna systems fail to perform according to thespeci0cations (Fig. 2).

The MSE is dormant through the cruise and is deliveredto its destination from either the MMO orbit (split launch1) or the MPO orbit (single launch). The MSE is targetedto a landing site at 85◦ latitude, where its design life is oneweek. The design life of the MMO and the MPO is one year(four Mercury years, two Mercury days).

4. Spacecraft composite

The satellite design concept is thoroughly modular in or-der to accommodate without con6ict the requirements aris-ing from various environments and mission scenarios. The0ve modules are combined in diJerent ways according to

5000 0 5000 1 104

1.5 104

1 104

5000

0

5000

1 104

X (km)

Z(k

m) Direction of the sun

at perihelion

MMO orbit

MPO orbit

Fig. 2. MPO and MMO orbit geometry.

the requirements of each mission scenario. The scienti0cmodules (MMO, MPO, and MSE) are designed to 0t eitherthe single-launch or the split-launch scenario with minimalchanges. The propulsion modules (SEPM, CPM) for thesingle-launch scenario are roughly twice the size of thosedesigned for the split-launch option. In the split-launch op-tion, the propulsion modules for the 0rst and second launchare nearly identical, thus oJering the advantages of recur-rent module procurement.

4.1. Single-launch system con#guration

In the single-launch scenario, all modules are deliveredby one dedicated launch of an Ariane-5 (Fig. 3). The SEPpropulsion includes four (+1 cold-redundant) 200 mNthrusters. Either ion or SPT thrusters can be accommodated.The solar array employs GaAs cells with 20% optical solarre6ectors, in two wings (Fig. 4). Throughout the interplan-etary and early-Mercury orbit phase, the command and con-trol tasks are centralised in the MPO, providing overall mon-itoring and control tasks, and managing telecommunicationsto the Earth. In Mercury orbit, the MPO and MMO are inde-pendent, and the MSE data relay is managed by the MPO.

Table 6 presents the mass budget of the launch con-0gurations with the ion thrusters and the SPT thrusters.The availability of Ariane-5 with a restartable upper stage(Ariane-5V) has been assumed, which provides ¿ 20%mass margin regardless of which thrusters are used. Thetrade-oJ between thruster types is further discussed inSection 4.3.

4.2. Split-launch system con#guration

In the split-launch scenario, one launch of a Soyuz–Fregatvehicle delivers a composite made of SEPM + CPM+


Fig. 3. Ariane-5 single-launch con0guration.

MMO+MSE, while another launch delivers a composite ofSEPM + CPM + MPO. The SEPM includes three 200-mNion thrusters and a 5:5 kW GaAs cell solar array in two wings(Fig. 5). Throughout the interplanetary and early Mercuryorbit phase, the command and control tasks are centralisedin the respective Orbiters. The MMO manages the MSE datarelay in the 0rst of the two split-launch missions.

Table 7 shows the mass budgets for launches fromBaikonur and Kourou in the 2009 launch window. Massmargins for direct launch from Baikonur are low (4%).Kourou (latitude 5◦N) is better suited for the 2009 launcheswith a low asymptote declination (5◦S). Should a Soyuz

Table 6Ariane-5 single-launch mass budget summary (2009 launch)

Launch mass budget (kg) Ion thrusters SPT thrustersAriane-5 single-launch con0guration (2009)

MMO 165.3 165.3MSE 234.3 234.3MPO 357.3 357.3CPM dry 167.4 167.4Subtotal 1 (dry mass at Mercury) 924.2 924.2Bipropellant 336.5 336.5Subtotal 2 (mass after jettison) 1260.8 1260.8SEPM Dry 665.1 622.5Subtotal 3 (mass before jettison) 1925.8 1883.2Cruise propellant 477.3 807.4Launch vehicle adapter 92.0 92.0Launch mass 2495.1 2782.6Ariane-5 limit launch mass 3500.0 3500.0System margin (kg) 1004.9 717.4System margin (%) 40 26

Fig. 4. Ariane-5 single-launch cruise con0guration.

launch pad at Kourou be available in 2009, the launchmass margins would reach ∼ 20%. Similar margins can bereached by a moon-bound launch from Baikonur, followedby a lunar swing-by.

The split spacecrafts are also compatible with a sharedAriane-5 launch, using the Speltra adapter, with the samemass margins as in Table 6.


Fig. 5. Soyuz–Fregat split-launch cruise con0gurations (left: MMO+MSE,right: MPO).

4.3. Solar-electric-propulsion (SEP) module

Three candidate thruster designs were considered inthe study. Radiofrequency Ion Thrusters (RITA-XT, de-veloped by Astrium-Germany) and Electron Bombard-ment Thrusters (T6-IPS, developed by Astrium-UK andDERA) have similar characteristics, while StationaryPlasma Thrusters (SPT-140, developed by SNECMA andAstrium-France) have signi0cantly lower speci0c impulseand power demands. In the single-launch mission option,any of the three thruster types can be accommodated; inthe split-launch option, due to the reduced capability of theSoyuz–Fregat vehicle, only the ion thrusters with grids canbe employed. A signi0cant 0gure of merit for SEP systemsis the speci0c mass of the power system (ratio of propul-sion system mass including thrusters, solar array, powerequipment and cabling to peak power). The planned spe-ci0c mass of the ion-thruster based SEP systems consideredfor BepiColombo is 53 kg=kW.

The SEPM is a simple rectangular prism. A central thrustcone, housing the Xenon propellant tanks, is the main struc-tural element, and transmits the loads to the launcher inter-face. Two solar array wings, with multiple panels equippedwith GaAs cells, provide the cruise power. On each wing, asolar array drive mechanism provides one rotational degreeof freedom around the yoke axis, suQcient for the array ori-entation needs during the cruise. The propulsion drive andpower conditioning electronics equipment is mounted on thebottom platform. Thruster number and con0guration, andsolar array size, depend on the launch scenario and thrustertype.

4.4. Chemical propulsion module (CPM)

The main function of the CPM is to host the propul-sion system employed for attitude control during the cruiseand for Mercury orbit insertion. The attitude control func-tions are served by a redundant set of eight 20 N thrusters,while the planetary capture and Mercury orbit acquisitionmanoeuvres are performed by a 400 N engine. A pressurisedbipropellant system is used (Monomethyl hydrazine fuel andNitrogen tetroxide oxidiser in 1.64 mixture ratio) providinga speci0c impulse of 315 s. The CPM also serves as a struc-tural interface between the SEPM and the scienti0c mod-ules, and as a support for additional equipment (split launch1). The general features of the CPM are common to all mis-sion options. The CPM is a short truncated cone, cappingthe SEPM thrust cone. The propellant tanks are supportedby the main structure, and protrude below it into the SEPM,in order to minimise the length of the structure. The lowerside of the CPM is thermally insulated all over in order toprovide thermal protection after the jettisoning of the SEPM.In the mission, options including the MPO (Ariane and splitlaunch 2), the CPM is mounted laterally, in order to havethe MPO in the same attitude to the Sun in both the cruiseand the operational phase.

4.5. Mercury planetary orbiter (MPO)

The con0guration of the MPO is driven by the thermaldesign, the purpose of which is to reject as much as possible,the very large heat inputs from the Sun and the planet, bymeans of high-eQciency insulation over all the body and alarge radiator. The external shape (Fig. 6) is a 6at prism withslanting sides, tilted by 20◦ to reduce the view factor to theplanet. Three sides (±X;+Y ) are partially covered with so-lar cells, mounted on an Al substrate with, on average, a 30%cell 0lling factor, the remaining 70% being covered with Op-tical Solar Re6ectors. Power is generated at any permittedsolar incidence, due to the 20◦ slant. The maximum powerdemand is 420 W (perihelion). The −Y side is covered bya 1:5 m2 radiator, the size of which is driven by the internalpower dissipation (limited to about 200 W). The radiatoris protected from the Sun by the nadir pointing attitude(Fig. 7); a manoeuvre rotating the spacecraft around theZ-axis by 180◦ is made every half Mercury year. It is alsoprotected from the planet by a deployable shield, largeenough to block the IR radiation for any permitted viewfactor to the planet (about 3:4 m2 for 400-km minimumaltitude). The IR shield is stowed at launch against theradiator, because of launcher fairing constraints.

Three star sensors view through the radiator side andare mounted to a stable optical bench with the payloadcameras. The optical payload instruments are recessed intothe spacecraft and view the planet through dichroic mir-rors. Multi-layer insulation (MLI) sheets are used for thenon-optical instruments. High-temperature MLI insulates


Table 7Soyuz–Fregat split-launch mass budget summary (2009 launch)

Launch mass budget (kg) Baikonur Baikonur KourouSoyuz–Fregat split-launch con0guration January 09 July 09 January 09

Direct injection Lunar swing-by Direct injection

Launch 1 — MMO+MSE compositeMMO 165.3 165.3 165.3MSE 219.5 219.5 219.5CPM dry 89.4 89.4 89.4Subtotal 1 (dry mass at Mercury) 474.2 474.2 474.2CPM propellant 130.9 131.0 130.9Subtotal 2 (mass after jettison) 605.1 605.2 605.1SEPM dry 366.5 366.5 366.5Subtotal 3 (mass before jettison) 971.6 971.7 971.6SEPM propellant 263.0 252.0 263.0Launch vehicle adapter 49.0 49.0 49.0Launch mass 1283.6 1272.7 1283.6Soyuz–Fregat limit launch mass 1330.0 1550.0 1510.0System margin (kg) 46.4 277.3 226.4System margin (%) 4 22 18Launch 2 — MPO compositeMPO 357.3 357.3 357.3CPM dry 77.1 77.1 77.1Subtotal 1 (dry mass at Mercury) 434.4 434.4 434.4CPM propellant 158.0 158.0 158.0Subtotal 2 (mass after jettison) 592.3 592.4 592.3SEPM dry 365.5 365.5 365.5Subtotal 3 (mass before jettison) 957.8 957.9 957.8SEPM propellant 259.3 248.5 259.3Launch vehicle adapter 49.0 49.0 49.0Launch mass 1266.1 1255.4 1266.1Soyuz–Fregat limit launch mass 1330.0 1550.0 1510.0System margin (kg) 63.9 294.6 243.9System margin (%) 5 24 19

Fig. 6. MPO con0guration.

the interior everywhere, but, at the radiator and the instru-ment apertures.

The major externally mounted element is a deployable,two-axis articulated, 1:5 m-diameter High Gain Antenna(HGA), mounted on a short boom on the zenith side. Asuitable latch mechanism restrains the antenna at launch.The scienti0c data are transmitted to Earth at Ka band, in

suppressed carrier mode, at a rate variable with the distanceto the Earth, for about 26% of the time (assuming oneground station and after subtracting the time lost for planetoccultations). Link availability is calculated assuming theusual 3 dB link margins. The overall data return in 1 yr is1550 Gbit. An ultra-high frequency (UHF) dipole array,mounted to the nadir side, is used for communications withthe MSE.

The MPO launch mass budget is 357 kg.

4.6. Mercury magnetospheric orbiter (MMO)

The con0guration of the MMO (Fig. 8) is driven by thethermal design both in shape and size. MMO is a spinningsatellite with the spin axis perpendicular to Mercury’s equa-tor. The external shape is a 6at cylinder. The side wall is ex-posed to solar illumination and, hence, also carries the solarcells, whereas the top and bottom sides are used as radia-tors. The main structural element is a truncated cone man-ufactured from carbon-0bre-reinforced plastic. A cold gastank is positioned inside of the cone with the centre closeto the centre of mass. Two equipment mounting panels areattached with 6anges to the inner cone. All the scienti0cinstruments are mounted on one of the outer half rings or


Fig. 7. MPO 6ight attitude.

Fig. 8. MMO con0guration.

directly to the equipment panels. Due to the large radiativearea needed to maintain an acceptable temperature at themaximum-temperature excursion, active temperature con-trol is needed, and is implemented by two sets of louverssituated on the ±Z faces. Bimetallic spring actuators controlthe opening and closing of the louvers. When the louversare fully open the equipment radiator panels are exposed tospace, when closed, the internal temperature is maintained.A mixture of solar cells, second-surface mirrors and thermalblankets covers the outer ring. The maximum power deliv-ered by the solar panels is 186 W.

For communication with Earth, a de-spun high-gain oJsetantenna operating at X-band and two X-band medium-gain

Fig. 9. MSE assembly (single launch version).

antennas (MGA), orientable around one axis, are employed.The HGA is used as the main link antenna to the Earthduring the Mercury mission, when the MMO operates ina spin-stabilised mode. The oJset con0guration provides±12◦ orientation capability to cover the variations in theelevation of the Earth. The HGA is positioned about thespacecraft spin axis to enable the whole assembly to be de-spun when the satellite is operating in spin-stabilised mode.The antenna is stowed during the launch phase and de-ployed after launch. The data return performance of the HGAin 1 yr is 160 Gbit (X band, residual carrier mode, 20 WRF transmitter power, 3 dB link margins assumed). Duringthe cruise, when the satellite operates in a 3-axis stabilisedmode, the medium-gain antennas are used. Communicationwith the Mercury lander is achieved by a microstrip UHFpatch antenna situated on the −Z face of the satellite in-side of the interface ring. Deployable Cluster-type booms(Cluster-II Project Team, 2000) and the associated instru-ments are stowed for the launch and cruise phase of themission and are located at the −Z face on specially hard-ened supports and hold-downs. They are arranged on eitherside of the interface cone and positioned so that they candeploy radially outwards below the lower edge of the outerring.

The MMO launch mass budget is 165 kg.

4.7. Mercury surface element (MSE)

The hard-lander variant of the MSE has two versions:

• equipped with two motors (insertion and descent), in thesingle-launch scenario (Fig. 9),

• equipped with descent motor only, for application in thesplit-launch scenario, where the CPM performs the inser-tion manoeuvre.

With the exception of the insertion motor and its interfacestructure, the basic design of the MSE is identical. The MSE


Fig. 10. Deployed MSE Aftbody and Forebody.

has an Aftbody and a Forebody as its central core. The Aft-body contains all the geochemical experimentation, cameras,power supplies and communication. The Forebody holds thegeophysical experiments and is connected to the Aftbodyvia an umbilical cord.

One or two solid propellant motors are attached to theAftbody structure. The 0rst is used to insert the MSE intoa ballistic-entry trajectory (single-launch scenario); the sec-ond controls the descent velocity of the MSE. At a givenheight, lateral thrust motors rotate the MSE so that it is ori-entated vertically, and then it is allowed to fall freely. Theaxial crushing of the main motor casing controls the Aft-body deceleration load, while the Forebody is free to impactand penetrate the surface of Mercury under its own inertia(Fig. 10).

A 2-Hz spin is imparted to the MSE prior to the separa-tion as an aid to trajectory stabilisation. This is accomplishedby either the CPM (split-launch scenario) or a dedicatedspin-eject device (Ariane-5 scenario). A small lightweightradar altimeter monitors the descent rate and altitude. Ve-locity and attitude rates are monitored by solid state rategyros. The main central processing unit is supplied with suf-0cient information to detect the height and rate of changeof velocity. Discrete thrusters, radially orientated around thebase of the motor case, rotate the MSE through ∼80◦ from ahorizontal to vertical position. A separate module providesthe power supply for activating and controlling individualthrusters.

The wet mass is 220 kg (234 kg with insertion motor).The mass prior to impact is 63:8 kg, which from a freefall drop height of 1:5 km equates to an impact energy of351 kJ at an impact speed of 105 m=s. Upon impact thespent motor casing=nozzle is utilised as the main energyabsorption device. It is constructed as a 0lament-wound

carbon composite, whose thickness (2 mm) and 0bre ori-entation are optimised to withstand the internal pressurisa-tion during motor burn and yet deform readily under theaxial-impact loading. The collapse force is governed bythe 2500 m=s2 deceleration loading imposed on the Aft-body.

The Aftbody is constructed from aluminium honey-comb skinned with carbon 0bre, making the structurelightweight with high stiJness. All the scienti0c equip-ment, control and communication packages are attachedto the main platform. The micro-rover restraints are in-corporated into the Aftbody, and their release is activatedsimultaneously with the opening of the access panel lo-cated in the side wall and the deployment of the com-munications antenna. The micro-rover is free to exit theAftbody but remains connected to the MSE by an um-bilical cable. The camera is positioned on top of the Aft-body, inside the well of the Forebody penetrator supportstructure. The Forebody is retained within the Aftbodycentral support hub, enveloped by the main rocket ini-tiator. Due to the high thermal loads, the penetrator isprotected by an additional layer of intumescent mate-rial to limit its temperature below 100◦C. The penetrator,containing the geophysics package, is manufactured fromtitanium. An armoured spoolable umbilical cord connectsthe Forebody to the Aftbody data handling and storageequipment. As the Aftbody decelerates on impact, theinertia forces acting on the Forebody are more than suQ-cient to fail the shear-limited restraints (set at 500 m=s2).Once separated, the Forebody is free to pass through thedeforming motor casing=nozzle, impact and penetrate thesurface.

The penetration depth of the Forebody into the regolith,from a free fall height of 1:5 km, is up to 7 m.

Given the stringent mass constraints, it was decidedto supply the MSE with a primary high-energy-densitybattery only. At 85◦ latitude, the probability that a solargenerator is not illuminated is high (40% of the terrain isestimated to be in shadow). A lightweight option wouldconsist of a small solar generator (0:5 kg) directly feedingthe load or part of it, without charging or discharging a sec-ondary battery. The energy available from the 2 kg primaryLithium-ion battery (a development item, including impacthardening) is 472 Wh. Communications, data handling andauxiliary power consume 127 Wh leaving 345 Wh to thepayload (259 Wh secondary power at 75% converter eQ-ciency). The scienti0c data are stored in a mass memoryand transmitted to the companion Orbiter (MMO or MPOdepending on the scenario) on each overhead pass. In thecase of the MSE-to-MMO link, a mean usable data rate of8:7 kb=s is provided by the telemetry system. The amountof downlink capability the payload can use accumulates to68 Mbit for 7 d of operations (18 contact periods of 480 s).About twice this amount (138 Mbit) is collected in the caseof the MSE-to-MPO link (more contact periods due to thelower near-circular orbit).


Table 8Possible launch combinationsa

Payload Launch vehicle and launch site Launch date Arrival at Mercury

Light composite 1 Soyuz–Fregat, Baikonur May 2007 September 2009Light composite 2 Soyuz–Fregat, Baikonur or Kourou January 2009 August 2011Heavy composite 1 Soyuz–Fregat, Baikonur + lunar swing-by January 2008 August 2011Heavy composite 2 Soyuz–Fregat, Baikonur + lunar swing-by August 2009 October 2012Heavy composite Soyuz–Fregat, Baikonur + lunar swing-by January 2008 August 2011Light composite Soyuz–Fregat, Baikonur January 2009 August 2011Heavy composite Soyuz–Fregat, Baikonur + lunar swing-by August 2009 October 2012Light composite Soyuz–Fregat, Baikonur August 2010 October 2012Single composite Ariane-5, Kourou January 2009 August 2011Single composite Ariane-5, Kourou August 2010 October 2012

aBold characters identify the current baseline scenarios. “Heavy composite” is the current split-launch con0guration, “Light composite” applies if eitherthe mass is reduced or the Soyuz launch capability is improved. “Single composite” is the con0guration applicable to the Ariane-5 launch.

5. Conclusions

The study demonstrated the ample 6exibility of launchoptions compatible with the mission requirements and thecapability of the mission to be adapted to varying pro-grammatic constraints. Results have been obtained for threealternative system con0guration options (single-launch,and split-launch with or without lunar swing-by). Thesplit-launch mission with lunar swing-by is the currentbaseline, providing the required ∼ 20% mass margin at thecost of a one-year longer transit time to Mercury. However,other possibilities for increasing the mass margins havebeen identi0ed and will be pursued in order to reduce thecruise time to 2:5 yr. They include a more realistic eval-uation of the speci0c impulse of ion thrusters, and furthermass reduction measures in the composite launch con0g-uration (optimisation of the SEPM and CPM, merging ofthe CPM with the MSE propulsion system in Launch 1).Moreover, the Soyuz–Fregat performance may be improvedwith respect to the present manifest by 2007, and even moreso for a 2009 launch. In particular, if a launch facility inKourou is developed for this launcher, the launch capabilityfor escape at 2:3 km=s increases to 1510 kg, providing therequired ∼ 20% margin (Table 7). Plans for implementa-tion of a Soyuz launch pad in Kourou already exist, subjectto a commercial decision.

Assuming launch in 2009, if suQcient mass margin canbe recovered in both composites, then the mission param-eters in column 3 of Table 5 apply. If not, the compositesmust use the lunar swing-by option (column 5 of Table 5).If only one of the two composites can be built within themass margins, the heavier composite can be launched in thelunar swing-by window, with the lighter composite follow-ing one year later without lunar swing-by. In that case, bothwill reach Mercury approximately at the same time. Thisis of interest for science operations, but it requires two dif-ferent sets of mission operations to be developed (with andwithout lunar swing-by). Table 8 shows a summary of pos-sible combinations of split-launch scenarios, compared withtwo single-launch scenarios.

In parallel with the system study, that addressed asolid-propulsion-based MSE design, a study was carriedout by ESTEC on a soft lander equipped with bipropel-lant propulsion (Novara, 2000). The results of both studiesdemonstrate the 6exibility of options available for the de-sign of the MSE. The hard-lander described in Section 4.7is retained as the baseline for the estimation of the systembudgets. However, numerous options have been generatedfrom the parallel study, which are worth considering forincorporation at a later stage of the design evolution. Inparticular, using liquid propulsion has the potential to re-duce the overall mass at the system level. This option,coupled with more capable guidance, navigation and con-trol, would allow the adoption of a lower landing speedand of an airbag landing system. These are more suitablefor the operation of geochemical payload in a chemicallyuncontaminated environment. On the other hand, extendedlife time on the surface may well be too ambitious withinthe overall mission context, especially considering the risksinduced by the signi0cant portion (40%) of the landing areapotentially without solar illumination and by the long-termoperation of mechanisms in a harsh thermal environment.

An important task of these system studies was to identifythe new technology developments required for a successfulimplementation of the mission. A core Mercury technologydevelopment plan covering essential technologies for ther-mal control, solar arrays, antennas and radiation protectionhas been laid out. Additional design-speci0c activities willaddress the MSE. Finally, requirements for minor adaptationof existing technologies to the Mercury environment havebeen identi0ed. The technology plan will be implemented inthe next three years, in parallel with the further developmentof the mission and system design.

Acknowledgements

The Mercury Cornerstone study was a collaborativeendeavour of the engineering teams at Alenia Spazio,Astrium-Germany=Dornier and Hunting Engineering Ltd.,


the engineering team at ESTEC, and the Scienti0c Advi-sory Group (SAG), whose collective work is gratefullyacknowledged. Special thanks are due to Yves Langevinof IAS Orsay, member of the SAG, whose gravity-assisttrajectories provided the breakthrough to a realistic andfeasible system design.

References

Cluster-II Project Team, 2000. The Cluster-II Mission. ESA Bulletin 102,47–53.

Grard, R., Novara, M., Scoon, G.E.N., 2000. BepiColombo—amultidisciplinary mission to a hot planet. ESA Bulletin 103, 11–19.

Iafolla, V., Milyukov, V., Morbidini, A., Nozzoli, S., 2000. Mercuryplanetary orbiter radio science experiments. IFSI-CNR Report.

Iess, L., 2000. Radio Science Experiments For The ESA MercuryCornerstone. UniversitTa La Sapienza, Rome.

Langevin, Y., 1999. Chemical and solar electric propulsion options fora Mercury cornerstone mission, IAF-99-A.2.04. 50th Congress of theInternational Astronautical Federation, Amsterdam, The Netherlands,4–8 October 1999.

Milani, A., Rossi, A., Vokrouhlicky, D., Villani, D., 2000. BepiColomboMission: Simulation Of The Full Cycle Data Processing For The RadioScience Experiments. UniversitTa di Pisa, June 30, 2000.

Novara, M. (Ed.), 2000. Mercury Surface Element Assessment StudyReport, ESA CDF-04 (A), March 2000.

Racca, G.D., Marini, A.E., 2000. Smart-1: preparing the next-generationmission to Mercury. IAA-00-IAA.11.2.06. 51st Congress of theInternational Astronautical Federation, Rio de Janeiro, Brazil, 2–6October 2000.

Scoon, G.E.N., Anselmi, A., HoJmann, H.-J., 1999. The Mercurycornerstone mission, IAF-99-Q.2.07. 50th Congress of the InternationalAstronautical Federation, Amsterdam, The Netherlands, 4–8 October1999.

Documents

BepiColombo, ESA's Mercury Cornerstone mission