- Email: [email protected]
The BepiColombo ESA cornerstone mission to Mercury
Pergamon www.elsevier.com/locate/actaastro
THE BEPICOLOMBO
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Acra Astronautica Vol. 51, No. 1-9, pp. 387-39.5, 2002 0 2002 Elsevier Science Ltd. All rights reserved Printed in Great Britain SOO94-5765(02)00065-6 0094-5765/02 $ - seefront matter
ESA CORNERSTONE MISSION TO MERCURY
Mauro Novara Scientific Projects Department, ESPJESTEC, Noordwijk, The Netherlands e-mail: [email protected]
The ESA Science Programme Committee (WC) in September 1999 named the mission in honour of Giuseppe (Bepi) Colombo (1920-1984). The Italian scientist explained Mercury’s peculiar rotation - it rotates three times for every two orbits around the Sun - and suggested to NASA that an appropriate orbit for Mariner 10 would allow several Mercury flybys in 1974-75.
ABSTRACT The ESA BepiColombo Mercury Cornerstone mission will be launched in 2009. The current baseline foresees 3 scientific elements (2 orbiters and a lander), to be launched by two SoyuzRegat vehicles from Baikonur, and chemical and solar electrical propulsion modules. The scientific elements are a Mercury Planetary Orbiter (MPO), a Mercury Magnetospheric Orbiter (MMO) and a Mercury Surface Element (MSE). The MM0 segment of the mission will be. contributed by ISAS of Japan. An alternative launch configuration foresees a single launch on an Ariane 5 from Kourou. A Definition Study is being carried out competitively by two industrial teams, in preparation for an Announcement of Opportunity for the payload, to be issued during 2002, and a Phase B to be initiated in mid 2003. A dedicated technology development programme is also on-going, to cover key technologies such as high-insolation/high-temperature equipment, miniaturised avionics and sensors, and landing systems. 0 2002 Elsevier Science Ltd. All rights reserved.
BepiColombo competed in 2000 with the GAIA mission proposal for the fifth ESA Cornerstone (CS) mission slot. In October 2000, the SPC approved a package of missions for 2008-2013: BepiColombo was selected as CS-5 (2009) and GAIA as CS-6 (2012). At the same time, the Institute of Space and Astronautical Science (ISAS) of Japan expressed the intention of participating into the BepiColombo mission in cooperation with ESA. SCIENTIFIC OBJECTIVES The Mariner 10 US probe remains the only visitor to Mercury, so there are many questions to be answered, including: What will be found on the unseen hemisphere? How did the planet evolve geologically? Why is Mercury’s density so high? What is its internal structure and is there a liquid outer core?
INTBODUCTION
What is the origin of Mercury’s magnetic field?
As the nearest planet to the Sun, Mercury has an important role in learning how planets form. Mercury, Venus, Earth and Mars make up the family of terrestrial planets, each carrying information that is essential for tracing the history of the whole group. Knowledge about their origin and evolution is a key to understanding how conditions supporting life arose in the Solar System, and possibly elsewhere. As long as Earth-like planets orbiting other stars remain inaccessible to astronomers, the Solar System is the only laboratory where we can test models applicable to other planetary systems. The exploration of Mercury is therefore of fundamental importance for answering questions of astrophysical and philosophical significance, such as “Are Earth-like planets common in the Galaxy?“.
What is the chemical composition of the surface? Is there any water ice in the polar regions? Which volatile materials atmosphere (exosphere)?
form
the
vestigial
How does the planet’s magnetic field interact with the solar wind? BepiColombo’s other objectives go beyond the exploration of the planet and its environment, to take advantage of Mercury’s proximity to the Sun:
A Mercury mission was proposed in May 1993 for the ESA M3 mission selection (eventually won by Planck). AIthough the assessment study showed it to bee too costly for a medium-class mission, it was viewed so positively by ESA that, when the Horizon 2000 science programme was extended in 1994 with Horizon 2000 Plus, the three new Cornerstones included a Mercury orbiter. Since that time, it has been the object of several investigations by the Agency, the European scientific community, and industrial contractors.
.
Fundamental science: is Einstein’s theory of gravity correct?
.
Impact threat: what asteroids lurk on the sunward side of the Earth? SYSTEM DESIGN
A one-year System and Technology Study completed in April 1999 revealed that the best way to fulfil the scientific goals is to fly two spacecraft in orbit around Mercury (a Planetary Orbiter and a Magnetospheric Orbiter), as well as a Surface Element, necessary for the detailed geochemical and geophysical exploration of the
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surface. The BepiColombo system design is therefore based on 3 scientific elements and 2 propulsion modules. The satellite design concept is thoroughly modular, so as to meet the environmental and mission requirements, minimise the complexity of the scientific elements, and take advantage of the most recent technology developments. The scientific elements are the Mercury Planetary Orbiter (MPO), the Mercury Magnetospheric Orbiter (MMO), and the Mercury Surface Element (MSE). The propulsion modules, the Solar Electrical propulsion Module (SBPM) and the Chemical Propulsion Module (CPM), are identical for the two composites of the split-launch baseline (described hereunder). in order to minim& the procurement cost.‘*z 3
Instrument
Range
[kg]
Power [WI
IhSS
Narrow angle camera CNAC)
350-loo0 nm
I
13
Wide angle cameta (WAC)
400-loo0 nm
3
5
IR mapping spectro~ (IMS)
800-2800 nm
6
10
LTVspectrometer (UVS)
70-330 nm
3
3
X-ray spectrometer (MXS)
0.5-10 keV
4
2
Gamma-ray spectrometer (MGS)
0.1-8 MeV
13
2
Neutron spectrometer (MNS)
O-5 MeV
5 1
Results are not yet available for publication, due to the competitive nature of the work. Hence data repotted in this paper reflect the baseline at the start of the Definition Study. In particular, the split-launch system configuration reflects the baseline presented to SPC for approyal in October 2000. However, the on-going studies are considering alternative concepts, such as a single spacecraft composite launch on an Ariane 5. ISAS of Japan will provide the MM0 element, which is being defined under ISAS responsibility. MODEL PAYLOAD Table 1 and Table 2 show the MPO and MSE Model Payload currently used by ESA and industry for the purpose of system definition. Not all candidate instruments listed in those tables can be accommodated at once, the payload complement of the MPO and MSE being limited to 70 kg and 7 kg, respectively.
3 I
Radioscience (RSE): Transponder (COM)
9
Accelerometer :ACC)
6 Table 2: MSE Model Payl&
Laser altimeter (TOP)
An Announcement of Opportunity (AO) for the selection of the actual MPO and MSE scientific payloads will be issued in early 2002. An A0 for the MM0 instruments will be issued by ISAS somewhat later.
Vear-Earth Objects &scope (NET) Hagnetometer :MAG)
i256,i4096 nT
3
Veutral particle malyser (NPA)
O-100 keV
6
8
hst mass pectrometer @MS)
1-loc0ou
2
4
R radiometer (IRR)
2ooo-6oooo nm
5
6
3 MISSION DESIGN The mission requires the delivery of a mass of about 1,100 kg in Mercury orbit. The study has demonstrated the existence of a variety of options compatible with the mission requirements and programmatic constraints. The scenarios that have been analysed in detail are: .
A single launch (Ariane 5) with chemical propulsion only, or with a mixed chemical and solar electric propulsion (SEP) system.
.
Split launch of MPO and MMO-MSE (SoyuzlFngat) with chemical and solar electric propulsions (with or without lunar swingby).
Table 1: MPO Model Payloaa! Two industrial Definition Studies are being conducted from May 2001 to late 2002 to prepare for the one-year Phase B to begin by mid-2003 and Phase UD in 2004.
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Both chemical and mixed propulsion options use gravity assists at Venus; the launch windows are spaced at intervals of 1.6 years. A split launch of MPO and MMOMSE can take place either in the same window or in successive windows. Launches in May 2007 and January 2009 (or about half a year later, with a lunar swingby) have been considered. A significant drawback of the purely chemical option is the long cruise time (6 years or more) compared to that of the SEP option (2.1 to 2.5 years, without lunar swingby, 3.5 years with lunar swingby). The combination of SEP with gravity assists provides a set of mission oppottunities with short cruise times and outstandlng flexibility. The mass margin is largest with At-lane 5 and SEP, and adequate in all other options. Considerations of cost and schedule resulted in the selection of a split launch with SEP in 2009 as the BepiColombo baseline. This is currently being reviewed in the industrial definition studies, taking into account new requirements and constraints dictated by the need to integrate an ISAS-provided MMO, and a constantly maturing design of the MSE segment of the mission. In order to increase the mass margin, both spacecraft are launched into a high-apogee Earth orbit, and achieve escape velocity by a lunar swingby. With this strategy, the launch capability of the Soyux/Fregat launcher provides sufficient mass margin at the cost of a one-year longer transit time to Mercury. The Soyuz/Fregat performance used for the study is the one expected for the Mars Express launch in 2003, corresponding to a 6070kg increase with respect to the performance announced by Starsem for end 2001. Such increase corresponds to the adoption of new stage drop zones in Kazakhstan. In addition, the Soyuz ST+ version of the Soyuz launcher (with a new 3rd stage engine) is currently under evaluation by Starsem and may be available already as early as 2002, providing an increase of 150-200 kg on top of the end-2001 launch performance, and thus further launch mass margin for BepiColombo. Table 3 summarizes the main parameters of the splitlaunch option with lunar swingby and SEP. Figure 1 shows the trajectory for an August 2009 launch (Mercury arrival in October 2012).
Table 3: Baseline mission paramerers with .Soyuz/Fregar.
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The DSl mission of NASA has successfully tested SEP in space on a modest scale. SMART-l, to be launched in late 2OOUearly 2003, will validate all system aspects of a mission associating SEP with gravity assists. The specific impulse of ion thrusters used for this study (3,400 s) is a realistic average between the beginning-of-life and the actual BepiColombo end-of-life (~7,000 h) values. The thrust times are compatible with the demonstrated lifetime (10,000 h), and there are recovery strategies for any single thruster failure and some double thruster failures.
I.
1.. -t1
.
I....
-as
I....)
aa
.
(u
.
.
.
II La
Figure 1: Ecliptic projection of Eonh to Mercury trajectory for the August 2005’ opporfuniry using lunar swingby. A first thrust arc increases the veloci@ relative to the Earth from 1.4 to 2.8 km/J. Afer the Earth swingby, two Venus gravity assists (with a 180” tratufer from Venus to Venus) and fwo Mercury gravity assists take place beforefinal encounter with Mercury.
Capture by Mercury and manoeuvring in Mercury orbit would expose the solar array to high IR fluxes from the planet on top of direct solar radiation. This would require a complex approach strategy (seveml months long) to keep the array temperature within a tolerable envelope. Therefore, the entire SEPM is jettisoned when close to Mercury, and the capture and ins&on manoeuvres are performed by much more efficient chemical propulsion (the 4kN bi-propellant engine of the CPM). This chemical propulsion capability also provides recovery options if there is a total failure of SEP at a late stage of the cruise phase. The MM0 orbit (400 x 11,800 km, polar) is reached first in both launches. In the MPO launch, another bum of the 4kN engine lowers the apocentre to 1,500 km, as required by the MPO; the CPM is jettisoned after both acquisition manoeuvres have been completed. In the MMO-MSE launch, the MM0 is released first on its orbit; the MSE, which is dormant throughout the cruise and approach phases, is then delivered to its destination by the CPM from the MM0 orbit. The MSE is targeted to a landing site at z&t” latitude, where its design life is 1 week.
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The SEP arrival conditions are constrained such that both the MPO and MM0 operational orbits have their line of apsides on Mercury equator, and their periherm on the anti-solar side at perihelion, which creates a more benign environment for the thermal control (Figure 2). Moreover, the MPO and MM0 orbits are resonant (4 MM0 revolutions per MPO revolution), so that there for backup inter-orbiter exists an opportunity communications should one of the two high-gain antenna systems fall to perform according to specifications. The design life of the MM0 and MPO is 1 Earth year (4 Mercury years, or 2 Mercury days).
(6.5 kW for the SPT thrusters and 10.5 kW for the ion thrusters).
Figure 3: MPQ cruise configuration.
SYSTEM
CONFIGURATION
Solit-Launch Svstem Confitruration In the baseline splitlaunch scenario, a first launch of a SoyuzYFregat vehicle delivers a composite made of SEPM-CPM-MPO, while another launch delivers a composite consisting of SBPMCPM-MMO-MSE. Figure 3 and Figure 4 show the cruise configurations for the first (MPO) and second (MMOMSE) launch. Table 4 shows the mass budgets for 2009 Soyuflregat launches from Balkonur. The propulsion elements feature three 200 mN ion thrusters (SEF’M) and one 4 kN bi-propellant thruster (IV = 315 s) plus eight 20 N thrusters for attitude control (CPM). lhe SEPM features a 5.5 kW GaAs cell solar array (with 20% Optical Solar Reflectors), in two wings. Through the interplanetary and early Mercury orbit phase, the command and control tasks are centralised in the respective Orbiters. Sinde-Launch Svstem Confitzuration’ln the single-launch scenario, all modules are delivered by a dedicated launch of Ariane 5 (Figure 5 and Figure 6). In this case, the SEPM can accommodate either gridded-ion or SPT thrusters. The resulting launch mass is either 2,674 kg (ion) or 3,026 kg (SPT), versus a launch capability of Ariane 5 of 3,500 kg into the required escape orbit with 2.&m/s hyperbolic excess velocity, thus mass margins are adequate (30% for ion, 16% for SPT). This flexibility is advantageous from the point of view of the commercial availability of either type of thruster. The main difference is in the required propellant mass and the solar array size
F+we
4: MMO-MSE cruiw conjiguration
MPO
MMOMSE
Magnetospheric Orbiter
0.0
165.3
Surface Element
0.0
44.1
Planetary Orbiter
357.3
0.0
Launch
Chemical Propulsion Module dry
71.1
71.1
Subtotal 1 (dry mass at Mercury)
428.4
280.5
Bi-propellant
155.9
333.6
Subtotal 2 (mass after jettison)
584.3
614.1
SEPM Dry
365.5
365.5
Subtotal 3 (mass before jettison)
949.8
979.6
Cruise Propellant
230.3
237.5
Launch vehicle adapter
49.0
49.0
Launch mass
1229.0
1266.1
Soyuz@regat limit launch mass
1509.6
1509.6
System margin (kg)
280.6
243.5
System margin (96)
23%
19%
Table 4: Split-launch
mass budget summa ry &I.
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Through interplanetary and early Mercury orbit phase, the command and control tasks are centralised in the MPO, providing overall monitoring and control tasks, and managing telecommunications to Earth, while both the MM0 and MSE are dormant. It has further been verified that the split spacecraft designed for the baseline are also compatible with a shared Ariane 5 launch, using the Speltra adapter. This option is also expensive, but may be considered as a backup in case of non-availability of the Soyuz/Fregat launcher in the time frame of relevance to BepiColombo. The launch window is the same as for the baseline singlecomposite Ariane 5 scenario, and the mass margin remains >20 %.
SYSTEM ELEMENTS Solar Electric Prooulsion Module Electrical propulsion is optimal for slow cruise manoeuvres, but unfit for quick insertion needs. Moreover, the large cruise arrays can hardly stand the near-Mercury thermal environment without major new developments. Therefore, a Solar Electric Propulsion Module (SEPM) design based on standard (modified) solar arrays from communication satellite applications (with a maximum operational temperature of 150” C), and thermal control optimised for deep-space conditions, has been adopted. As the spacecraft approaches the sun, the power delivered by the array increases, and so does the array temperature. After the spacecraft has reached a distance from the sun at which the temperature has risen to 150” C, the array is progressively tilted away from the sun. At Mercury (0.32 AU from the sun), the array ends up tilted about 65”. By this strategy, the power delivered is approximately constant (around twice the IAU power) from 0.6 AU onwards. Table 5 summarises the main SEPM characteristics. 1 Number of thrusters
1 3 (either 1 or 2 used) 1
Table 5: SEPM design parameters.
The resulting SEPM vehicle may be regarded as the prototype of a family of upper stages for the Soyufiregat medium launcher, usable for other solar system exploration missions requiring a high post-launch velocity increment. Due to the relatively limited launch capability of Soyuz/Fregat, ion thrusters are the sole candidate technology for electrical propulsion. Two candidate thruster designs with similar characteristics, i.e.. BP Ion Thrusters (RITA-XT) and Electron Bombardment thrusters (T6 IPS) have been considered in the study. Their main characteristics are summarised in Table 6.
I Thrust
1 150-2OOmN 1 Thruster system dry mass (3 units) 1 85 kg level
Beam voltage
13cOv
Specific impulse, BOLEOL
35OOs132OOs
Power demand @ 170 mN
5.0 kW
Table 6: Ion thruster characteristics.
Figure 6: Single-composite
cruise configuration.
The SEPM is a simple rectangular prism. A central thrust cone is the main structural element, and transmits the loads to the launcher interface. The Xenon propellant tanks are housed within it. Two solar array wings, with multiple panels equipped with GaAs cells, provide the
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cruise power. On each wing, a solar array drive mechanism provides one rotational degree of freedom, around the yoke axis, sufficient for the array orientation needs during the cruise. The propulsion drive and power conditioning electronics equipment are mounted on the bottom platform. Chemical Pronulsion Module The main function of the Chemical Propulsion Module (CPM) is to host the bipropellant propulsion system employed for attitude control during the cruise, for Mercury orbit insertion, and for MSE de-orbit and landing. Table 7 gives the main CPM design parameters.
slant. The maximum power demand is 420 W (perihelion). The -Y side is covered by a 1.5 m* radiator, the size of which is driven by the internal power dissipation (limited to about 200 W). The radiator is protected from the Sun by the Nadir pointing attitude; a manoeuvre rotating the spacecraft around the Z-2s by 180” is made every half Mercury year. The radiator is also protected from the planet by a deployable shield. large enough to block the IR radiation for any permitted view factor to the planet (about 3.4 m* for 400 km minimum altitude). The IR shield is stowed at launch against the radiator.
;I
I’1including 18.8 kg Interface Cone Table 7: CPM design parameters. The attitude control functions are served by a redundant set of eight 20 N thrusters, while the planetary capture, Mercury orbit acquisition, de-orbit and landing manoeuvres are performed by a 4,000 N engine. Alternative engines with lower thrust (down to 1,500 N) may also be used, depending on their off-the-shelf availability. The CPM also serves as a structural interface between the SEPM and the scientific modules, and as a carrier for AGCS sensors and for the CLAM-D camera head. The CPM structure consists essentially of a main platform holding together the elements of the propulsion system, and of a short Interface Cone, capping the SEPM thrust cone and providing a mounting interface for the MM0 and MPG. The propellant tanks protrude under the platform into the SEPM, in order to minirnise the total height of the structure. The lower side of the CPM is thermally insulated in order to provide thermal protection after the jettisoning of the SEPM. In the MPG launch, the CPM is mounted laterally, in order to maintain the same attitude with respect to the Sun and planet as in the MPG operational phase. In the MMG-MSE launch, the MM0 is used to remove the Interface Cone (which is jettisoned afterwards) when separating from the CPM-MSE composite spacecraft, shown in Wgure 7. The MSE is supported by brackets welded to the propulsion tanks. Planetarv Orbiter The configuration of the Planetary Orbiter is driven by the thermal design, the purpose of which is to reject as much as possible the very large heat inputs from the Sun and the planet, by means of highefficiency insulation over all the body and a large radiator. The external shape (Figure 8) is a flat prism with slanting sides, tilted by 20” to reduce the view factor to the planet. Three sides (kX, +Y) are partially covered with solar cells, mounted on an Al substrate with, on average, a 30% cell filling factor, the remaining 70% being covered with Optical Solar Reflectors. Power is generated at any permitted solar incidence, due to the 20”
Figure 7: CPM-MSE composite afrer separation from MMO. Three star sensors view through the radiator side, and are mounted to a stable bench-like structure, to which also the payload cameras are rigidly connected. The optical payload instruments are recessed into the spacecmft and view the planet through dichroic mirrors. MLJ sheets are used for the non-optical instruments. High-temperature MLI insulates the interior everywhere but at the radiator and the instrument apertures.
Figure 8: MPO configuration. The major externally mounted element is a deployable, 2axis articulated, 1.5m diameter high-gain antenna (HGA), mounted on a short boom on the Zenith side. A latch mechanism restrains the antenna at launch. The scientific data are transmitted to Earth in Ka-band, in suppressedcarrier mode, at a rate variable with the distance from
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Earth, for about 25% of the time (assuming one ground station and subtracting the time lost for planet occultations); the overall data return in 1 year is 1.550 Gbit. A UHF dipole array, mounted on the Nadir side, is used for communications with the MSE. Table 8 shows the MPO mass budget.
I
Subsystem
1 Mass [kg1 1
Structure & Mechanisms
88.7
Thermal Control
47.1
RCS including Propellant
24.7
Power & Harness
48.6
Solar generator elements
above, the Definition Study is re-evaluating options which minimise the utilisation of MM0 subsystems during the cruise. The need for duplicating some of the key control and communications functions may lead to tbe adoption of a dedicated Service Module (SM) as part of the cruise composite spacecraft. Surface Element The MSE provides Mercury-surface ground-truth measurements for a duration of about one week. MSE landing is targeted at regions (and epochs) in which the mean surface temperature stays between -50’ and +65’ C (i.e. >84’ latitude in either hemisphere). The 0.9m diameter MSE has a mass of 44 kg, of which 7 kg have been allocated to the science payload. Overall configuration and mass budget are given in Figure 10 and Table 9.
5.6
AOCS
23.6
Data Handling
26.0
‘IT&C (X/Ka-band)
33.0
Payload
60.0
Total
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357.3
Table 8: UP0 mass budget. Mannetosoheric Orbiter The MM0 is a small spinning satellite (Figure 9). which will be placed in an eccentric, 400 by 11,800 km polar orbit around Mercury. It will be equipped with a range of fields and particles science instruments to allow the analysis of the magnetospheric physics of Mercury. ISAS is currently developing the MM0 segment of the mission! Figure 10: MSE deployed confisumtion.
Mass[kg1 15.3 Mechanisms
1.2
Thermal Control
3.2
’
11.2 1.4 0.6 1.4 ‘IT&C (UHF Relay)
2.5
Payload Total
7.2
I
44.1
Figure 9: MM0 conjigumtion (courtesy of ISAS).
Table 9: MSE mass budget.
With respect to the previous ESA baseline, the ISASdesigned MM0 presents an increased mass (from 165 to 199 kg). In addition, the use of the MM0 navigation sensors, avionics and communications during the cruise to Mercury (split-launch option) was baselined by ESA during the previous study. However, such an approach becomes somewhat complex in terms of development and verification, if the MM0 is under the responsibility of an external agency such as ISAS. As a consequence of the
After release of the MMO, the MSE is deployed by a CPM bum on a low-perihenn (1Okm) orbit. After a 75minute autonomous descent, a final braking manoeuvre is also carried out by the CPM 4kN thruster, from an altitude of about 10 km above the surface, until zero velocity is reached at an altitude of ~120 m. Control during the tinal manoeuvre is based on gyro/accelerometer information and optical range/rangerate sensing. The MSE then separates from the CPM,
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which crashes at ~100 m distance from the MSE (thus avoiding chemical contamination of the MSE landing site). Airbags are used as impact attenuator, rather than a crushable structure (which would be more sensitive to the nature of the terrain). A maximum 3Om/s touchdown speed is attained after a 12Om free fall: this corresponds to the terminal velocity used by parachute-based landing systems under development for Mars, allowing the same airbag design to be used. A maximum impact deceleration of 200 g (for 20 ms) is experienced under these circumstances; a whole range of rugged&d equipment (payload instruments, power system, avionics) to be used on Mars landers will be available for re-use in the MSE. The complete deployment sequence for the MMO-MSE launch is depicted in Figure 11 and Figure 12.
Figure 1 I: MM0 and MSE deployment sequence: SEPM jettison, separation of MMOfrom CPM-MSE, jettison of Inte+ce Cone from MMO.
discharging a secondary battery, thus allowing extended surface operations. No mechanisms are used in the MSE design, except for the deployment of individual payload components (tethered microrover and mole), therefore increasirTg the system reliability. The craft is fully insulated, to cope with the low-temperature environment in a shadowed area. Should landing occur in sunlight, a jettisonable cover is expelled, and a topside radiator is used to dump waste heat from the MSE. ‘lhe radiator is slightly recessed inside the MSE structute. in order to prevent direct sunlight to shine on the radiating surface, up to at least 20“ incidence. Communications, data handling and power subsystems will consume about 1.4 kWh. leaving 300 Wh of energy from primary power to the payload. The scientific data are stored in a mass memory and transmitted to the companion orbiter at each overhead pass. Either the MPG or the MM0 can be used as relay. In the case of MSE to MM0 link, a mean usable data rate of 8.7 kWs is provided by the UHF telemetry system The payload can use a total of 75 Mb for the 7 days of operations (18 contact periods of 480 s). Indicatively, 68 Mbit may be used for imagery (compressed), to rettieve highresolution and colour images taken by the CLAM-D during descent, stereo pairs of the surface from the CLAM-S, and close-up images from the MIRGCAM on the microrover. The rest of the data volume is shared among the & (e.g. 2 Mbit from 73 hours of operations, using 120 Wh including mole operations), the APXS and MIMOS II (2 Mbit, 6 hours of measurements, 150 Wh including microrover operations), and the MLMAG (3 Mbit, 100 hours, 30 Wh). Nearly twice this total data volume (138 Mb) is provided in the case of MSE to MPG link (more frequent contact periods due to the lower nearcircular orbit). In addition, by appropriately phasing the MSE descent with the MPG orbit, it is possible to ensure visibility of the MPG from the MSE for at least a portion of the descent trajectory, so that telemetry of vital parameters from the MSE at a low data rate can be maintained until landing. TECHNOLQGY DEVELOPMENT
Figure 12: Separation of MSEfrom CPM, MSE landing and airbag separation, MSE on ground with deployed payload (thennal protection cover is jettisoned).
MSE main design choices are largely driven by the severe mass constraints, and by the risk inherent in the need to operate in a largely uncharted terrain, with consequent risk of landing in a shadowed area, or anyway with limited sun elevation. An energy-limited MSE design has been therefore adopted, with a primary battery of high energy density (1.7 kWh available energy). This approach allows operations to be carried out even in a completely shadowed area for about one week; the minimum mission therefore does not rely on availability of solar energy, and has a high probability of success. The probability of landing in shadow at a latitude of around 8.5”is high (40% of the terrain is estimated to be in shadow), but an optional small solar generator (about 0.5 kg) could nevertheless be added and used to feed directly the load, or part of it, without charging or
In parallel to the Definition Study, BepiColombo Technology Development Activities (TDAs) are being carried out. Their overall objective is to achieve demonstration of feasibility of critical technologies by end 2002. in order to allow a selected industrial consortium to start the BepiColombo realisation phase in early 2003 with a robust and consolidated development plan. The key technologies to be developed for BepiColombo include:
.
Thermal control for high temperatures (thermal control materials, heat pipes embedded into structural panels, louvres).
.
Solar generators for high insolation, high temperatures (GaAs cells and cell module assemblies, arrays, and array drive mechanisms).
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Antenna technologies for high temperatures (reflector materials, feed and pointing/de-spin mechanisms).
REFERENCES 1.
“BepiColombo - An Interdisciplinary Cornerstone Mission to the Planet Mercury - System and Technology Study Report”, ESA-SCI(2000)1, April 2oo0, httn://solarsvstem.estec.esa.nl/Mercurv/MercurvAss esshtm.
2.
“BepiColombo - Interdisciplinary Mission to Planet Mercury”, ESA BR-165, September 2000. “Assessment Study Report - Mercury Surface Element”, ESA CDFG(A), March 2000, ftD://ftn.estec.esa.nl/uub/sci-snd/beuicol/CDF. “ISAS Feasibility Study on the BepiColombo/MMO Spacecraft Design”, H. Yamakawa, H. Ogawa, H. Hayakawa, T. Mukai, M. Adachi, IAF-Ol-Q.2.03, 52nd International Astronautical Congress, l-5 October 2001.
Miniaturised avionics (Highly Integrated Control & Data Systems). Landing systems for the MSE (vision-based range/range-rate system for landing navigation, miniaturised sensors, adaptation of airbags). Deployment booms for both orbiters, and Automation & Robotics and scientific instruments for high-temperature surface operations.
3.
Information initiatives among the concerned scientific community are being undertaken (e.g. a Mercury Lander Workshop in December 2000) for the purpose of obtaining a timely feedback on the technology development needs from a scientific payload perspective.
4.
The high temperature-related technologies are of broad applicability to missions to the inner solar system, while the avionics/navigation-related technologies are enabling for a wide range of small planetary missions (miniaturised orbiters and landers). In addition, the SMART-l lunar mission will demonstrate the use of an electric thruster for primary propulsion, combined with planetary swingbys.
ACKNOWLEDGEMENT Data included in this paper are largely baaed on input generated during the BepiColombo System & Technology Study by the industrial team (led by A. Anselmi of Alenia Spazio) and by ESA colleagues, in particular the BepiColombo Project Scientist (R. Grard) for the scientific part. Their contribution is gratefully acknowledged.
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Acra Astronautica Vol. 51, No. 1-9, pp. 387-39.5, 2002 0 2002 Elsevier Science Ltd. All rights reserved Printed in Great Britain SOO94-5765(02)00065-6 0094-5765/02 $ - seefront matter
ESA CORNERSTONE MISSION TO MERCURY
Mauro Novara Scientific Projects Department, ESPJESTEC, Noordwijk, The Netherlands e-mail: [email protected]
The ESA Science Programme Committee (WC) in September 1999 named the mission in honour of Giuseppe (Bepi) Colombo (1920-1984). The Italian scientist explained Mercury’s peculiar rotation - it rotates three times for every two orbits around the Sun - and suggested to NASA that an appropriate orbit for Mariner 10 would allow several Mercury flybys in 1974-75.
ABSTRACT The ESA BepiColombo Mercury Cornerstone mission will be launched in 2009. The current baseline foresees 3 scientific elements (2 orbiters and a lander), to be launched by two SoyuzRegat vehicles from Baikonur, and chemical and solar electrical propulsion modules. The scientific elements are a Mercury Planetary Orbiter (MPO), a Mercury Magnetospheric Orbiter (MMO) and a Mercury Surface Element (MSE). The MM0 segment of the mission will be. contributed by ISAS of Japan. An alternative launch configuration foresees a single launch on an Ariane 5 from Kourou. A Definition Study is being carried out competitively by two industrial teams, in preparation for an Announcement of Opportunity for the payload, to be issued during 2002, and a Phase B to be initiated in mid 2003. A dedicated technology development programme is also on-going, to cover key technologies such as high-insolation/high-temperature equipment, miniaturised avionics and sensors, and landing systems. 0 2002 Elsevier Science Ltd. All rights reserved.
BepiColombo competed in 2000 with the GAIA mission proposal for the fifth ESA Cornerstone (CS) mission slot. In October 2000, the SPC approved a package of missions for 2008-2013: BepiColombo was selected as CS-5 (2009) and GAIA as CS-6 (2012). At the same time, the Institute of Space and Astronautical Science (ISAS) of Japan expressed the intention of participating into the BepiColombo mission in cooperation with ESA. SCIENTIFIC OBJECTIVES The Mariner 10 US probe remains the only visitor to Mercury, so there are many questions to be answered, including: What will be found on the unseen hemisphere? How did the planet evolve geologically? Why is Mercury’s density so high? What is its internal structure and is there a liquid outer core?
INTBODUCTION
What is the origin of Mercury’s magnetic field?
As the nearest planet to the Sun, Mercury has an important role in learning how planets form. Mercury, Venus, Earth and Mars make up the family of terrestrial planets, each carrying information that is essential for tracing the history of the whole group. Knowledge about their origin and evolution is a key to understanding how conditions supporting life arose in the Solar System, and possibly elsewhere. As long as Earth-like planets orbiting other stars remain inaccessible to astronomers, the Solar System is the only laboratory where we can test models applicable to other planetary systems. The exploration of Mercury is therefore of fundamental importance for answering questions of astrophysical and philosophical significance, such as “Are Earth-like planets common in the Galaxy?“.
What is the chemical composition of the surface? Is there any water ice in the polar regions? Which volatile materials atmosphere (exosphere)?
form
the
vestigial
How does the planet’s magnetic field interact with the solar wind? BepiColombo’s other objectives go beyond the exploration of the planet and its environment, to take advantage of Mercury’s proximity to the Sun:
A Mercury mission was proposed in May 1993 for the ESA M3 mission selection (eventually won by Planck). AIthough the assessment study showed it to bee too costly for a medium-class mission, it was viewed so positively by ESA that, when the Horizon 2000 science programme was extended in 1994 with Horizon 2000 Plus, the three new Cornerstones included a Mercury orbiter. Since that time, it has been the object of several investigations by the Agency, the European scientific community, and industrial contractors.
.
Fundamental science: is Einstein’s theory of gravity correct?
.
Impact threat: what asteroids lurk on the sunward side of the Earth? SYSTEM DESIGN
A one-year System and Technology Study completed in April 1999 revealed that the best way to fulfil the scientific goals is to fly two spacecraft in orbit around Mercury (a Planetary Orbiter and a Magnetospheric Orbiter), as well as a Surface Element, necessary for the detailed geochemical and geophysical exploration of the
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surface. The BepiColombo system design is therefore based on 3 scientific elements and 2 propulsion modules. The satellite design concept is thoroughly modular, so as to meet the environmental and mission requirements, minimise the complexity of the scientific elements, and take advantage of the most recent technology developments. The scientific elements are the Mercury Planetary Orbiter (MPO), the Mercury Magnetospheric Orbiter (MMO), and the Mercury Surface Element (MSE). The propulsion modules, the Solar Electrical propulsion Module (SBPM) and the Chemical Propulsion Module (CPM), are identical for the two composites of the split-launch baseline (described hereunder). in order to minim& the procurement cost.‘*z 3
Instrument
Range
[kg]
Power [WI
IhSS
Narrow angle camera CNAC)
350-loo0 nm
I
13
Wide angle cameta (WAC)
400-loo0 nm
3
5
IR mapping spectro~ (IMS)
800-2800 nm
6
10
LTVspectrometer (UVS)
70-330 nm
3
3
X-ray spectrometer (MXS)
0.5-10 keV
4
2
Gamma-ray spectrometer (MGS)
0.1-8 MeV
13
2
Neutron spectrometer (MNS)
O-5 MeV
5 1
Results are not yet available for publication, due to the competitive nature of the work. Hence data repotted in this paper reflect the baseline at the start of the Definition Study. In particular, the split-launch system configuration reflects the baseline presented to SPC for approyal in October 2000. However, the on-going studies are considering alternative concepts, such as a single spacecraft composite launch on an Ariane 5. ISAS of Japan will provide the MM0 element, which is being defined under ISAS responsibility. MODEL PAYLOAD Table 1 and Table 2 show the MPO and MSE Model Payload currently used by ESA and industry for the purpose of system definition. Not all candidate instruments listed in those tables can be accommodated at once, the payload complement of the MPO and MSE being limited to 70 kg and 7 kg, respectively.
3 I
Radioscience (RSE): Transponder (COM)
9
Accelerometer :ACC)
6 Table 2: MSE Model Payl&
Laser altimeter (TOP)
An Announcement of Opportunity (AO) for the selection of the actual MPO and MSE scientific payloads will be issued in early 2002. An A0 for the MM0 instruments will be issued by ISAS somewhat later.
Vear-Earth Objects &scope (NET) Hagnetometer :MAG)
i256,i4096 nT
3
Veutral particle malyser (NPA)
O-100 keV
6
8
hst mass pectrometer @MS)
1-loc0ou
2
4
R radiometer (IRR)
2ooo-6oooo nm
5
6
3 MISSION DESIGN The mission requires the delivery of a mass of about 1,100 kg in Mercury orbit. The study has demonstrated the existence of a variety of options compatible with the mission requirements and programmatic constraints. The scenarios that have been analysed in detail are: .
A single launch (Ariane 5) with chemical propulsion only, or with a mixed chemical and solar electric propulsion (SEP) system.
.
Split launch of MPO and MMO-MSE (SoyuzlFngat) with chemical and solar electric propulsions (with or without lunar swingby).
Table 1: MPO Model Payloaa! Two industrial Definition Studies are being conducted from May 2001 to late 2002 to prepare for the one-year Phase B to begin by mid-2003 and Phase UD in 2004.
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Both chemical and mixed propulsion options use gravity assists at Venus; the launch windows are spaced at intervals of 1.6 years. A split launch of MPO and MMOMSE can take place either in the same window or in successive windows. Launches in May 2007 and January 2009 (or about half a year later, with a lunar swingby) have been considered. A significant drawback of the purely chemical option is the long cruise time (6 years or more) compared to that of the SEP option (2.1 to 2.5 years, without lunar swingby, 3.5 years with lunar swingby). The combination of SEP with gravity assists provides a set of mission oppottunities with short cruise times and outstandlng flexibility. The mass margin is largest with At-lane 5 and SEP, and adequate in all other options. Considerations of cost and schedule resulted in the selection of a split launch with SEP in 2009 as the BepiColombo baseline. This is currently being reviewed in the industrial definition studies, taking into account new requirements and constraints dictated by the need to integrate an ISAS-provided MMO, and a constantly maturing design of the MSE segment of the mission. In order to increase the mass margin, both spacecraft are launched into a high-apogee Earth orbit, and achieve escape velocity by a lunar swingby. With this strategy, the launch capability of the Soyux/Fregat launcher provides sufficient mass margin at the cost of a one-year longer transit time to Mercury. The Soyuz/Fregat performance used for the study is the one expected for the Mars Express launch in 2003, corresponding to a 6070kg increase with respect to the performance announced by Starsem for end 2001. Such increase corresponds to the adoption of new stage drop zones in Kazakhstan. In addition, the Soyuz ST+ version of the Soyuz launcher (with a new 3rd stage engine) is currently under evaluation by Starsem and may be available already as early as 2002, providing an increase of 150-200 kg on top of the end-2001 launch performance, and thus further launch mass margin for BepiColombo. Table 3 summarizes the main parameters of the splitlaunch option with lunar swingby and SEP. Figure 1 shows the trajectory for an August 2009 launch (Mercury arrival in October 2012).
Table 3: Baseline mission paramerers with .Soyuz/Fregar.
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The DSl mission of NASA has successfully tested SEP in space on a modest scale. SMART-l, to be launched in late 2OOUearly 2003, will validate all system aspects of a mission associating SEP with gravity assists. The specific impulse of ion thrusters used for this study (3,400 s) is a realistic average between the beginning-of-life and the actual BepiColombo end-of-life (~7,000 h) values. The thrust times are compatible with the demonstrated lifetime (10,000 h), and there are recovery strategies for any single thruster failure and some double thruster failures.
I.
1.. -t1
.
I....
-as
I....)
aa
.
(u
.
.
.
II La
Figure 1: Ecliptic projection of Eonh to Mercury trajectory for the August 2005’ opporfuniry using lunar swingby. A first thrust arc increases the veloci@ relative to the Earth from 1.4 to 2.8 km/J. Afer the Earth swingby, two Venus gravity assists (with a 180” tratufer from Venus to Venus) and fwo Mercury gravity assists take place beforefinal encounter with Mercury.
Capture by Mercury and manoeuvring in Mercury orbit would expose the solar array to high IR fluxes from the planet on top of direct solar radiation. This would require a complex approach strategy (seveml months long) to keep the array temperature within a tolerable envelope. Therefore, the entire SEPM is jettisoned when close to Mercury, and the capture and ins&on manoeuvres are performed by much more efficient chemical propulsion (the 4kN bi-propellant engine of the CPM). This chemical propulsion capability also provides recovery options if there is a total failure of SEP at a late stage of the cruise phase. The MM0 orbit (400 x 11,800 km, polar) is reached first in both launches. In the MPO launch, another bum of the 4kN engine lowers the apocentre to 1,500 km, as required by the MPO; the CPM is jettisoned after both acquisition manoeuvres have been completed. In the MMO-MSE launch, the MM0 is released first on its orbit; the MSE, which is dormant throughout the cruise and approach phases, is then delivered to its destination by the CPM from the MM0 orbit. The MSE is targeted to a landing site at z&t” latitude, where its design life is 1 week.
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The SEP arrival conditions are constrained such that both the MPO and MM0 operational orbits have their line of apsides on Mercury equator, and their periherm on the anti-solar side at perihelion, which creates a more benign environment for the thermal control (Figure 2). Moreover, the MPO and MM0 orbits are resonant (4 MM0 revolutions per MPO revolution), so that there for backup inter-orbiter exists an opportunity communications should one of the two high-gain antenna systems fall to perform according to specifications. The design life of the MM0 and MPO is 1 Earth year (4 Mercury years, or 2 Mercury days).
(6.5 kW for the SPT thrusters and 10.5 kW for the ion thrusters).
Figure 3: MPQ cruise configuration.
SYSTEM
CONFIGURATION
Solit-Launch Svstem Confitruration In the baseline splitlaunch scenario, a first launch of a SoyuzYFregat vehicle delivers a composite made of SEPM-CPM-MPO, while another launch delivers a composite consisting of SBPMCPM-MMO-MSE. Figure 3 and Figure 4 show the cruise configurations for the first (MPO) and second (MMOMSE) launch. Table 4 shows the mass budgets for 2009 Soyuflregat launches from Balkonur. The propulsion elements feature three 200 mN ion thrusters (SEF’M) and one 4 kN bi-propellant thruster (IV = 315 s) plus eight 20 N thrusters for attitude control (CPM). lhe SEPM features a 5.5 kW GaAs cell solar array (with 20% Optical Solar Reflectors), in two wings. Through the interplanetary and early Mercury orbit phase, the command and control tasks are centralised in the respective Orbiters. Sinde-Launch Svstem Confitzuration’ln the single-launch scenario, all modules are delivered by a dedicated launch of Ariane 5 (Figure 5 and Figure 6). In this case, the SEPM can accommodate either gridded-ion or SPT thrusters. The resulting launch mass is either 2,674 kg (ion) or 3,026 kg (SPT), versus a launch capability of Ariane 5 of 3,500 kg into the required escape orbit with 2.&m/s hyperbolic excess velocity, thus mass margins are adequate (30% for ion, 16% for SPT). This flexibility is advantageous from the point of view of the commercial availability of either type of thruster. The main difference is in the required propellant mass and the solar array size
F+we
4: MMO-MSE cruiw conjiguration
MPO
MMOMSE
Magnetospheric Orbiter
0.0
165.3
Surface Element
0.0
44.1
Planetary Orbiter
357.3
0.0
Launch
Chemical Propulsion Module dry
71.1
71.1
Subtotal 1 (dry mass at Mercury)
428.4
280.5
Bi-propellant
155.9
333.6
Subtotal 2 (mass after jettison)
584.3
614.1
SEPM Dry
365.5
365.5
Subtotal 3 (mass before jettison)
949.8
979.6
Cruise Propellant
230.3
237.5
Launch vehicle adapter
49.0
49.0
Launch mass
1229.0
1266.1
Soyuz@regat limit launch mass
1509.6
1509.6
System margin (kg)
280.6
243.5
System margin (96)
23%
19%
Table 4: Split-launch
mass budget summa ry &I.
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Through interplanetary and early Mercury orbit phase, the command and control tasks are centralised in the MPO, providing overall monitoring and control tasks, and managing telecommunications to Earth, while both the MM0 and MSE are dormant. It has further been verified that the split spacecraft designed for the baseline are also compatible with a shared Ariane 5 launch, using the Speltra adapter. This option is also expensive, but may be considered as a backup in case of non-availability of the Soyuz/Fregat launcher in the time frame of relevance to BepiColombo. The launch window is the same as for the baseline singlecomposite Ariane 5 scenario, and the mass margin remains >20 %.
SYSTEM ELEMENTS Solar Electric Prooulsion Module Electrical propulsion is optimal for slow cruise manoeuvres, but unfit for quick insertion needs. Moreover, the large cruise arrays can hardly stand the near-Mercury thermal environment without major new developments. Therefore, a Solar Electric Propulsion Module (SEPM) design based on standard (modified) solar arrays from communication satellite applications (with a maximum operational temperature of 150” C), and thermal control optimised for deep-space conditions, has been adopted. As the spacecraft approaches the sun, the power delivered by the array increases, and so does the array temperature. After the spacecraft has reached a distance from the sun at which the temperature has risen to 150” C, the array is progressively tilted away from the sun. At Mercury (0.32 AU from the sun), the array ends up tilted about 65”. By this strategy, the power delivered is approximately constant (around twice the IAU power) from 0.6 AU onwards. Table 5 summarises the main SEPM characteristics. 1 Number of thrusters
1 3 (either 1 or 2 used) 1
Table 5: SEPM design parameters.
The resulting SEPM vehicle may be regarded as the prototype of a family of upper stages for the Soyufiregat medium launcher, usable for other solar system exploration missions requiring a high post-launch velocity increment. Due to the relatively limited launch capability of Soyuz/Fregat, ion thrusters are the sole candidate technology for electrical propulsion. Two candidate thruster designs with similar characteristics, i.e.. BP Ion Thrusters (RITA-XT) and Electron Bombardment thrusters (T6 IPS) have been considered in the study. Their main characteristics are summarised in Table 6.
I Thrust
1 150-2OOmN 1 Thruster system dry mass (3 units) 1 85 kg level
Beam voltage
13cOv
Specific impulse, BOLEOL
35OOs132OOs
Power demand @ 170 mN
5.0 kW
Table 6: Ion thruster characteristics.
Figure 6: Single-composite
cruise configuration.
The SEPM is a simple rectangular prism. A central thrust cone is the main structural element, and transmits the loads to the launcher interface. The Xenon propellant tanks are housed within it. Two solar array wings, with multiple panels equipped with GaAs cells, provide the
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cruise power. On each wing, a solar array drive mechanism provides one rotational degree of freedom, around the yoke axis, sufficient for the array orientation needs during the cruise. The propulsion drive and power conditioning electronics equipment are mounted on the bottom platform. Chemical Pronulsion Module The main function of the Chemical Propulsion Module (CPM) is to host the bipropellant propulsion system employed for attitude control during the cruise, for Mercury orbit insertion, and for MSE de-orbit and landing. Table 7 gives the main CPM design parameters.
slant. The maximum power demand is 420 W (perihelion). The -Y side is covered by a 1.5 m* radiator, the size of which is driven by the internal power dissipation (limited to about 200 W). The radiator is protected from the Sun by the Nadir pointing attitude; a manoeuvre rotating the spacecraft around the Z-2s by 180” is made every half Mercury year. The radiator is also protected from the planet by a deployable shield. large enough to block the IR radiation for any permitted view factor to the planet (about 3.4 m* for 400 km minimum altitude). The IR shield is stowed at launch against the radiator.
;I
I’1including 18.8 kg Interface Cone Table 7: CPM design parameters. The attitude control functions are served by a redundant set of eight 20 N thrusters, while the planetary capture, Mercury orbit acquisition, de-orbit and landing manoeuvres are performed by a 4,000 N engine. Alternative engines with lower thrust (down to 1,500 N) may also be used, depending on their off-the-shelf availability. The CPM also serves as a structural interface between the SEPM and the scientific modules, and as a carrier for AGCS sensors and for the CLAM-D camera head. The CPM structure consists essentially of a main platform holding together the elements of the propulsion system, and of a short Interface Cone, capping the SEPM thrust cone and providing a mounting interface for the MM0 and MPG. The propellant tanks protrude under the platform into the SEPM, in order to minirnise the total height of the structure. The lower side of the CPM is thermally insulated in order to provide thermal protection after the jettisoning of the SEPM. In the MPG launch, the CPM is mounted laterally, in order to maintain the same attitude with respect to the Sun and planet as in the MPG operational phase. In the MMG-MSE launch, the MM0 is used to remove the Interface Cone (which is jettisoned afterwards) when separating from the CPM-MSE composite spacecraft, shown in Wgure 7. The MSE is supported by brackets welded to the propulsion tanks. Planetarv Orbiter The configuration of the Planetary Orbiter is driven by the thermal design, the purpose of which is to reject as much as possible the very large heat inputs from the Sun and the planet, by means of highefficiency insulation over all the body and a large radiator. The external shape (Figure 8) is a flat prism with slanting sides, tilted by 20” to reduce the view factor to the planet. Three sides (kX, +Y) are partially covered with solar cells, mounted on an Al substrate with, on average, a 30% cell filling factor, the remaining 70% being covered with Optical Solar Reflectors. Power is generated at any permitted solar incidence, due to the 20”
Figure 7: CPM-MSE composite afrer separation from MMO. Three star sensors view through the radiator side, and are mounted to a stable bench-like structure, to which also the payload cameras are rigidly connected. The optical payload instruments are recessed into the spacecmft and view the planet through dichroic mirrors. MLJ sheets are used for the non-optical instruments. High-temperature MLI insulates the interior everywhere but at the radiator and the instrument apertures.
Figure 8: MPO configuration. The major externally mounted element is a deployable, 2axis articulated, 1.5m diameter high-gain antenna (HGA), mounted on a short boom on the Zenith side. A latch mechanism restrains the antenna at launch. The scientific data are transmitted to Earth in Ka-band, in suppressedcarrier mode, at a rate variable with the distance from
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Earth, for about 25% of the time (assuming one ground station and subtracting the time lost for planet occultations); the overall data return in 1 year is 1.550 Gbit. A UHF dipole array, mounted on the Nadir side, is used for communications with the MSE. Table 8 shows the MPO mass budget.
I
Subsystem
1 Mass [kg1 1
Structure & Mechanisms
88.7
Thermal Control
47.1
RCS including Propellant
24.7
Power & Harness
48.6
Solar generator elements
above, the Definition Study is re-evaluating options which minimise the utilisation of MM0 subsystems during the cruise. The need for duplicating some of the key control and communications functions may lead to tbe adoption of a dedicated Service Module (SM) as part of the cruise composite spacecraft. Surface Element The MSE provides Mercury-surface ground-truth measurements for a duration of about one week. MSE landing is targeted at regions (and epochs) in which the mean surface temperature stays between -50’ and +65’ C (i.e. >84’ latitude in either hemisphere). The 0.9m diameter MSE has a mass of 44 kg, of which 7 kg have been allocated to the science payload. Overall configuration and mass budget are given in Figure 10 and Table 9.
5.6
AOCS
23.6
Data Handling
26.0
‘IT&C (X/Ka-band)
33.0
Payload
60.0
Total
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357.3
Table 8: UP0 mass budget. Mannetosoheric Orbiter The MM0 is a small spinning satellite (Figure 9). which will be placed in an eccentric, 400 by 11,800 km polar orbit around Mercury. It will be equipped with a range of fields and particles science instruments to allow the analysis of the magnetospheric physics of Mercury. ISAS is currently developing the MM0 segment of the mission! Figure 10: MSE deployed confisumtion.
Mass[kg1 15.3 Mechanisms
1.2
Thermal Control
3.2
’
11.2 1.4 0.6 1.4 ‘IT&C (UHF Relay)
2.5
Payload Total
7.2
I
44.1
Figure 9: MM0 conjigumtion (courtesy of ISAS).
Table 9: MSE mass budget.
With respect to the previous ESA baseline, the ISASdesigned MM0 presents an increased mass (from 165 to 199 kg). In addition, the use of the MM0 navigation sensors, avionics and communications during the cruise to Mercury (split-launch option) was baselined by ESA during the previous study. However, such an approach becomes somewhat complex in terms of development and verification, if the MM0 is under the responsibility of an external agency such as ISAS. As a consequence of the
After release of the MMO, the MSE is deployed by a CPM bum on a low-perihenn (1Okm) orbit. After a 75minute autonomous descent, a final braking manoeuvre is also carried out by the CPM 4kN thruster, from an altitude of about 10 km above the surface, until zero velocity is reached at an altitude of ~120 m. Control during the tinal manoeuvre is based on gyro/accelerometer information and optical range/rangerate sensing. The MSE then separates from the CPM,
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which crashes at ~100 m distance from the MSE (thus avoiding chemical contamination of the MSE landing site). Airbags are used as impact attenuator, rather than a crushable structure (which would be more sensitive to the nature of the terrain). A maximum 3Om/s touchdown speed is attained after a 12Om free fall: this corresponds to the terminal velocity used by parachute-based landing systems under development for Mars, allowing the same airbag design to be used. A maximum impact deceleration of 200 g (for 20 ms) is experienced under these circumstances; a whole range of rugged&d equipment (payload instruments, power system, avionics) to be used on Mars landers will be available for re-use in the MSE. The complete deployment sequence for the MMO-MSE launch is depicted in Figure 11 and Figure 12.
Figure 1 I: MM0 and MSE deployment sequence: SEPM jettison, separation of MMOfrom CPM-MSE, jettison of Inte+ce Cone from MMO.
discharging a secondary battery, thus allowing extended surface operations. No mechanisms are used in the MSE design, except for the deployment of individual payload components (tethered microrover and mole), therefore increasirTg the system reliability. The craft is fully insulated, to cope with the low-temperature environment in a shadowed area. Should landing occur in sunlight, a jettisonable cover is expelled, and a topside radiator is used to dump waste heat from the MSE. ‘lhe radiator is slightly recessed inside the MSE structute. in order to prevent direct sunlight to shine on the radiating surface, up to at least 20“ incidence. Communications, data handling and power subsystems will consume about 1.4 kWh. leaving 300 Wh of energy from primary power to the payload. The scientific data are stored in a mass memory and transmitted to the companion orbiter at each overhead pass. Either the MPG or the MM0 can be used as relay. In the case of MSE to MM0 link, a mean usable data rate of 8.7 kWs is provided by the UHF telemetry system The payload can use a total of 75 Mb for the 7 days of operations (18 contact periods of 480 s). Indicatively, 68 Mbit may be used for imagery (compressed), to rettieve highresolution and colour images taken by the CLAM-D during descent, stereo pairs of the surface from the CLAM-S, and close-up images from the MIRGCAM on the microrover. The rest of the data volume is shared among the & (e.g. 2 Mbit from 73 hours of operations, using 120 Wh including mole operations), the APXS and MIMOS II (2 Mbit, 6 hours of measurements, 150 Wh including microrover operations), and the MLMAG (3 Mbit, 100 hours, 30 Wh). Nearly twice this total data volume (138 Mb) is provided in the case of MSE to MPG link (more frequent contact periods due to the lower nearcircular orbit). In addition, by appropriately phasing the MSE descent with the MPG orbit, it is possible to ensure visibility of the MPG from the MSE for at least a portion of the descent trajectory, so that telemetry of vital parameters from the MSE at a low data rate can be maintained until landing. TECHNOLQGY DEVELOPMENT
Figure 12: Separation of MSEfrom CPM, MSE landing and airbag separation, MSE on ground with deployed payload (thennal protection cover is jettisoned).
MSE main design choices are largely driven by the severe mass constraints, and by the risk inherent in the need to operate in a largely uncharted terrain, with consequent risk of landing in a shadowed area, or anyway with limited sun elevation. An energy-limited MSE design has been therefore adopted, with a primary battery of high energy density (1.7 kWh available energy). This approach allows operations to be carried out even in a completely shadowed area for about one week; the minimum mission therefore does not rely on availability of solar energy, and has a high probability of success. The probability of landing in shadow at a latitude of around 8.5”is high (40% of the terrain is estimated to be in shadow), but an optional small solar generator (about 0.5 kg) could nevertheless be added and used to feed directly the load, or part of it, without charging or
In parallel to the Definition Study, BepiColombo Technology Development Activities (TDAs) are being carried out. Their overall objective is to achieve demonstration of feasibility of critical technologies by end 2002. in order to allow a selected industrial consortium to start the BepiColombo realisation phase in early 2003 with a robust and consolidated development plan. The key technologies to be developed for BepiColombo include:
.
Thermal control for high temperatures (thermal control materials, heat pipes embedded into structural panels, louvres).
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Solar generators for high insolation, high temperatures (GaAs cells and cell module assemblies, arrays, and array drive mechanisms).
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Antenna technologies for high temperatures (reflector materials, feed and pointing/de-spin mechanisms).
REFERENCES 1.
“BepiColombo - An Interdisciplinary Cornerstone Mission to the Planet Mercury - System and Technology Study Report”, ESA-SCI(2000)1, April 2oo0, httn://solarsvstem.estec.esa.nl/Mercurv/MercurvAss esshtm.
2.
“BepiColombo - Interdisciplinary Mission to Planet Mercury”, ESA BR-165, September 2000. “Assessment Study Report - Mercury Surface Element”, ESA CDFG(A), March 2000, ftD://ftn.estec.esa.nl/uub/sci-snd/beuicol/CDF. “ISAS Feasibility Study on the BepiColombo/MMO Spacecraft Design”, H. Yamakawa, H. Ogawa, H. Hayakawa, T. Mukai, M. Adachi, IAF-Ol-Q.2.03, 52nd International Astronautical Congress, l-5 October 2001.
Miniaturised avionics (Highly Integrated Control & Data Systems). Landing systems for the MSE (vision-based range/range-rate system for landing navigation, miniaturised sensors, adaptation of airbags). Deployment booms for both orbiters, and Automation & Robotics and scientific instruments for high-temperature surface operations.
3.
Information initiatives among the concerned scientific community are being undertaken (e.g. a Mercury Lander Workshop in December 2000) for the purpose of obtaining a timely feedback on the technology development needs from a scientific payload perspective.
4.
The high temperature-related technologies are of broad applicability to missions to the inner solar system, while the avionics/navigation-related technologies are enabling for a wide range of small planetary missions (miniaturised orbiters and landers). In addition, the SMART-l lunar mission will demonstrate the use of an electric thruster for primary propulsion, combined with planetary swingbys.
ACKNOWLEDGEMENT Data included in this paper are largely baaed on input generated during the BepiColombo System & Technology Study by the industrial team (led by A. Anselmi of Alenia Spazio) and by ESA colleagues, in particular the BepiColombo Project Scientist (R. Grard) for the scientific part. Their contribution is gratefully acknowledged.