The BepiColombo Spacecraft, its Mission to Mercury and its Thermal



The BepiColombo Spacecraft, its Mission to Mercury and its Thermal
The BepiColombo Spacecraft, its Mission
to Mercury and its Thermal 5erikcation
1oger ) 6iKson anC Markus ScheKkKe irbus #S &mb' %rieCrichshaEen &ermany
!eOi"oKombo is a Ioint $S ) 7 mission to OKace 2 sOacecraEt in orbits arounC Mercury
Airbus DS is prime contractor for the European industrial part
Functional Breakdown
From Launch to Mercury Orbit
Mission Objectives
3hD main sciDntijc obIDctiUDs oE thD !DOi"oKombo mission to
MDrcurX comOrisD thD inUDstigation oE thD origin anC DUoKution
oE a OKanDt cKosD to thD OarDnt star anC a comOrDhDnsiUD stuCX
oE thD OKanDt itsDKE 3his ViKK bD achiDUDC bX OKacing tVo sOacD
craEt in CiEEDrDnt OoKar orbits arounC MDrcurX
!y courtesy oE ) 7
The )aOanese MMO Mercury
MagnetosOheric Orbiter carries
instruments Vith a totaK oE 11 sensors
The $uroOean MPO Mercury PKanetary Orbiter carries
11 instruments some Vith muKtiOKe sensors
MPO continuously nadir pointing:
• Nadir pointing provides maximum science return with continuous
planet observation
• Causes sun illumination of “all” spacecraft faces
¨Solution needed to provide a heat rejection radiator
• rianD $" into DscaOD orbit
Spacecraft Driving
MMO spins when free Ļying, therefore
needs shading during the 3-axis stabilised cruise phase:
• A sunshade and interface structure are required
¨a separate module is needed
3hD M"2 Mercury ComOosite SOacecraEt consists oE 4
oOtimiseC moCuKes
MPO l is oOtimiseC Eor its oOerationaK mission
• PerEorms commanC anC controK Eor M"2
• PerEorms aOOroach OroOuKsion anC Mercury orbit KoVering
Deceleration of 7 km/s needed to reach the innermost planet
• Planetary gravity assists are used to provide braking and
4.4 km/s braking is provided by electric propulsion
(requiring 10 kW power)
¨a separate module is needed
Operations and Control Requirements:
• Communication delays: maximum one-way signal time
14 minutes
• Solar conjunctions of 20 days in cruise and 7 days in
Mercury orbit – no ground contact possible
• Temperature control by maintenance of safe attitude
(especially of solar arrays)
Interplanetary Cruise Phase
• 1 W $arth 2 W 5Dnus anC W MDrcurX graUitX assist manoDuUrDs
• $KDctric ProOuKsion Eor braking bDtVDDn graUitX assists shortDns transEDr
• M3M sDOaration on 111202 aEtDr 1 orbits arounC thD sun
Mercury Approach Phase
• -aUigation bDEorD caOturD thDn orbit KoVDring
• %rDD graUitX caOturD on 01012024
• 1000 ms manoDuUrDs ODrEormDC bX MPO
• MMO sDOaration in MMO orbit
• MO2(% sDOaration
• #DscDnt to MPO orbit
Operational orbits around Mercury can only be inertially ĺxed
¨polar orbit of 0° inclination chosen:
• Orbit offset to manage thermal environment.
MPO: 480 km x 1,500 km, MMO: 590 km x 11,640 km
• Apoherm towards sun at perihelion to constrain planet
IR load to 5200 W/m2 (5400 W/m2 at aphelion)
Communications System
• 7 anC *a banCs Eor 10 &bityr science Cata anC 77
7*a *a*a ranging
• ntennas oE titanium to surUiUe thermaK enUironment
Power System
• SoKar rrays oOerating at 10¦" TemOerature KimiteC by tiKting
OaneKs unCer O"S controK to ¦ Erom sun
• 1 OS1s on MPO SoKar rray
AOCS (AttitudeControl)
• "ontroK oE sOacecraEt anC soKar array attituCes in an
enUironment Vhere 10 seconC CeKay can cause oUerheating
• "aters Eor OhysicaK conjgurations
Thermal Environment Boundary Conditions due to Mercury’s
proximity to the Sun:
• Solar intensity varying between 6,300 W/m2 and 14,500 W/m2
(> 10 solar constants), plus IR >5,200 W/m2
• -ominaK mission 1 $arth year l 2024
• $WtenCeC mission 1 $arth year Eoreseen
Data Management
• %"$ %aiKure "ontroK $Kectronics ensures continueC O"S
saEe oOeration Curing reboot oE main comOuter
• %irst sOacecraEt Vith netVork aOOKication oE SOaceVire
interEaces Eor science Cata
Electric Propulsion System
• 4 W 14 m- T ion thrusters oOerateC singKy or in Oairs
IMH attachment
Launch mass
MPO mass on orbit
MLI comOrises a singKe -eWteK outer Kayer Vith CimOKeC titanium
Kayers seOarateC by gKass sOacers %reeKy suOOorteC oUer Kengths
oE uO to 2 m
The LSS Large SOace SimuKator at $ST$" is being useC to test anC
UeriEy each moCuKe
• The LSS Vas moCijeC anC has suOOorteC !eOi"oKombo since
SeOtember 2010 Vhen the MMO thermaK moCeK Vas testeC
n intensity oE soKar constants is achieUabKe at Š2 m
• The MOSI% MMO Vere successEuKKy testeC thereaEter
• The MPO STM EoKKoVeC a year Kater shoVing notabKe CeUiations Erom
the OreCicteC OerEormance reUieV oE the CetaiKeC Cesign anC the
MLI construction Vas OerEormeC
• In autumn 2012 a Karge scaKe test 2 W m samOKes UerijeC the
imOroUeC MLI Cesign Eor the MPO kightmoCeK
• The MTM STM Vas successEuKKy testeC in sOring 201 its P%M test
is stiKK to come
• The EuKKy ePuiOOeC MPO P%M Vas testeC in -oUember 2014 This
conjrmeC the imOroUeC OerEormance oE the thermaK Cesign anC aKso
the correct Eunctioning oE the eKectricaK systems oUer the mission
temOerature ranges
contains embeCCeC anC surEace heatOiOes anC uses MLI CeriUeC
Erom the MPO Cesign
MPO hightemperature MLI kWation
4120 kg
1240 kg
Thermal 5erikcation Programme
– including MPO light model
Thermal Design for the severe thermal environment
The MPO unCergoes a kiOoUer manoeuUre tVice Oer Mercury
year in orCer to OroUiCe a singKe raCiator surEace Eor heat reIection
• 'eatOiOes are embeCCeC in the ePuiOment mounting OaneKs anC
the raCiator OaneK to transEer anC Cistribute heat
• LouUres in Eront oE the raCiator rekect the OKanet inErareC raCiation
VhiKst aKKoVing the raCiator a UieV to sOace
• The entire MPO boCy is coUereC Vith high temOerature ML(
CeUeKoOeC Eor !eOi"oKombo in orCer to combat temOerature anC
restrict heat inOut into the boCy
Outer heat shieKC comOrises 2 Kayers oE -eWteK ceramic cKoth
EoKKoVeC by 11 aKuminium Kayers 2 aKuminiseC 4OiKeW Kayers
anC 10 aKuminiseC MyKar Kayers in Oackets comOKete
the MLI
The -eWteK Kayers reach 0¦"
MTM Mercury TransEer MoCuKe
• ProUiCes braking by means oE eKectric OroOuKsion
• ProUiCes oUeraKK OoVer source Curing cruise
• "hemicaK OroOuKsion Eor naUigation anC O"S
The CriUing rePuirements haUe imOacts beyonC the mechanicaK
anC thermaK systems
M"S conjguration anC seOarations aEter years oE cruise
• The centraK structure oE the M"S at Kaunch is comOoseC oE mo
CuKe structures IoineC by intermoCuKe eKements IM' Inter
MoCuKe HarCVare l incK 2 W 2 eKectricaK connections
• 4Ooint attachments at MTMMPO anC MPOMOSI% interEaces
emOKoyeC to enabKe minimisation oE Oarasitic heat inOut once in
Mercury orbit
The IM' eKements Oass through the #T" DeOKoyabKe ThermaK
CoUer Erames to connect the structures
Eter seOaration the MLI Ciscs oE the
#T"s are CeOKoyeC to thermaKKy cKose
the aOertures in the main MLI
MOSIF MMO Sun2hieKC anC InterFace Structure
• 3hermaK Orotection Eor the MMO
• MechanicaK anC eKectricaK interEaces Eor the MMO
Subsystem and Hardware
Module Attachment and Separation
MPO Mercury Orbit Phase
• 2Oins Curing its oOerationaK mission
• (s OassiUe Curing cruise
MPO PFM - thermal testing in LSS
MOSIF + MMO – thermal test models in LSS
4th Lunar anC PKanetary Science "onEerence The 6ooCKanCs TW
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification. Roger J. Wilson1 and
Markus Schelkle1, 1 Airbus Defence and Space, Friedrichshafen, Germany ([email protected])
The spacecraft and subsystem designs are strongly driven by the severe demands of the thermal environment experienced at Mercury (whilst the same
solar intensities are also experienced during cruise)
and by the staging necessary to evolve the mechanical/electrical configuration from the 4-module composite at launch to the two free-flying orbiters.
Equipment not required in Mercury orbit is ejected
before capture into orbit around the planet. Due to the
nature of the MPO mission (in a low, inertial, polar
orbit) all external equipments are of bespoke hightemperature design.
BepiColombo is an interdisciplinary mission performed in a partnership between ESA (European
Space Agency) and JAXA (Japan Aerospace Exploration Agency). The mission aims to place 2 spacecraft in complementary orbits around Mercury and
perform scientific investigations of the planet. The
BepiColombo scientific results will add to the
knowledge already gained from the Mariner 10 flybys and the Messenger in-situ measurements.
JAXA provides the MMO (Mercury Magnetospheric Orbiter), whilst Airbus Defence and Space is
prime contractor for ESA, providing the MPO (Mercury Planetary Orbiter) and all other spacecraft hardware. The scientific payload is provided by national
This paper provides an overview of the mission
(including its trajectory to Mercury), spacecraft design (in particular its staging characteristics) and the
specific solutions implemented in the spacecraft
subsystems to meet the peculiar BepiColombo needs.
It further addresses the thermal testing performed to
verify the ability of the spacecraft and its equipments
to survive the harsh thermal environments experienced during the cruise phase and in Mercury orbit.
Mercury is the innermost planet of the solar system, is therefore difficult to reach and until now has
been visited by only two spacecraft. Mariner 10 flewby Mercury in March 1974, September 1974 and
March 1975 and Messenger has been in orbit around
Mercury since March 2011.
The BepiColombo mission will be the first European mission to Mercury. It was named after the
Italian scientist Giuseppe “Bepi” Colombo (1929 –
1984), who proposed the trajectory of Mariner 10,
and discovered the planet’s 3:2 spin-orbit resonance.
The BepiColombo mission will place the European MPO (Mercury Planetary Orbiter) and the Japanese MMO (Mercury Magnetospheric Orbiter) in low
polar orbits around Mercury. The orbiters will be
delivered to Mercury by means of a combined launch
aboard an Ariane 5 ECA and with the assistance of
dedicated propulsion and protection modules during
the 7 year cruise phase (transfer to Mercury). The
cruise phase includes multiple planetary gravity assists in addition to the braking provided by the spacecraft.
Figure 1:
The Mercury Composite Spacecraft
Whilst addressing the above topics in more detail,
this paper also pays particular attention to the verification of the thermal control performance. In total, 8
module sized tests will be performed (starting in
autumn 2010 with the MMO under JAXA responsibility) at up to 8 solar constants in the LSS (Large
Space Simulator) at ESTEC.
Overview of the Scientific Objectives
The BepiColombo mission will perform scientific
investigations [ 1 ] of Mercury in the areas of:
 Origin and evolution
 Interior, structure, geology, composition
 Exosphere composition and dynamics
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
 Magnetosphere structure and dynamics
 Origin of Mercury's magnetic field
 Test of Einstein's theory of general relativity
recorded the major drivers to the spacecraft design,
including an indication of the direct consequences
Two spacecraft in single launch
The MPO and MMO shall be launched in a single
flight configuration – whereby the MMO is passive
during the cruise to Mercury and (as a normally spinning spacecraft) will require thermal protection during the 3-axis stabilised cruise. This requirement
alone leads to some form of stacked configuration.
The payloads of the MPO and MMO are provided
by national agencies, with the objectives and leaderships as per Table 1.
The nadir pointed attitude of the MPO is aimed at
providing continuous viewing of Mercury for the
payload over the complete orbit. The only restrictions
to payload operation result from temperature and
power availability constraints around perihelion. The
mission thus enables visibility of the complete Mercury surface.
Deceleration to reach Mercury
Mercury is the innermost planet of the solar system and a braking ∆V of ca. 7 km/s is required to
arrive at Mercury with a velocity suitable for capture
into orbit. Planetary gravity assist manoeuvres can be
used to achieve this deceleration but then time and
the number of fly-bys are both large. For BepiColombo it was decided to implement an Electric
Propulsion System using 2 x 145 mN ion thrusters to
reduce the cruise phase duration. Since operating
these 2 thrusters requires 10 kW electrical power a
dedicated propulsion module was conceived for the
cruise phase. This module then serves as the first
stage of the flight composite – to which were added
all functions not required in Mercury orbit.
Table 1: MPO and MMO Payloads
BepiColombo Laser Altimeter
Italian Spring Accelerometer
Magnetic Field Investigation
Mercury Radiometer and
Thermal Imaging Spectrometer
Mercury Gamma-Ray and
Neutron Spectrometer
Mercury Imaging X-ray Spectrometer
Mercury Orbiter Radio Science Experiment
Probing of Hermean Exosphere by UV Spectroscopy
Search for Exospheric Refilling and Emitted Natural
Solar Intensity X-ray and
particle Spectrometer
Spectrometers and Imagers
for MPO BepiColombo Integrated Observatory
Mercury Magnetometer
Mercury Plasma Particle Experiment
Plasma Wave Instrument
Mercury Sodium Atmospheric Spectral Imager
Mercury Dust Monitor
Co-PI: N. Thomas, CH
Co-PI: T. Spohn, D
PI: V. Iafolla, I
PI: K.H. Glassmeier
PI: H. Hiesinger, D
PI: I. Mitrofanov, RUS
Free Gravity Capture
A free gravity capture, using the Mercury – Sun
Lagrange point, was implemented to enable a safe
capture without the need for time-critical operations.
The approach navigation must ensure suitable trajectory and velocity to guarantee the capture. The initially weak capture must be consolidated by apoherm
lowering, but this is also a not time critical activity.
PI: E. Bunce, UK
PI: L. Iess, I
PI: E. Quémerais, F
PI: S. Orsini, I
Operational Orbits around Mercury
The MPO will be in an inertially fixed polar orbit
of 0° inclination with periherm distance 480 km and
apoherm distance 1500 km. The orbit is orientated to
place the apoherm towards the sun at perihelion,
thereby crossing the equator far from the planet’s
sub-solar point which reaches over 400°C. The MPO
has an orbit period of 2.3 hours.
The MMO will be placed in a coplanar orbit with
the MPO, with periherm distance 590 km and
apoherm distance 11,640 km. The MMO has an orbit
period of 9.3 hours.
PI: J. Huovelin, FIN
PI: E. Flamini, I
PI: W. Baumjohann, A
PI: Y. Saito, JPN
PI: Y. Kasaba, JPN
PI: I. Yoshikawa, JPN
MPO Observation Concept and Attitude
The MPO will fly permanently nadir pointed to
provide continuous scientific observation.
Thermal Environment Boundary Conditions
Mercury orbits the sun with perihelion distance
0.31 AU and aphelion distance 0.47 AU. Resulting
PI: H. Shibata, JPN
Mission and Spacecraft Driving Requirements
The spacecraft design is greatly influenced by the
mission and the environment of Mercury. Below are
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
from this orbit is a solar intensity varying between
6,300 W/m2 and 14,500 W/m2 (> 10 solar constants).
The Mercury sub-solar point reaches >400°C at
perihelion – which results in an infrared intensity of
5,200 W/m2 on the MPO. The choice of orbit results
in an infrared intensity of 5,400 W/m2 on the MPO
when Mercury is at aphelion.
The inertially fixed orbit entails that, over a Mercury year, an orbiting body experiences sun illumination from all directions. At any time, the MPO requires at least 1 surface for heat rejection. A single
radiator surface is made available for thermal control
by means of flip-over manoeuvres which rotate the
spacecraft 180° around the nadir-pointed axis at each
perihelion and aphelion (thereby inverting the flight
Figure 2:
BepiColombo the maximum one-way signal time is
14.4 minutes.
Mercury is regularly behind the sun, as is the
spacecraft during the cruise phase: these solar conjunctions inhibit communication between spacecraft
and ground. The maximum conjunction during cruise
is 20 days and 7 days when in MPO orbit, during
which the spacecraft must operate without assistance
from ground.
In addition to this long-term autonomy requirement, the avionics must ensure that a safe attitude is
maintained, thereby avoiding that damaging temperatures are generated on external surfaces. Critical
durations are in the order of only a few 10s of seconds.
System Breakdown and Functional Apportionment
From the Mission and Spacecraft Driving Requirements it is quickly clear that the BepiColombo
launch configuration is comprised of 4 modules. As
the mission evolves, then the number of modules
decreases. These evolving configurations are a composite of x modules, hence the various configurations
are known as MCSn for Mercury Composite Spacecraft (where “n” represents the states L for launch, C
for Cruise, A for Approach and O for Orbit).
The MPO, whatever tasks it may perform during
cruise, is ultimately a free-flying spacecraft containing all the capabilities needed to perform its scientific
mission – for which careful optimisation was necessary when considering the thermal environment. The
MPO therefore contains most of the capabilities also
needed during cruise. In order not to compromise the
MPO design by taking unnecessary hardware into
Mercury orbit, hardware needed solely for cruise is
accommodated in a separate MTM (Mercury Transfer
The MMO is eventually also a free-flying spacecraft containing all the capabilities needed to perform
its scientific mission. However with the spacecraft
capabilities controlled from the MPO during cruise,
the MMO then remains passive throughout (apart
from periodic check-outs). Since the MMO is a normally spinning spacecraft, it requires thermal protection during the 3-axis stabilised cruise. The 4th module of the MCS derives from the MMO’s needs: the
MOSIF (MMO SunShield and InterFace Structure)
providing thermal protection as well as all the interfaces between MPO and MMO.
The MCSL and MCSC are composed of:
 MPO (Mercury Planetary Orbiter)
• Spacecraft optimised for its operational mission
The MPO orbit around Mercury
The thermal environment of the MPO orbit necessitates that all external items of the MPO must be
capable of withstanding high solar intensity and high
temperatures, whilst in addition the MLI (multi-layer
insulation) must restrict the heat input to the spacecraft body to manageable levels.
Since the cruise trajectory entails repeatedly visiting the perihelion distance of Mercury from the sun
(whilst progressively reducing the aphelion distance)
the MTM and MOSIF are also subjected to > 10 solar
constants. As the MCS flys in a sun pointed attitude
the constraints are less severe than for the MPO,
nevertheless the sunward face and the solar arrays
must withstand high solar intensity.
Operations and Control Requirements
As is typical with inter-planetary missions, communication and control are constrained by the oneway signal time between spacecraft and Earth which
inhibits real-time commanding from ground. For
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
• Performs command & control for MCS
(with only minor hardware modification for
MCS configurations, notably the size of the
reaction wheels to control the MCS)
• Chemical Propulsion System not used during cruise, however after MTM separation
the MPO performs approach propulsion and
apoherm lowering of Mercury orbit
 MMO (Mercury Magnetospheric Orbiter)
• Spins during operational mission
• Is passive during cruise – apart from checkouts
 MOSIF (MMO SunShield and InterFace
• Thermal protection for the MMO
• Mechanical interface for the MMO
• Harness routing between MPO and MMO
 MTM (Mercury Transfer Module)
• Provides MEPS (MTM Electric Propulsion
System) plus chemical propulsion (for cruise
AOCS and navigation correction)
• Provides power for electric propulsion system and for MPO +MMO
• Separated before capture into Mercury orbit
The MCSA is created upon MTM separation. On
reaching the MMO orbit, the MMO is released to
create the MCSO. The MOSIF is ejected shortly
afterwards to leave the MPO.
Spacecraft Configuration
MPO Configuration
The MPO was conceived and optimised to meet
the needs of its operational orbit, before undergoing
minor adaptations to fit within the MCS.
The spacecraft envelope is compact in order to
minimise the heat input through the MLI, and hence
the heat load to the single radiator. The dedicated
radiator is sized for a maximum heat rejection (at
maximum temperature) of 1500 W – where 20% of
this quantity is from parasitic heat leaks into the
spacecraft. The radiator locally fills the Ariane fairing. Heatpipes transport heat from the equipment
panels to the radiator and radiator heatpipes provide
heat distribution. The equipment mounting and the
heatpipe network are concentrated on 2 aluminium
sandwich panels which constitute the major elements
of the central double-H structure (completed by 2
smaller lateral panels). The panel interfaces of the
double-H primary structure provide the nodes for the
4-point interfaces to the MTM and MOSIF.
Figure 4:
Figure 3:
MPO – showing equipment panel perpendicular to radiator
The majority of the instruments point to nadir
and, in particular, a CFRP sandwich optical bench
accommodates those instruments with tight pointing
requirements co-aligned with the 3 Star Trackers and
an Inertial Measurement Unit. The Star Trackers
view through the radiator to avoid sun illumination.
The steerable MGA (medium gain antenna) and
the Ø 1100 mm HGA (high gain antenna) provide
communication links in the MPO orbit when the
Earth can be located in any direction with respect to
Composition of MCS, showing module
functions and contributions to MCS
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
the spacecraft body. Both antennas are accommodated to provide viewing capability in MPO orbit and
during cruise. The MGAMA is mounted on a long
boom which can rotate about the boom axis as well
as about the boom attachment point – this 2-axis
rotation provides the capability to view around the
obstruction of the MPO body and also past the obstructions generated by the MTM and MOSIF in the
MCS configuration. The HGAMA has 2-axis rotation
capability providing it with a large field of view
around the MPO body, whilst having a restricted
view in the MCS configuration.
Figure 5:
The MMO in stowed configuration can be accommodated in a cylinder Ø 1800 x 1200 mm.
MOSIF Configuration
The MOSIF acronym (MMO SunShield and InterFace Structure) very much describes its functionality, which then leads to its configuration.
The MOSIF consists of two major assemblies:
 The Adapter has at its centre the circular interface to the MMO, the arms provide the attachments for the Sunshield, the arms provide
the 4 separable attachments to the MPO
 The Sunshield structure supports the MLI
which shades the MMO. The Sunshield is
truncated at 18° to allow MCS/MOSIF rotation towards the sun without illuminating the
MMO. The Sunshield is conical at 16° halfangle to tolerate wobble of the MMO during
its slowly spinning separation
MPO deployed configuration
The MPO is 1.7 m high by 3.6 m across the radiator and in its flight configuration with MLI installed
can be seen in Figure 1and Figure 15
MMO Configuration
The MMO has an octagonal body, deployable antennas and deployable instrument booms. The circular high-gain antenna is deployed before separation
from the MCSO, the other deployments are performed post-separation. The MMO provides a circular, bolted interface to the MOSIF. The Spin Ejection
Device provided by JAXA remains bolted to the
Figure 7:
MTM Configuration
The configuration of the MTM is mainly driven
by the MEPS. The power demand of 5 kW per thruster (of which two will operate simultaneously) leads
to the large solar arrays, whilst the efficiency of the
power system and the MEPS electronics determine
the dissipated heat to be rejected. Three radiator panels are needed which are themselves in a 20° wedge
configuration, allowing the MCS to rotate slightly
about its longitudinal axis whilst avoiding sun illumination of the radiators. The 4 MEPS thrusters are
accommodated within the launcher interface ring,
where they can thrust through the spacecraft centre of
mass, and are protected by a sunshade to allow rotation of the thrust vector towards the sun without the
thrusters being illuminated.
Photo courtesy of JAXA
Figure 6:
MOSIF – with MMO installed
MMO flight model
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
Figure 8:
The Mission from Launch to MPO Orbit
The MCS will be launched by an Ariane 5 ECA
from the Centre Spatial Guyanais in Kourou, French
Guiana in December 2016. At 4121 kg the MCS will
constitute the lightest single payload launched by an
Ariane 5 ECA, nevertheless the full capability of the
launcher is required to deliver the needed escape
MTM deployed configuration
MCS Configuration
The MCS configuration is represented by a stacking of the above described modules.
Figure 9:
Figure 10: From launch to MPO orbit showing
gravity assists and separations
Cross-section through the MCS
Table 2: Overall Mass Budget
Mass Budget
Dry Mass
A deceleration of 7 km/sec is required to reach
Mercury. 4.2 km/sec will be provided by the MEPS –
enabling the spacecraft to more rapidly synchronise
with each subsequent planet gravity assist. The remainder of the braking will be achieved by gravity
assists at Earth, Venus and Mercury. The cruise until
MTM separation in November 2023 takes 6 years 10
months, includes 17 orbits around the sun and 8 gravity assist manoeuvres: 1 x EGA, 2 x VGA and
5 x MGA. During the electric propulsion thrust phases the MCS is orientated with the thrusters pointing
Cruise CPS
Total Mass
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
in the flight direction to provide braking and with
thruster steering supporting the attitude control system. During coast phases the MCS is stabilised by
rotation around the sun vector, thereby minimising
chemical propellant consumption.
After MTM separation, the MPO takes on the
propulsion tasks of the MCSA for the MOI (Mercury
Orbit Insertion), which begins with navigation to the
Mercury-Sun Lagrange point. A free gravity capture
is employed, avoiding the risks of time-critical manoeuvres prior to Mercury capture. The MCSA is
weakly captured into an initial orbit 178,000 x 670
km. The MOI includes a series of 15 orbit lowering
manoeuvres, the first of which is needed to stabilise
the capture, thereafter the manoeuvres lead to the
respective operational orbits of the MMO and MPO.
The MPO orbit will be reached in January 2024.
From the many changes of flight configuration,
the number of actuators employed and the stringent
safe and survival modes the AOCS consists of 17
operational modes.
Data Management System (DHS)
The basic MPO DHS comprises redundant OBCs
(On-Board Computer) and an internally redundant
SSMM (Solid State Mass Memory) for payload and
spacecraft data storage. A MIL-STD-1553B bus is
used for spacecraft telemetry and telecommand whilst
all payload TM/TC and science data interfaces use
SpaceWire. BepiColombo is the first spacecraft with
a network application of SpaceWire interfaces.
The MPO provides all the intelligence during
cruise and is enhanced with additional data buses to
the MMO and MTM for this purpose.
Permanent availability of a functioning processor
to guarantee safe and prompt attitude control is provided in Survival Mode by redundant FCEs (Failure
Control Electronics) which take over the control
functions in the event of an OBC reboot. The FCEs
retain control for 7 minutes after which it is taken
over by the reconfigured OBC.
Subsystem Designs and BepiColombo specific
Attitude and Orbit Control System (AOCS)
For MPO science operations the AOCS must provide continuous nadir pointing whilst meeting accuracy and stability requirements. Two Star Trackers
(plus a 3rd for redundancy) and an Inertial measurement Unit (IMU) are co-mounted with instruments on
an optical bench whilst 4 reaction wheels serve as
actuators (with 5 N thrusters used for wheel offloading). The AOCS also controls the thermally critical orientation of the solar array and the 22 N thrusters for orbit manoeuvres during Mercury Orbit Insertion.
This basic AOCS is enhanced with sun sensors
for survival mode and is further enhanced for the
MCS configuration when the MEPS thrusters and
MTM 10 N thrusters serve as actuators. As for the
MPO, MTM Solar Arrays are also thermally controlled by the appropriate orientation.
During cruise the AOCS controls the MEPS
thruster orientation and corresponding MCS attitude
as required by the uploaded mission timeline – with
fine pointing of the MEPS thrusters minimising momentum accumulation by the reaction wheels.
The thermal environment experienced in the
MPO orbit and during cruise allows (for a number of
thermally critical items) only deviations from nominal attitudes in the order of seconds before overheating and damage occurs. In the event of an OBC reboot, the Survival Mode will be entered and the
AOCS control will be transferred to the FCE and a
second IMU. In Survival Mode the AOCS uses Sun
Sensors as the attitude reference . Different Survival
attitudes apply for the various spacecraft configurations.
Communications System
The communications system [ 2 ] is equipped
with 2 x LGA (Low Gain Antenna), a 2-axis steerable
MGA (Medium Gain Antenna) and a 2-axis steerable
HGA (High Gain Antenna). The LGAs provide
spherical coverage, can support X-band downlink and
uplink near Earth, whilst also providing uplink reception over greater distances. The X-band horn MGA is
steerable around the MPO or MCS obstructions in
order to view Earth and is the primary antenna during
cruise. The Ø 1100 mm HGA is steerable around the
MPO and is the primary antenna during science operations when it supports up- and downlinks in both Xand Ka-band.
The newly developed DST (Deep Space Transponder) supports telecommanding uplink in X-band
with telemetry downlink in both X- and Ka-bands to
enable the downlink of 1550 Gb/year of science data.
The DST supports ranging in X/X band and X/Ka
band whilst Ka/Ka band ranging is provided with the
inclusion of the payload-provided MORE translator.
This ranging strategy is related to the Radio Science
Experiment and requires high stability of the HGA.
Power amplification is by TWTAs for both Xand Ka-bands. All antennas are exposed to the severe
thermal environment and are based on titanium. The
antenna pointing mechanisms for HGA and MGA are
capable of operating at 250°C.
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
tilts the array to a maximum of 75° from the sun. The
OSRs occupy 17% of the panel area. The tilted, single wing configuration is needed to achieve sun illumination whilst permanently nadir pointing throughout the Mercury year – this entails near continuous
rotation controlled by the AOCS.
During the cruise all power is provided by the
MTM solar array. The MTM provides a 100 V regulated bus for the electric propulsion system and a
28 V regulated bus for other equipment plus the
MPO/MMO. During gravity assist manoeuvres, when
in some cases eclipses occur, the MEPS is not active.
The MTM includes a 12 Ah lithium-ion battery for
damping of MEPS surge current (in the event of a
beam-out) and also provides power during the short
eclipses. The solar array provides 13 kW which is
particularly driven by the 2 x 5 kW demand of the
The MTM solar array is a high temperature design operating at maximum 190°C which uses the
same technologies as the MPO, but without the need
for in-plane thermally conductive CFRP facesheets.
The thermal control of the array is achieved by the
AOCS which tilts the array to a maximum of 76°
from the sun. The size of the array is driven by the
need to provide maximum power also at 0.31 AU.
Whilst approaching the sun the solar array output
initially increases, of course accompanied by an increase of temperature. Once the temperature has
reached 190°C (at about 0.5 AU) the array must be
tilted, thereby reducing its projected area and limiting
its output. The two wings total 40 m2 and have a
mass of 290 kg.
MTM Electric Propulsion System (MEPS)
The MEPS contains the 4 electric propulsion
(Kaufmann) thrusters, their power processing electronics, 4 thruster pointing mechanisms and the xenon storage and feed system.
The thrusters are the QinetiQ T6, Ø 22 cm,
145 mN thruster derived from the T5 flown on the
GOCE mission. The system is planned to operate
over 25 thrust arcs totalling 880 days, with the longest continuous operation being for 167 days. The
system typically operates using two thrusters.
Photo courtesy of QinetiQ
Figure 11: T6 thruster firing test
The thrusters are mounted on individual pointing
mechanisms which enable the thrust vector to point
through the MCS centre of mass – either for a single
thruster or the plane of an operating pair of thrusters.
By off-pointing of the thrust vector a moment is created for use by the AOCS for wheel off-loading.
The xenon system, with its flow control units,
pressure regulators and 3 tanks, is able to store and
deliver 580 kg of xenon which can provide 5400 m/s
delta V.
Chemical Propulsion Systems (CPS)
The MPO CPS is tasked with the 15 MOI manoeuvres and attitude control, for which it is
equipped with redundant 4 x 22 N and redundant
4 x 5 N thrusters. The 22 N thrusters are bi-propellant
whilst the 5 N thrusters are mono-propellant: these
are combined into the first dual-mode propulsion
system implemented on a European spacecraft. The
system uses hydrazine and MON (Mixed Oxides of
Nitrogen). 669 kg of propellant are carried, giving a
capability of 1000 m/s delta V plus attitude control.
The MTM CPS employs redundant 12 x 10 N bipropellant thrusters, using MMH (Mono-Methyl
Hydrazine) and MON (Mixed Oxides of Nitrogen).
This system is derived from Eurostar 3000 systems [
4 ]. As well as attitude control capability, the MTM
CPS can provide axial thrust for cruise navigation.
157 kg of propellant are carried, giving a capability
of 68 m/s delta V plus attitude control.
The MPO uses a 28 V regulated power bus [ 3 ].
After MTM separation the power is provided by the
MPO solar array which is kept edge-on to the sun
during cruise to minimise UV degradation. An 96 Ah
lithium-ion battery provides power during eclipses.
The solar array provides TBD W which is sufficient
for full payload operation and alternatively sufficient
to provide 300 W to the MMO until its separation.
The MPO solar array is a high temperature design
operating at maximum 190°C – for which components have required dedicated development. The
thermal control of the array is achieved by its design,
involving a mix of cells and OSRs (optical solar
reflector) enhanced by in-plane thermally conductive
CFRP facesheets, and control by the AOCS which
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
 7 dimpled titanium layers separated by glass
 It is freely supported over lengths of up to
2.5 m and must withstand the vibration and
acoustic environments of the launch
Thermal Control
The MPO TCS must regulate the equipment temperatures (achieving standard equipment levels) [ 5 ],
transfer heat to the single radiator, shield the radiator
from planet infrared illumination, reject 1200 W of
dissipated heat from the payload and spacecraft
equipments and reject up to 300 W of parasitic heat
which enters the MPO body. These functions are
achieved by means of:
 Heatpipes embedded in the equipment mounting panels to collect and transfer the heat the
radiator panel
 Spreader heatpipes in the radiator panel, thermally connected to the equipment panels by
90° linking heatpipes
 97 heatpipes are used, of which a few are 3dimensional hence difficult to test on ground
 Fixed louvres are mounted in front of the radiator to reflect the planet infrared radiation
away from the radiator whilst allowing the radiator an extensive view to space
 The entire MPO body is covered with high
temperature MLI developed for BepiColombo
 The outer heat shield comprises 2 layers of
Nextel ceramic cloth followed by 11 aluminium layers. The Nextel layers reach 380°C
 Moving inwards to lower temperature, 26 layers of aluminised Upilex are followed by 10
layers of aluminised Mylar
 Spacers of glass fibre and AAerofoam are
used to separate the layers in the 4 packets,
whilst kapton rosettes separate the packets
 The installed MLI has a thickness of 65 mm.
The total MLI mass is 94 kg.
The MTM TCS must regulate the equipment temperatures, distribute heat in the radiators, reject
2000 W of dissipated heat equipments and reject up
to 300 W of parasitic heat which enters the MTM
body. These functions are achieved by means of:
 Heatpipes embedded in the radiator panels
(which also serve for equipment mounting)
 The embedded heatpipe network is enhanced
by surface heatpipes
 63 heatpipes are used
 Derivatives of the high-temperature MLI are
Further MLI applications result from the stack
configuration and the separation interfaces of the
 Whilst the modules are protected as described
above, solar illumination gaps between modules can not be tolerated
 Elaborate Gap Closure MLI is implemented
between MTM-MPO and MPO-MOSIF. This
MLI is in contact with the MPO and is attached to the separating modules (see Figure
 The 4-point mechanical interfaces between
modules leave holes of Ø140 mm in MLI (to
these add 2 x Ø170 mm holes for the connectors at each interface). These holes are closed
by DTCs (Deployable Thermal Covers) containing MLI disks to drastically reduce the
heat load. The 12 DTCs are mounted between
the MLI layers with the cylinder and ring
(white coated) protruding through the MLI
Figure 12: Section through MPO MLI
The MOSIF MLI must shade the MMO and limit
the infrared heat load to the MMO. The MOSIF MLI
is characterised by:
 A single Nextel outer layer
Figure 13: Deployable Thermal Cover to close
separation apertures
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
Structure and Inter-Module Hardware
The MPO structure is dominated by the need of
aluminium heatpipes for thermal control reasons
(necessitating aluminium equipment mounting panels
and radiator) along with minimalisation of parasitic
heat inputs at the interfaces to MOSIF and MTM
after separation – where 4-point interfaces are implemented. The structure uses aluminium sandwich
throughout – with the exception of the CFRP sandwich Optical Bench. The configuration can be visualised from Figure 4.
The MOSIF structure is characterised by:
 The Adapter: a central cylindrical piece interfacing to the MMO, supported on box construction arms from the IMH interfaces to the
MPO – the arms extend further to support the
Sunshield, both parts use aluminium
 The Sunshield: an aluminium truss structure
of circular section tubes to support the MLI
Figure 14: MTM-MPO IMH element
Spacecraft Operations and Autonomy
Operations will generally be performed off-line,
using an on-board master schedule called the Mission
Timeline (MTL). The MTL is enhanced by On-Board
Control Procedures (OBCPs), file transfer capability,
high level command functionality, and a flexible onboard monitoring function.
The autonomous functions on-board are tasked
with the execution of the MTL whilst ensuring that
each item is in a safe and correct configuration by
means of monitoring and FDIR (Failure Detection,
Isolation and Recovery). The hierarchical FDIR enables isolation of unambiguously identified anomalies
by local reconfiguration. For other anomalies the
safety of the spacecraft is guaranteed by entry in
Survival then Safe Mode. Return to Normal Mode is
by ground command. The possible durations in
Safe/Survival modes are driven by the solar conjunctions. By this means, for example, it will be possible
to continue MEPS operation during solar conjunctions – for which back-up thruster configurations and
corresponding attitudes are stored on-board.
In the post-launch, near-Earth phase ground contact will be provided for up to 24 hours/day. During
cruise the contacts will be reduced to only 8
hours/week – except around planet fly-bys. In MPO
orbit a daily 10 hour pass is planned, of which 9
hours are allocated for data downlinking.
ESA’s Cebreros Ø35 m ground station is the primary station for all mission phases.
The MTM structure provides a circular launcher
interface to the Payload Adapter System 1666MVS,
from which a CFRP sandwich main cone extends
upwards to the 4-point interface to the MPO. The
xenon and CPS tanks are mounted off the main cone,
whilst CFRP sandwich floors and shear panels accommodate smaller items. The floors and shear panels support the aluminium radiators (carrying the
power and MEPS electronics) and the solar arrays.
The thruster floor, within the main cone, carries the
MEPS thrusters and pointing mechanisms.
The MTM, MPO and MOSIF structures provide
the core of the MCS structure at launch. These elements (MTM and MPO structures are separated by
130 mm, with a smaller gap to the MOSIF) are connected by Inter Module Hardware (IMH) completing
the MCS structure at launch and providing the separation functions after cruise. The IMH also includes
attached hardware and separable connectors for the
inter-module electrical connections (excluding to
MMO). To provide the structural performance over
the module-module distances, along with the separation functions and dynamics (both electrical and
mechanical), the IMH requires a mass of 47 kg.
Overview of the Integration and Verification
The AIV (Assembly Integration and Verification)
for BepiColombo is longer than other spacecraft.
More modules must be verified but these must also
be verified in the MCS configuration. With much
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
new hardware developed for BepiColombo, an extensive AIV programme was undertaken.
An ATB (Avionics Test Bench) was employed to
test and verify the avionics and software, using real
hardware in the loop.
An ETB (Electrical Test Bench) was employed to
test and verify MPO equipments working together in
a representative electrical and mechanical configuration. This enabled development and testing with
spacecraft hardware and instrument models and was
extended to electrically represent the MCS configuration with MTM and MMO representations. This
programme enabled extensive testing and risk reduction before embarking on the PFM programme.
An STM (Structural Thermal Model) was employed to demonstrate the functioning and performance of the all-new structures and thermal control
systems (using dummy electronic units). The sequence began in autumn 2010 with thermal testing of
the MOSIF + MMO and ended (in autumn 2010)
with MCS separation tests after successful vibration
and acoustic testing.
A PFM (Proto-Flight Model) for which electrical,
functional and EMC verification will be performed at
module and MCS level. The mechanical and thermal
tests will be repeated to provide final confirmation or
demonstrate adequate workmanship. The modules
were transferred to ESTEC in July 2014 for the continuation of assembly and the verification programme.
mance. The high temperatures experienced warranted
a review of the thermal design implementation and in
particular reconsideration, and verification, of the
MLI design.
Figure 15: MPO PFM in the LSS
The flight model MLI design includes an extra
Nextel layer, more layers in total and a thermally
improved attachment system. This design was successfully verified in a large scale back-to-back comparison against the STM design (2 x 3 m samples of
MLI tested in the LSS) in October 2012.
The MTM STM was tested in spring 2013 and the
performance fitted satisfactorily to the predictions.
A thermal balance / thermal vacuum test (TB/TV)
of the fully equipped MPO PFM was performed in
November 2014. This validated the performance of
the thermal design modifications, along with the
improved MLI, and also the correct functioning of
the electrical systems over the mission temperature
ranges. The temperatures measured inside the MPO
were typically 4°C cooler than predicted.
Still to be performed in the LSS are a test of the
complete high gain antenna and TB/TV test of the
Thermal Verification
Each module is of a completely new mechanical
and thermal design and employs the newly developed
high temperature MLI. The thermal environment to
be experienced is severe, when compared with nearEarth or the colder environments experienced at the
outer planets. The need to demonstrate the thermal
performance at solar intensities close to the >10 solar
constants to be experienced was considered paramount.
The LSS (Large Space Simulator) at ESTEC [ 6 ]
was modified and has supported BepiColombo since
September 2010. All 19 sun-simulating lamps are
used at high power and the mirrors of the normally
parallel Ø6 m beam have been adjusted to conical
such that more than 8.5 solar constants illuminate the
test object with Ø2.7 m.
The MMO thermal model was the first tested, in
September 2010, and was followed shortly afterwards
by the MOSIF + MMO combination (Figure 7).
The MPO STM was tested in September 2011
and the value of the test was clearly demonstrated by
the notable deviations from the predicted perfor-
The mechanical performance of the MCS was
demonstrated on the STM, along with the ability to
separate the modules. The thermal performance of the
MOSIF and MTM were demonstrated on the respective STMs. The avionics and electrical systems have
been developed and demonstrated on their respective
The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification
46th Lunar and Planetary Science Conference, The Woodlands, March 2015
test benches. The PFM programme is now in full
flow, with the capabilities of the MPO having been
extensively verified during its TB/TV test. With the
arrival of MMO flight model at ESTEC in April 2015
the programme is well on its way towards departure
for Kourou 5 months before launch in December
MPO and MMO in orbit around Mercury
[ 5 ] Federica Tessarin, Domenico Battaglia and
Tiziano Malosti, Daniele Stramaccioni, Jürgen Schilke, “Thermal Control Design of Mercury Planetary
Orbiter”, AIAA 2010-6090, 40th International Conference on Environmental Systems
[ 6 ] Large Space Simulator (LSS), at ESTEC,
Noordwijk, The Netherlands
[ 1 ] Johannes Benkhoff, Jan van Casteren,
Hajime Hayakawa, Masaki Fujimoto, Harri Laakso,
Mauro Novara, Paolo Ferri, Helen R. Middleton,
Ruth Ziethe, “BepiColombo—Comprehensive exploration of Mercury: Mission overview and science
goals”, Planetary and Space Science 58 (2010) 2–20
[ 2 ] M. Mascarello, P. Pablos, R. Heinze,
A Busso, R. Carbone, “The BepiColombo X/X/Ka
TT&C Subsystem”, TTC 2010, 5th ESA Workshop on
Telemetry, Tracking & Communications Systems for
Space Applications, ESA-ESTEC 21 23 September
[ 3 ] Pierluigi Morsaniga, Giuseppe Gervasio,
Giuseppe Cuzzocrea, “BepiColombo Electrical Power System”, Proc. ‘9th European Space Power Conference’, Saint Raphaël, France, 6–10 June 2011
(ESA SP-690, October 2011)
[ 4 ] Eurostar 3000 satellite, Airbus DS