The BepiColombo , its Mission to and its Thermal 5erik cation 1oger ) 6iKson anC Markus ScheKkKe irbus #S &mb' %rieCrichshaEen &ermany

!eOi"oKombo is a Ioint $S  ) 7 mission to OKace 2 sOacecraEt in arounC Mercury DS is prime contractor for the European industrial part

From Launch to Mercury Functional Breakdown Mission Objectives

3hD main sciDntij c obIDctiUDs oE thD !DOi "oKombo mission to MDrcurX comOrisD thD inUDstigation oE thD origin anC DUoKution oE a OKanDt cKosD to thD OarDnt 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 The $uroOean MPO Mercury PKanetary Orbiter carries Magnetos Oheric Orbiter carries 11 instruments some Vith muKtiOKe sensors  instruments Vith a totaK oE 11 sensors Spacecraft Driving Requirements MPO continuously nadir pointing: • Nadir pointing provides maximum return with continuous MMO spins when free Ļ ying, therefore observation needs shading during the 3-axis stabilised cruise phase: • Causes illumination of “all” spacecraft faces • A sunshade and interface structure are required ¨Solution needed to provide a heat rejection radiator ¨a separate module is needed

3hD M"2 Mercury ComOosite SOacecraEt consists oE 4 oOtimiseC moCuKes MPO l is oOtimiseC Eor its oOerationaK mission Deceleration of 7 km/s needed to reach the innermost planet • PerEorms commanC anC controK Eor M"2 Mercury: • PerEorms aOOroach OroOuKsion anC Mercury orbit KoVering • Planetary assists are used to provide braking and 4.4 km/s braking is provided by electric propulsion MMO (requiring 10 kW power) • 2Oins Curing its oOerationaK mission Launch Operations and Control Requirements: ¨a separate module is needed • (s OassiUe Curing cruise • rianD  $" into DscaOD orbit • Communication delays: maximum one-way signal time 14 minutes MOSIF MMO Sun2hieKC anC InterFace Structure Interplanetary Cruise Phase • Solar conjunctions of 20 days in cruise and 7 days in • 3hermaK Orotection Eor the MMO • 1 W $arth 2 W 5Dnus anC  W MDrcurX graUitX assist manoDuUrDs Mercury orbit – no ground contact possible • MechanicaK anC eKectricaK interEaces Eor the MMO • $KDctric ProOuKsion Eor braking bDtVDDn graUitX assists shortDns transEDr • control by maintenance of safe attitude • M3M sDOaration on 1 11 202 aEtDr 1 orbits arounC thD sun (especially of arrays) MTM Mercury TransEer MoCuKe Operational orbits around Mercury can only be inertially ĺ xed • ProUiCes braking by means oE eKectric OroOuKsion ¨ Mercury Approach Phase of 0° inclination chosen: • ProUiCes oUeraKK OoVer source Curing cruise • -aUigation bDEorD caOturD thDn orbit KoVDring • Orbit offset to manage thermal environment. • "hemicaK OroOuKsion Eor naUigation anC O"S • %rDD graUitX caOturD on 01 01 2024 MPO: 480 km x 1,500 km, MMO: 590 km x 11,640 km • 1000 ms manoDuUrDs ODrEormDC bX MPO • Apoherm towards sun at perihelion to constrain planet Subsystem and Hardware • MMO sDOaration in MMO orbit IR load to 5200 W/m2 (5400 W/m2 at aphelion) • MO2(% sDOaration Implications • #DscDnt to MPO orbit The CriUing rePuirements haUe imOacts beyonC the mechanicaK Module Attachment and Separation anC thermaK systems Communications System M"S conj guration anC seOarations aEter  years oE cruise • 7 anC *a banCs Eor 10 &bityr science Cata anC 77 • The centraK structure oE the M"S at Kaunch is comOoseC oE mo 7*a *a*a ranging CuKe structures IoineC by inter moCuKe eKements IM'  Inter • ntennas oE titanium to surUiUe thermaK enUironment MoCuKe Har CVare l incK 2 W 2 eKectricaK connections • 4 Ooint attachments at MTM MPO anC MPO MOSI% interEaces Power System emOKoyeC to enabKe minimisa tion oE Oarasitic heat inOut once in • SoKar rrays oOerating at 10¦" TemOerature KimiteC by tiKting Mercury orbit OaneKs unCer O"S controK to ¦ Erom sun The IM' eKements Oass through the #T" DeOKoyabKe ThermaK • 1 OS1s on MPO SoKar rray CoUer Erames to connect the structures Eter seOaration the MLI Ciscs oE the AOCS (AttitudeControl) #T"s are CeOKoyeC to thermaKKy cKose • "ontroK oE sOacecraEt anC soKar array attituCes in an the aOertures in the main MLI enUironment Vhere 10 seconC CeKay can cause oUerheating • "aters Eor  OhysicaK conj gurations Thermal Environment Boundary Conditions due to Mercury’s proximity to the Sun: Data Management • Solar intensity varying between 6,300 W/m2 and 14,500 W/m2 • %"$ %aiKure "ontroK $Kectronics ensures continueC O"S (> 10 solar constants), plus IR >5,200 W/m2 saEe oOeration Curing reboot oE main comOuter • %irst sOacecraEt Vith netVork aOOKication oE SOaceVire interEaces Eor science Cata

Electric Propulsion System DTC MPO to MTM • 4 W 14 m- T ion thrusters oOerateC singKy or in Oairs MPO Mercury Orbit Phase IMH attachment • -ominaK mission 1 $arth year l 2024 • $WtenCeC mission 1 $arth year Eoreseen Launch 4120 kg MPO mass on orbit 1240 kg

Thermal Design for the severe thermal environment Thermal 5erik cation Programme

The MPO unCergoes a k oUer manoeuUre tVice Oer Mercury MOSIF – including MPO l ight model year in orCer to OroUiCe a singKe raCiator surEace Eor heat reIection MLI comOrises a singKe -eWteK outer Kayer Vith  CimOKeC titanium then Kayers seOarateC by gKass sOacers %reeKy suOOorteC oUer Kengths The LSS Large SOace SimuKator at $ST$" is being useC to test anC • 'eatOiOes are embeCCeC in the ePuiOment mounting OaneKs anC oE uO to 2  m UeriEy each moCuKe the raCiator OaneK to transEer anC Cistribute heat • The LSS Vas moCij eC anC has suOOorteC !eOi"oKombo since • LouUres in Eront oE the raCiator rek ect the OKanet inErareC raCiation MTM SeOtember 2010 Vhen the MMO thermaK moCeK Vas testeC VhiKst aKKoVing the raCiator a UieV to sOace contains embeCCeC anC surEace heatOiOes anC uses MLI CeriUeC n intensity oE  soKar constants is achieUabKe at Š2  m • The entire MPO boCy is coUereC Vith high temOerature ML( Erom the MPO Cesign • The MOSI% MMO Vere successEuKKy testeC thereaEter CeUeKoOeC Eor !eOi"oKombo in orCer to combat temOerature anC • The MPO STM EoKKoVeC a year Kater shoVing notabKe CeUiations Erom restrict heat inOut into the boCy the OreCicteC OerEormance reUieV oE the CetaiKeC Cesign anC the Outer heat shieKC comOrises 2 Kayers oE -eWteK ceramic cKoth MLI construction Vas OerEormeC EoKKoVeC by 11 aKuminium Kayers 2 aKuminiseC 4OiKeW Kayers • In autumn 2012 a Karge scaKe test 2 W  m samOKes Uerij eC the anC 10 aKuminiseC MyKar Kayers in  Oackets comOKete imOroUeC MLI Cesign Eor the MPO k ight moCeK the MLI • The MTM STM Vas successEuKKy testeC in sOring 201 its P%M test The -eWteK Kayers reach 0¦" is stiKK to come • The EuKKy ePuiOOeC MPO P%M Vas testeC in -oUember 2014 This conj rmeC the imOroUeC OerEormance oE the thermaK Cesign anC aKso the correct Eunctioning oE the eKectricaK systems oUer the mission temOerature ranges MPO high temperature MLI k Wation 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 Conference, The Woodlands, 2015

The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification. Roger J. Wilson1 and Markus Schelkle1, 1 , Friedrichshafen, ([email protected])

Abstract The spacecraft and subsystem designs are strong- BepiColombo is an interdisciplinary mission per- ly driven by the severe demands of the thermal envi- formed in a partnership between ESA (European ronment experienced at Mercury (whilst the same Space Agency) and JAXA ( Aerospace Explo- solar intensities are also experienced during cruise) ration Agency). The mission aims to place 2 space- and by the staging necessary to evolve the mechani- craft in complementary orbits around Mercury and cal/electrical configuration from the 4-module com- perform scientific investigations of the planet. The posite at launch to the two free-flying orbiters. BepiColombo scientific results will add to the Equipment not required in Mercury orbit is ejected knowledge already gained from the fly- before capture into orbit around the planet. Due to the bys and the Messenger in-situ measurements. of the MPO mission (in a low, inertial, polar JAXA provides the MMO (Mercury Magneto- orbit) all external equipments are of bespoke high- spheric Orbiter), whilst Airbus Defence and Space is temperature design. prime contractor for ESA, providing the MPO (Mer- cury Planetary Orbiter) and all other spacecraft hard- ware. The scientific payload is provided by national agencies. This paper provides an overview of the mission (including its trajectory to Mercury), spacecraft de- sign (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 experi- enced during the cruise phase and in Mercury orbit. Introduction Mercury is the innermost planet of the solar sys- tem, is therefore difficult to reach and until now has been visited by only two spacecraft. Mariner 10 flew- by 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 Euro- pean mission to Mercury. It was named after the Italian Giuseppe “Bepi” Colombo (1929 – Figure 1: The Mercury Composite Spacecraft 1984), who proposed the trajectory of Mariner 10, Whilst addressing the above topics in more detail, and discovered the planet’s 3:2 spin-orbit resonance. this paper also pays particular attention to the verifi- The BepiColombo mission will place the Europe- cation of the thermal control performance. In total, 8 an MPO (Mercury Planetary Orbiter) and the Japa- module sized tests will be performed (starting in nese MMO (Mercury Magnetospheric Orbiter) in low autumn 2010 with the MMO under JAXA responsi- polar orbits around Mercury. The orbiters will be bility) at up to 8 solar constants in the LSS (Large delivered to Mercury by means of a combined launch Space Simulator) at ESTEC. aboard an ECA and with the assistance of dedicated propulsion and protection modules during Overview of the Scientific Objectives the 7 year cruise phase (transfer to Mercury). The The BepiColombo mission will perform scientific cruise phase includes multiple planetary gravity as- investigations [ 1 ] of Mercury in the areas of: sists in addition to the braking provided by the space-  Origin and evolution craft.  Interior, structure, geology, composition  composition and dynamics

2 The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification 46th Lunar and Planetary Science Conference, The Woodlands, March 2015

structure and dynamics recorded the major drivers to the spacecraft design,  Origin of Mercury's including an indication of the direct consequences  Test of Einstein's theory of Two spacecraft in single launch

The MPO and MMO shall be launched in a single The payloads of the MPO and MMO are provided flight configuration – whereby the MMO is passive by national agencies, with the objectives and leader- during the cruise to Mercury and (as a normally spin- ships as per Table 1. ning spacecraft) will require thermal protection dur- The nadir pointed attitude of the MPO is aimed at ing the 3-axis stabilised cruise. This requirement providing continuous viewing of Mercury for the alone leads to some form of stacked configuration. payload over the complete orbit. The only restrictions to payload operation result from temperature and Deceleration to reach Mercury power availability constraints around perihelion. The Mercury is the innermost planet of the solar sys- mission thus enables visibility of the complete Mer- tem and a braking ∆V of ca. 7 km/s is required to cury surface. arrive at Mercury with a suitable for capture into orbit. Planetary manoeuvres can be Table 1: MPO and MMO Payloads used to achieve this deceleration but then time and the number of fly-bys are both large. For Bepi- MPO Colombo it was decided to implement an Electric BepiColombo Laser Altimeter Co-PI: N. Thomas, CH Propulsion System using 2 x 145 mN ion thrusters to Co-PI: T. Spohn, D reduce the cruise phase duration. Since operating Italian Spring PI: V. Iafolla, I these 2 thrusters requires 10 kW electrical power a Magnetic Field Investigation PI: K.H. Glassmeier dedicated propulsion module was conceived for the Mercury Radiometer and PI: H. Hiesinger, D cruise phase. This module then serves as the first Thermal Imaging Spectrome- stage of the flight composite – to which were added ter all functions not required in Mercury orbit. Mercury Gamma-Ray and PI: I. Mitrofanov, RUS Neutron Spectrometer Free Gravity Capture Mercury Imaging X-ray Spec- PI: E. Bunce, UK A free gravity capture, using the Mercury – Sun trometer point, was implemented to enable a safe Mercury Orbiter Radio Sci- PI: L. Iess, I capture without the need for time-critical operations. ence Experiment The approach navigation must ensure suitable trajec- Probing of Hermean Exo- PI: E. Quémerais, F tory and velocity to guarantee the capture. The initial- sphere by UV Spectroscopy ly weak capture must be consolidated by apoherm Search for Exospheric Refill- PI: S. Orsini, I lowering, but this is also a not time critical activity. ing and Emitted Natural Operational Orbits around Mercury Abundances The MPO will be in an inertially fixed polar orbit Solar Intensity X-ray and PI: J. Huovelin, FIN of 0° inclination with periherm distance 480 km and particle Spectrometer apoherm distance 1500 km. The orbit is orientated to Spectrometers and Imagers PI: E. Flamini, I place the apoherm towards the sun at perihelion, for MPO BepiColombo Inte- thereby crossing the far from the planet’s grated Observatory sub-solar point which reaches over 400°C. The MPO MMO has an orbit period of 2.3 hours. Mercury PI: W. Baumjohann, A The MMO will be placed in a coplanar orbit with Mercury Particle Ex- PI: Y. Saito, JPN the MPO, with periherm distance 590 km and periment apoherm distance 11,640 km. The MMO has an orbit Plasma Wave Instrument PI: Y. Kasaba, JPN period of 9.3 hours. Mercury Atmospher- PI: I. , JPN MPO Observation Concept and Attitude ic Spectral Imager The MPO will fly permanently nadir pointed to Mercury Dust Monitor PI: H. Shibata, JPN provide continuous scientific observation. Thermal Environment Boundary Conditions Mercury orbits the sun with perihelion distance Mission and Spacecraft Driving Requirements 0.31 AU and aphelion distance 0.47 AU. Resulting The spacecraft design is greatly influenced by the mission and the environment of Mercury. Below are

2 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 BepiColombo the maximum one-way signal time is 6,300 W/m2 and 14,500 W/m2 (> 10 solar constants). 14.4 minutes. The Mercury sub-solar point reaches >400°C at Mercury is regularly behind the sun, as is the perihelion – which results in an intensity of spacecraft during the cruise phase: these solar con- 5,200 W/m2 on the MPO. The choice of orbit results junctions inhibit communication between spacecraft in an infrared intensity of 5,400 W/m2 on the MPO and ground. The maximum during cruise when Mercury is at aphelion. is 20 days and 7 days when in MPO orbit, during The inertially fixed orbit entails that, over a Mer- which the spacecraft must operate without assistance cury year, an orbiting body experiences sun illumina- from ground. tion from all directions. At any time, the MPO re- In addition to this long-term autonomy require- quires at least 1 surface for heat rejection. A single ment, the avionics must ensure that a safe attitude is radiator surface is made available for thermal control maintained, thereby avoiding that damaging tempera- by means of flip-over manoeuvres which rotate the tures are generated on external surfaces. Critical spacecraft 180° around the nadir-pointed axis at each durations are in the order of only a few 10s of sec- perihelion and aphelion (thereby inverting the flight onds. direction). System Breakdown and Functional Appor- tionment From the Mission and Spacecraft Driving Re- quirements 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 com- posite of x modules, hence the various configurations are known as MCSn for Mercury Composite Space- craft (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 contain- ing all the capabilities needed to perform its scientific mission – for which careful optimisation was neces- sary when considering the thermal environment. The MPO therefore contains most of the capabilities also needed during cruise. In order not to compromise the Figure 2: The MPO orbit around Mercury MPO design by taking unnecessary hardware into The thermal environment of the MPO orbit neces- Mercury orbit, hardware needed solely for cruise is sitates that all external items of the MPO must be accommodated in a separate MTM (Mercury Transfer capable of withstanding high solar intensity and high Module). , whilst in addition the MLI (multi-layer The MMO is eventually also a free-flying space- insulation) must restrict the heat input to the space- craft containing all the capabilities needed to perform craft body to manageable levels. its scientific mission. However with the spacecraft Since the cruise trajectory entails repeatedly visit- capabilities controlled from the MPO during cruise, ing the perihelion distance of Mercury from the sun the MMO then remains passive throughout (apart (whilst progressively reducing the aphelion distance) from periodic check-outs). Since the MMO is a nor- the MTM and MOSIF are also subjected to > 10 solar mally spinning spacecraft, it requires thermal protec- th constants. As the MCS flys in a sun pointed attitude tion during the 3-axis stabilised cruise. The 4 mod- the constraints are less severe than for the MPO, ule of the MCS derives from the MMO’s needs: the nevertheless the sunward face and the solar arrays MOSIF (MMO SunShield and InterFace Structure) must withstand high solar intensity. providing thermal protection as well as all the inter- Operations and Control Requirements faces between MPO and MMO. As is typical with inter-planetary missions, com- The MCSL and MCSC are composed of: munication and control are constrained by the one-  MPO (Mercury Planetary Orbiter) way signal time between spacecraft and which • Spacecraft optimised for its operational mis- inhibits real-time commanding from ground. For sion

3 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 The MCSA is created upon MTM separation. On (with only minor hardware modification for reaching the MMO orbit, the MMO is released to MCS configurations, notably the size of the create the MCSO. The MOSIF is ejected shortly reaction wheels to control the MCS) afterwards to leave the MPO. • Chemical Propulsion System not used dur- Spacecraft Configuration ing cruise, however after MTM separation the MPO performs approach propulsion and MPO Configuration apoherm lowering of Mercury orbit The MPO was conceived and optimised to meet  MMO (Mercury Magnetospheric Orbiter) the needs of its operational orbit, before undergoing • Spins during operational mission minor adaptations to fit within the MCS. • Is passive during cruise – apart from check- The spacecraft envelope is compact in order to outs minimise the heat input through the MLI, and hence  MOSIF (MMO SunShield and InterFace the heat load to the single radiator. The dedicated Structure) radiator is sized for a maximum heat rejection (at • Thermal protection for the MMO maximum temperature) of 1500 W – where 20% of • Mechanical interface for the MMO this quantity is from parasitic heat leaks into the • Harness routing between MPO and MMO spacecraft. The radiator locally fills the Ariane fair-  MTM (Mercury Transfer Module) ing. Heatpipes transport heat from the equipment • Provides MEPS (MTM Electric Propulsion panels to the radiator and radiator heatpipes provide System) plus chemical propulsion (for cruise heat distribution. The equipment mounting and the AOCS and navigation correction) heatpipe network are concentrated on 2 • Provides power for electric propulsion sys- sandwich panels which constitute the major elements tem and for MPO +MMO of the central double-H structure (completed by 2 • Separated before capture into Mercury orbit 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: MPO – showing equipment panel per- pendicular 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 Figure 3: Composition of MCS, showing module communication links in the MPO orbit when the functions and contributions to MCS Earth can be located in any direction with respect to

4 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 accommodat- The MMO in stowed configuration can be ac- ed to provide viewing capability in MPO orbit and commodated in a cylinder Ø 1800 x 1200 mm. during cruise. The MGAMA is mounted on a long MOSIF Configuration boom which can rotate about the boom axis as well The MOSIF acronym (MMO SunShield and In- as about the boom attachment point – this 2-axis terFace Structure) very much describes its functional- provides the capability to view around the ity, which then leads to its configuration. obstruction of the MPO body and also past the ob- The MOSIF consists of two major assemblies: structions generated by the MTM and MOSIF in the  The Adapter has at its centre the circular inter- MCS configuration. The HGAMA has 2-axis rotation face to the MMO, the arms provide the at- capability providing it with a large field of view tachments for the Sunshield, the arms provide around the MPO body, whilst having a restricted the 4 separable attachments to the MPO view in the MCS configuration.  The Sunshield structure supports the MLI which shades the MMO. The Sunshield is truncated at 18° to allow MCS/MOSIF rota- tion towards the sun without illuminating the MMO. The Sunshield is conical at 16° half- angle to tolerate wobble of the MMO during its slowly spinning separation

Figure 5: MPO deployed configuration

The MPO is 1.7 m high by 3.6 m across the radia- tor 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 an- tennas and deployable instrument booms. The circu- lar high-gain antenna is deployed before separation from the MCSO, the other deployments are per- formed post-separation. The MMO provides a circu- lar, bolted interface to the MOSIF. The Spin Ejection Figure 7: MOSIF – with MMO installed Device provided by JAXA remains bolted to the MOSIF. MTM Configuration The configuration of the MTM is mainly driven by the MEPS. The power demand of 5 kW per thrust- er (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 pan- els are needed which are themselves in a 20° wedge configuration, allowing the MCS to rotate slightly about its longitudinal axis whilst avoiding sun illumi- nation 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 rota- tion of the thrust vector towards the sun without the Photo courtesy of JAXA thrusters being illuminated. Figure 6: MMO flight model

5 The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification 46th Lunar and Planetary Science Conference, The Woodlands, March 2015

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 Figure 8: MTM deployed configuration launcher is required to deliver the needed . MCS Configuration The MCS configuration is represented by a stack- ing of the above described modules.

Figure 9: Cross-section through the MCS Figure 10: From launch to MPO orbit showing gravity assists and separations Table 2: Overall Mass Budget A deceleration of 7 km/sec is required to reach MMO MOSIF MPO MCSA MTM MCSL Mass Budget Mercury. 4.2 km/sec will be provided by the MEPS – [kg] [kg] [kg] [kg] [kg] [kg] enabling the spacecraft to more rapidly synchronise Dry Mass 285 127 1169 1580 1134 2714 with each subsequent planet gravity assist. The re- Propellant mainder of the braking will be achieved by gravity MOI CPS 602 602 602 assists at Earth, and Mercury. The cruise until MPO CPS 67 67 67 MTM separation in November 2023 takes 6 years 10 Cruise CPS n/a 157 157 months, includes 17 orbits around the sun and 8 grav- ity assist manoeuvres: 1 x EGA, 2 x VGA and Xenon n/a 581 581 5 x MGA. During the electric propulsion thrust phas- Total Mass 285 127 1838 2249 1872 4121 es the MCS is orientated with the thrusters pointing

6 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 From the many changes of flight configuration, thruster steering supporting the sys- the number of actuators employed and the stringent tem. During coast phases the MCS is stabilised by safe and survival modes the AOCS consists of 17 rotation around the sun vector, thereby minimising operational modes. chemical propellant consumption. Data Management System (DHS) After MTM separation, the MPO takes on the The basic MPO DHS comprises redundant OBCs propulsion tasks of the MCSA for the MOI (Mercury (On-Board Computer) and an internally redundant Orbit Insertion), which begins with navigation to the SSMM (Solid State Mass Memory) for payload and Mercury-Sun Lagrange point. A free gravity capture spacecraft data storage. A MIL-STD-1553B bus is is employed, avoiding the risks of time-critical ma- used for spacecraft telemetry and telecommand whilst noeuvres prior to Mercury capture. The MCSA is all payload TM/TC and science data interfaces use weakly captured into an initial orbit 178,000 x 670 SpaceWire. BepiColombo is the first spacecraft with km. The MOI includes a series of 15 orbit lowering a network application of SpaceWire interfaces. manoeuvres, the first of which is needed to stabilise The MPO provides all the intelligence during the capture, thereafter the manoeuvres lead to the cruise and is enhanced with additional data buses to respective operational orbits of the MMO and MPO. the MMO and MTM for this purpose. The MPO orbit will be reached in January 2024. Permanent availability of a functioning processor Subsystem Designs and BepiColombo specific to guarantee safe and prompt attitude control is pro- solutions vided in Survival Mode by redundant FCEs (Failure Attitude and Orbit Control System (AOCS) Control Electronics) which take over the control functions in the event of an OBC reboot. The FCEs For MPO science operations the AOCS must pro- retain control for 7 minutes after which it is taken vide continuous nadir pointing whilst meeting accu- over by the reconfigured OBC. racy and stability requirements. Two Star Trackers (plus a 3rd for redundancy) and an Inertial measure- Communications System ment Unit (IMU) are co-mounted with instruments on The communications system [ 2 ] is equipped an optical bench whilst 4 reaction wheels serve as with 2 x LGA (Low Gain Antenna), a 2-axis steerable actuators (with 5 N thrusters used for wheel off- MGA (Medium Gain Antenna) and a 2-axis steerable loading). The AOCS also controls the thermally criti- HGA (High Gain Antenna). The LGAs provide cal orientation of the solar array and the 22 N thrust- spherical coverage, can support X-band downlink and ers for orbit manoeuvres during Mercury Orbit Inser- uplink near Earth, whilst also providing uplink recep- tion. tion over greater distances. The X-band horn MGA is This basic AOCS is enhanced with sun sensors steerable around the MPO or MCS obstructions in for survival mode and is further enhanced for the order to view Earth and is the primary antenna during MCS configuration when the MEPS thrusters and cruise. The Ø 1100 mm HGA is steerable around the MTM 10 N thrusters serve as actuators. As for the MPO and is the primary antenna during science oper- MPO, MTM Solar Arrays are also thermally con- ations when it supports up- and downlinks in both X- trolled by the appropriate orientation. and Ka-band. During cruise the AOCS controls the MEPS The newly developed DST (Deep Space Tran- thruster orientation and corresponding MCS attitude sponder) supports telecommanding uplink in X-band as required by the uploaded mission timeline – with with telemetry downlink in both X- and Ka-bands to fine pointing of the MEPS thrusters minimising mo- enable the downlink of 1550 Gb/year of science data. mentum accumulation by the reaction wheels. The DST supports ranging in X/X band and X/Ka The thermal environment experienced in the band whilst Ka/Ka band ranging is provided with the MPO orbit and during cruise allows (for a number of inclusion of the payload-provided MORE translator. thermally critical items) only deviations from nomi- This ranging strategy is related to the Radio Science nal attitudes in the order of seconds before overheat- Experiment and requires high stability of the HGA. ing and damage occurs. In the event of an OBC re- Power amplification is by TWTAs for both X- boot, the Survival Mode will be entered and the and Ka-bands. All antennas are exposed to the severe AOCS control will be transferred to the FCE and a thermal environment and are based on titanium. The second IMU. In Survival Mode the AOCS uses Sun antenna pointing mechanisms for HGA and MGA are Sensors as the attitude reference . Different Survival capable of operating at 250°C. attitudes apply for the various spacecraft configura- tions.

7 The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification 46th Lunar and Planetary Science Conference, The Woodlands, March 2015

MTM Electric Propulsion System (MEPS) tilts the array to a maximum of 75° from the sun. The The MEPS contains the 4 electric propulsion OSRs occupy 17% of the panel area. The tilted, sin- (Kaufmann) thrusters, their power processing elec- gle wing configuration is needed to achieve sun illu- tronics, 4 thruster pointing mechanisms and the xen- mination whilst permanently nadir pointing through- on storage and feed system. out the Mercury year – this entails near continuous The thrusters are the QinetiQ T6, Ø 22 cm, rotation controlled by the AOCS. 145 mN thruster derived from the T5 flown on the During the cruise all power is provided by the GOCE mission. The system is planned to operate MTM solar array. The MTM provides a 100 V regu- over 25 thrust arcs totalling 880 days, with the long- lated bus for the electric propulsion system and a est continuous operation being for 167 days. The 28 V regulated bus for other equipment plus the system typically operates using two thrusters. MPO/MMO. During gravity assist manoeuvres, when in some cases 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 MEPS. The MTM solar array is a high temperature de- sign 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° Photo courtesy of QinetiQ from the sun. The size of the array is driven by the need to provide maximum power also at 0.31 AU. Figure 11: T6 thruster firing test Whilst approaching the sun the solar array output The thrusters are mounted on individual pointing initially increases, of course accompanied by an in- mechanisms which enable the thrust vector to point crease of temperature. Once the temperature has through the MCS centre of mass – either for a single reached 190°C (at about 0.5 AU) the array must be thruster or the plane of an operating pair of thrusters. tilted, thereby reducing its projected area and limiting By off-pointing of the thrust vector a moment is cre- its output. The two wings total 40 m2 and have a ated for use by the AOCS for wheel off-loading. mass of 290 kg. The xenon system, with its flow control units, Chemical Propulsion Systems (CPS) pressure regulators and 3 tanks, is able to store and The MPO CPS is tasked with the 15 MOI ma- deliver 580 kg of xenon which can provide 5400 m/s noeuvres and attitude control, for which it is delta V. equipped with redundant 4 x 22 N and redundant Power 4 x 5 N thrusters. The 22 N thrusters are bi-propellant The MPO uses a 28 V regulated power bus [ 3 ]. whilst the 5 N thrusters are mono-propellant: these After MTM separation the power is provided by the are combined into the first dual-mode propulsion MPO solar array which is kept edge-on to the sun system implemented on a European spacecraft. The during cruise to minimise UV degradation. An 96 Ah system uses hydrazine and MON (Mixed Oxides of lithium-ion battery provides power during eclipses. ). 669 kg of propellant are carried, giving a The solar array provides TBD W which is sufficient capability of 1000 m/s delta V plus attitude control. for full payload operation and alternatively sufficient The MTM CPS employs redundant 12 x 10 N bi- to provide 300 W to the MMO until its separation. propellant thrusters, using MMH (Mono-Methyl The MPO solar array is a high temperature design Hydrazine) and MON (Mixed Oxides of Nitrogen). operating at maximum 190°C – for which compo- This system is derived from Eurostar 3000 systems [ nents have required dedicated development. The 4 ]. As well as attitude control capability, the MTM thermal control of the array is achieved by its design, CPS can provide axial thrust for cruise navigation. involving a mix of cells and OSRs (optical solar 157 kg of propellant are carried, giving a capability reflector) enhanced by in-plane thermally conductive of 68 m/s delta V plus attitude control. CFRP facesheets, and control by the AOCS which

8 The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification 46th Lunar and Planetary Science Conference, The Woodlands, March 2015

Thermal Control  7 dimpled titanium layers separated by glass The MPO TCS must regulate the equipment tem- spacers peratures (achieving standard equipment levels) [ 5 ],  It is freely supported over lengths of up to transfer heat to the single radiator, shield the radiator 2.5 m and must withstand the vibration and from planet infrared illumination, reject 1200 W of acoustic environments of the launch dissipated heat from the payload and spacecraft equipments and reject up to 300 W of parasitic heat The MTM TCS must regulate the equipment tem- which enters the MPO body. These functions are peratures, distribute heat in the radiators, reject achieved by means of: 2000 W of dissipated heat equipments and reject up  Heatpipes embedded in the equipment mount- to 300 W of parasitic heat which enters the MTM ing panels to collect and transfer the heat the body. These functions are achieved by means of: radiator panel  Heatpipes embedded in the radiator panels  Spreader heatpipes in the radiator panel, ther- (which also serve for equipment mounting) mally connected to the equipment panels by  The embedded heatpipe network is enhanced 90° linking heatpipes by surface heatpipes  97 heatpipes are used, of which a few are 3-  63 heatpipes are used dimensional hence difficult to test on ground  Derivatives of the high-temperature MLI are  Fixed louvres are mounted in front of the radi- used ator to reflect the planet infrared radiation away from the radiator whilst allowing the ra- Further MLI applications result from the stack diator an extensive view to space configuration and the separation interfaces of the  The entire MPO body is covered with high MPO: temperature MLI developed for BepiColombo  Whilst the modules are protected as described  The outer heat shield comprises 2 layers of above, solar illumination gaps between mod- Nextel ceramic cloth followed by 11 alumini- ules can not be tolerated um layers. The Nextel layers reach 380°C  Elaborate Gap Closure MLI is implemented  Moving inwards to lower temperature, 26 lay- between MTM-MPO and MPO-MOSIF. This ers of aluminised Upilex are followed by 10 MLI is in contact with the MPO and is at- layers of aluminised Mylar tached to the separating modules (see Figure  Spacers of glass fibre and AAerofoam are 1) used to separate the layers in the 4 packets,  The 4-point mechanical interfaces between whilst rosettes separate the packets modules leave holes of Ø140 mm in MLI (to  The installed MLI has a thickness of 65 mm. these add 2 x Ø170 mm holes for the connect- The total MLI mass is 94 kg. ors at each interface). These holes are closed by DTCs (Deployable Thermal Covers) con- taining 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 heatshield.

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: Figure 13: Deployable Thermal Cover to close  A single Nextel outer layer separation apertures

9 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 im- plemented. The structure uses aluminium sandwich throughout – with the exception of the CFRP sand- wich Optical Bench. The configuration can be visual- ised from Figure 4.

The MOSIF structure is characterised by:  The Adapter: a central cylindrical piece inter- facing to the MMO, supported on box con- struction arms from the IMH interfaces to the Figure 14: MTM-MPO IMH element MPO – the arms extend further to support the Spacecraft Operations and Autonomy Sunshield, both parts use aluminium  The Sunshield: an aluminium truss structure Operations will generally be performed off-line, of circular section tubes to support the MLI using an on-board master schedule called the Mission Timeline (MTL). The MTL is enhanced by On-Board The MTM structure provides a circular launcher Control Procedures (OBCPs), file transfer capability, interface to the Payload Adapter System 1666MVS, high level command functionality, and a flexible on- from which a CFRP sandwich main cone extends board monitoring function. upwards to the 4-point interface to the MPO. The The autonomous functions on-board are tasked xenon and CPS tanks are mounted off the main cone, with the execution of the MTL whilst ensuring that whilst CFRP sandwich floors and shear panels ac- each item is in a safe and correct configuration by commodate smaller items. The floors and shear pan- means of monitoring and FDIR (Failure Detection, els support the aluminium radiators (carrying the Isolation and Recovery). The hierarchical FDIR ena- power and MEPS electronics) and the solar arrays. bles isolation of unambiguously identified anomalies The thruster floor, within the main cone, carries the by local reconfiguration. For other anomalies the MEPS thrusters and pointing mechanisms. safety of the spacecraft is guaranteed by entry in The MTM, MPO and MOSIF structures provide Survival then . Return to Normal Mode is the core of the MCS structure at launch. These ele- by ground command. The possible durations in ments (MTM and MPO structures are separated by Safe/Survival modes are driven by the solar conjunc- 130 mm, with a smaller gap to the MOSIF) are con- tions. By this means, for example, it will be possible nected by Inter Module Hardware (IMH) completing to continue MEPS operation during solar conjunc- the MCS structure at launch and providing the sepa- tions – for which back-up thruster configurations and ration functions after cruise. The IMH also includes corresponding attitudes are stored on-board. attached hardware and separable connectors for the In the post-launch, near-Earth phase ground con- inter-module electrical connections (excluding to tact will be provided for up to 24 hours/. During MMO). To provide the structural performance over cruise the contacts will be reduced to only 8 the module-module distances, along with the separa- hours/week – except around planet fly-bys. In MPO tion functions and dynamics (both electrical and orbit a daily 10 hour pass is planned, of which 9 mechanical), the IMH requires a mass of 47 kg. hours are allocated for data downlinking. ESA’s Cebreros Ø35 m ground station is the pri- mary station for all mission phases. Overview of the Integration and Verification Programme 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

10 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 exten- mance. The high temperatures experienced warranted sive AIV programme was undertaken. a review of the thermal design implementation and in An ATB (Avionics Test Bench) was employed to particular reconsideration, and verification, of the test and verify the avionics and software, using real MLI design. 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 configura- tion. This enabled development and testing with spacecraft hardware and instrument models and was extended to electrically represent the MCS configura- tion with MTM and MMO representations. This programme enabled extensive testing and risk reduc- tion before embarking on the PFM programme. An STM (Structural Thermal Model) was em- ployed to demonstrate the functioning and perfor- mance of the all-new structures and thermal control systems (using dummy electronic units). The se- quence 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 Figure 15: MPO PFM in the LSS demonstrate adequate workmanship. The modules The flight model MLI design includes an extra were transferred to ESTEC in July 2014 for the con- Nextel layer, more layers in total and a thermally tinuation of assembly and the verification pro- improved attachment system. This design was suc- gramme. cessfully verified in a large scale back-to-back com- Thermal Verification parison against the STM design (2 x 3 m samples of Each module is of a completely new mechanical MLI tested in the LSS) in October 2012. and thermal design and employs the newly developed The MTM STM was tested in spring 2013 and the high temperature MLI. The thermal environment to performance fitted satisfactorily to the predictions. be experienced is severe, when compared with near- A thermal balance / thermal vacuum test (TB/TV) Earth or the colder environments experienced at the of the fully equipped MPO PFM was performed in outer . The need to demonstrate the thermal November 2014. This validated the performance of performance at solar intensities close to the >10 solar the thermal design modifications, along with the constants to be experienced was considered para- improved MLI, and also the correct functioning of mount. the electrical systems over the mission temperature The LSS (Large Space Simulator) at ESTEC [ 6 ] ranges. The temperatures measured inside the MPO was modified and has supported BepiColombo since were typically 4°C cooler than predicted. September 2010. All 19 sun-simulating lamps are Still to be performed in the LSS are a test of the used at high power and the mirrors of the normally complete high gain antenna and TB/TV test of the parallel Ø6 m beam have been adjusted to conical MTM PFM. such that more than 8.5 solar constants illuminate the Conclusions test object with Ø2.7 m. The mechanical performance of the MCS was The MMO thermal model was the first tested, in demonstrated on the STM, along with the ability to September 2010, and was followed shortly afterwards separate the modules. The thermal performance of the by the MOSIF + MMO combination (Figure 7). MOSIF and MTM were demonstrated on the respec- The MPO STM was tested in September 2011 tive STMs. The avionics and electrical systems have and the value of the test was clearly demonstrated by been developed and demonstrated on their respective the notable deviations from the predicted perfor-

11 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 the programme is well on its way towards departure flow, with the capabilities of the MPO having been for Kourou 5 months before launch in December extensively verified during its TB/TV test. With the 2016. arrival of MMO flight model at ESTEC in April 2015

MPO and MMO in orbit around Mercury

References [ 5 ] Federica Tessarin, Domenico Battaglia and [ 1 ] Johannes Benkhoff, Jan van Casteren, Tiziano Malosti, Daniele Stramaccioni, Jürgen Schil- Hajime Hayakawa, Masaki Fujimoto, Harri Laakso, ke, “Thermal Control Design of Mercury Planetary Mauro Novara, Paolo Ferri, Helen R. Middleton, Orbiter”, AIAA 2010-6090, 40th International Con- Ziethe, “BepiColombo—Comprehensive explo- ference on Environmental Systems ration of Mercury: Mission overview and science [ 6 ] Large Space Simulator (LSS), at ESTEC, goals”, Planetary and Space Science 58 (2010) 2–20 Noordwijk, The Netherlands [ 2 ] M. Mascarello, P. Pablos, R. Heinze, http://www.esa.int/Our_Activities/Space_Engineerin A Busso, R. Carbone, “The BepiColombo X/X/Ka g_Technology/Large_Space_Simulator_LSS TT&C Subsystem”, TTC 2010, 5th ESA Workshop on Telemetry, Tracking & Communications Systems for Space Applications, ESA-ESTEC 21 23 September 2010 [ 3 ] Pierluigi Morsaniga, Giuseppe Gervasio, Giuseppe Cuzzocrea, “BepiColombo Electrical Pow- er System”, Proc. ‘9th European Space Power Con- ference’, Saint Raphaël, , 6–10 June 2011 (ESA SP-690, October 2011) [ 4 ] Eurostar 3000 , Airbus DS http://www.space- airbusds.com/en/programme/eurostar-series-czw.html

12