The BepiColombo Spacecraft, its Mission to Mercury and its Thermal
Transcription
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 Launch • rianD $" into DscaOD orbit Spacecraft Driving Requirements 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 Mercury: • 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 MPO to MTM IMH attachment Launch mass MPO mass on orbit MOSIF 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 2 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 MTM 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 then • '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 DTC MOSIF MMO Sun2hieKC anC InterFace Structure • 3hermaK Orotection Eor the MMO • MechanicaK anC eKectricaK interEaces Eor the MMO Subsystem and Hardware Implications Module Attachment and Separation MPO Mercury Orbit Phase MMO • 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. Abstract 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 agencies. 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. Introduction 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 2 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 MPO 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 Abundances Solar Intensity X-ray and particle Spectrometer Spectrometers and Imagers for MPO BepiColombo Integrated Observatory MMO 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 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 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 direction). 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 Module). 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 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 (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 Structure) • 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 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 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 MOSIF. 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 5 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 velocity. 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 MMO [kg] MOSIF [kg] MPO [kg] MCSA [kg] MTM [kg] MCSL [kg] 285 127 1169 1580 1134 2714 MOI CPS 602 602 MPO CPS 67 67 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 Propellant 602 67 Cruise CPS n/a 157 157 Xenon n/a 581 581 2249 1872 4121 Total Mass 285 127 1838 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 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 solutions 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. 7 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 MEPS. 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. Power 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 8 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 spacers 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 used Further MLI applications result from the stack configuration and the separation interfaces of the MPO: 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 1) 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 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: A single Nextel outer layer Figure 13: Deployable Thermal Cover to close 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 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 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 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 MTM PFM. 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- Conclusions 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 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 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 2016. 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 http://www.esa.int/Our_Activities/Space_Engineerin g_Technology/Large_Space_Simulator_LSS References [ 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 2010 [ 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 http://www.spaceairbusds.com/en/programme/eurostar-series-czw.html 12