The BepiColombo Spacecraft, its Mission to Mercury and its … · 2015. 3. 18. · The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification 46th Lunar and Planetary

Launch mass 4120 kgMPO mass on orbit 1240 kg

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

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

MMO spins when free ying, therefore needs shading during the 3-axis stabilised cruise phase:• A sunshade and interface structure are required

a separate module is needed

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)Operational orbits around Mercury can only be inertially xed

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)

From Launch to Mercury Orbit Functional BreakdownMission Objectives

h main sci nti c ob cti s o th i o ombo mission to M rcur com ris th in stigation o th origin an o ution

o a an t c os to th ar nt star an a com r h nsi stu o th an t its his i b achi b acing t o s ac

cra t in i r nt o ar orbits aroun M rcur

Launch• rian into sca orbit

Interplanetary Cruise Phase

• 1 arth 2 nus an M rcur gra it assist mano u r s• ctric Pro u sion or braking b t n gra it assists short ns trans r• M M s aration on 1 11 202 a t r 1 orbits aroun th sun

Mercury Approach Phase• a igation b or ca tur th n orbit o ring• r gra it ca tur on 01 01 2024• 1000 m s mano u r s r orm b MPO• MMO s aration in MMO orbit• MO s aration• sc nt to MPO orbit

h M Mercury Com osite S acecra t consists o 4 o timise mo u es MPO is o timise or its o erationa mission• Per orms comman an contro or M• Per orms a roach ro u sion an Mercury orbit o ering

MMO• ins uring its o erationa mission• s assi e uring cruise

MOSIF MMO Sun hie an InterFace Structure• herma rotection or the MMO• Mechanica an e ectrica inter aces or the MMO

MTM Mercury Trans er Mo u e• Pro i es braking by means o e ectric ro u sion• Pro i es o era o er source uring cruise• hemica ro u sion or na igation an O S

MPO Mercury Orbit Phase• omina mission 1 arth year 2024

• ten e mission 1 arth year oreseen

MOSIF + MMO – thermal test models in LSS

The MPO un ergoes a i o er manoeu re t ice er Mercury year in or er to ro i e a sing e ra iator sur ace or heat re ection then• eat i es are embe e in the e ui ment mounting ane s an

the ra iator ane to trans er an istribute heat• Lou res in ront o the ra iator re ect the anet in rare ra iation

hi st a o ing the ra iator a ie to s ace• The entire MPO bo y is co ere ith high tem erature ML

e e o e or e i o ombo in or er to combat tem erature an restrict heat in ut into the bo y

Outer heat shie com rises 2 ayers o e te ceramic c oth o o e by 11 a uminium ayers 2 a uminise i e ayers an 10 a uminise My ar ayers in ackets com ete the MLI

The e te ayers reach 0

MOSIF MLI com rises a sing e e te outer ayer ith im e titanium ayers se arate by g ass s acers ree y su orte o er engths o u to 2 m

MTM contains embe e an sur ace heat i es an uses MLI eri e rom the MPO esign

Subsystem and Hardware Implications

The ri ing re uirements ha e im acts beyon the mechanica an therma systems

Communications System• an a ban s or 1 0 bit yr science ata an

a a a ranging• ntennas o titanium to sur i e therma en ironment

Power System• So ar rrays o erating at 1 0 Tem erature imite by ti ting

ane s un er O S contro to rom sun• 1 OS s on MPO So ar rray

AOCS (AttitudeControl)• ontro o s acecra t an so ar array attitu es in an

en ironment here 10 secon e ay can cause o erheating• aters or hysica con gurations

Data Management• ai ure ontro ectronics ensures continue O S

sa e o eration uring reboot o main com uter• irst s acecra t ith net ork a ication o S ace ire

inter aces or science ata

Electric Propulsion System• 4 14 m T ion thrusters o erate sing y or in airs

Thermal eri cation Programme– including MPO ight model

The LSS Large S ace Simu ator at ST is being use to test an eri y each mo u e

• The LSS as mo i e an has su orte e i o ombo since Se tember 2010 hen the MMO therma mo e as teste

n intensity o so ar constants is achie ab e at 2 m• The MOSI MMO ere success u y teste therea ter• The MPO STM o o e a year ater sho ing notab e e iations rom

the re icte er ormance re ie o the etai e esign an the MLI construction as er orme

• In autumn 2012 a arge sca e test 2 m sam es eri e the im ro e MLI esign or the MPO ight mo e

• The MTM STM as success u y teste in s ring 201 its P M test is sti to come

• The u y e ui e MPO P M as teste in o ember 2014 This con rme the im ro e er ormance o the therma esign an a so the correct unctioning o the e ectrica systems o er the mission tem erature ranges

The uro ean MPO Mercury P anetary Orbiter carries11 instruments some ith mu ti e sensors

y courtesy o

The a anese MMO MercuryMagnetos heric Orbiter carries

instruments ith a tota o 11 sensors

4 th Lunar an P anetary Science on erence The oo an s T

Module Attachment and Separation

M S con guration an se arations a ter years o cruise• The centra structure o the M S at aunch is com ose o mo

u e structures oine by inter mo u e e ements IM InterMo u e Har are inc 2 2 e ectrica connections

• 4 oint attachments at MTM MPO an MPO MOSI inter aces em oye to enab e minimisa tion o arasitic heat in ut once in Mercury orbit

The IM e ements ass through the T De oyab e Therma Co er rames to connect the structures ter se aration the MLI iscs o the T s are e oye to therma y c ose the a ertures in the main MLI

MPO to MTM IMH attachment

DTC

Spacecraft Driving Requirements

Thermal Design for the severe thermal environment

MPO high temperature MLI ation MPO PFM - thermal testing in LSS

e i o ombo is a oint S mission to ace 2 s acecra t in orbits aroun MercuryAirbus DS is prime contractor for the European industrial part

The BepiColombo Spacecraft, its Mission to Mercury and its Thermal eri cation

oger i son an Markus Sche k e irbus S mb rie richsha en ermany

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

2

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])

Abstract BepiColombo is an interdisciplinary mission per-

formed in a partnership between ESA (European Space Agency) and JAXA (Japan Aerospace Explo-ration Agency). The mission aims to place 2 space-craft 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 fly-bys and the Messenger in-situ measurements.

JAXA provides the MMO (Mercury Magneto-spheric Orbiter), whilst Airbus Defence and Space is 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 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 Europe-an MPO (Mercury Planetary Orbiter) and the Japa-nese 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 as-sists in addition to the braking provided by the space-craft.

The spacecraft and subsystem designs are strong-ly driven by the severe demands of the thermal envi-ronment experienced at Mercury (whilst the same solar intensities are also experienced during cruise) and by the staging necessary to evolve the mechani-cal/electrical configuration from the 4-module com-posite 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 high-temperature design.

Figure 1: The Mercury Composite Spacecraft

Whilst addressing the above topics in more detail, this paper also pays particular attention to the verifi-cation of the thermal control performance. In total, 8 module sized tests will be performed (starting in autumn 2010 with the MMO under JAXA responsi-bility) 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

mailto:[email protected]


2

Magnetosphere structure and dynamics Origin of Mercury's magnetic field Test of Einstein's theory of general relativity

The payloads of the MPO and MMO are provided

by national agencies, with the objectives and leader-ships 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 Mer-cury surface.

Table 1: MPO and MMO Payloads MPO

BepiColombo Laser Altimeter Co-PI: N. Thomas, CH Co-PI: T. Spohn, D

Italian Spring Accelerometer PI: V. Iafolla, I Magnetic Field Investigation PI: K.H. Glassmeier Mercury Radiometer and Thermal Imaging Spectrome-ter

PI: H. Hiesinger, D

Mercury Gamma-Ray and Neutron Spectrometer

PI: I. Mitrofanov, RUS

Mercury Imaging X-ray Spec-trometer

PI: E. Bunce, UK

Mercury Orbiter Radio Sci-ence Experiment

PI: L. Iess, I

Probing of Hermean Exo-sphere by UV Spectroscopy

PI: E. Quémerais, F

Search for Exospheric Refill-ing and Emitted Natural Abundances

PI: S. Orsini, I

Solar Intensity X-ray and particle Spectrometer

PI: J. Huovelin, FIN

Spectrometers and Imagers for MPO BepiColombo Inte-grated Observatory

PI: E. Flamini, I

MMO Mercury Magnetometer PI: W. Baumjohann, A Mercury Plasma Particle Ex-periment

PI: Y. Saito, JPN

Plasma Wave Instrument PI: Y. Kasaba, JPN Mercury Sodium Atmospher-ic Spectral Imager

PI: I. Yoshikawa, JPN

Mercury Dust Monitor 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

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 spin-ning spacecraft) will require thermal protection dur-ing the 3-axis stabilised cruise. This requirement alone leads to some form of stacked configuration.

Deceleration to reach Mercury Mercury is the innermost planet of the solar sys-

tem 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 Bepi-Colombo 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.

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 trajec-tory and velocity to guarantee the capture. The initial-ly weak capture must be consolidated by apoherm lowering, but this is also a not time critical activity.

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.

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


3

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 Mer-cury year, an orbiting body experiences sun illumina-tion from all directions. At any time, the MPO re-quires 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: The MPO orbit around Mercury

The thermal environment of the MPO orbit neces-sitates 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 space-craft body to manageable levels.

Since the cruise trajectory entails repeatedly visit-ing 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, com-

munication and control are constrained by the one-way signal time between spacecraft and Earth which inhibits real-time commanding from ground. For

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 con-junctions 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 require-ment, the avionics must ensure that a safe attitude is maintained, thereby avoiding that damaging tempera-tures are generated on external surfaces. Critical durations are in the order of only a few 10s of sec-onds.

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 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 space-craft 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 nor-mally spinning spacecraft, it requires thermal protec-tion during the 3-axis stabilised cruise. The 4th mod-ule of the MCS derives from the MMO’s needs: the MOSIF (MMO SunShield and InterFace Structure) providing thermal protection as well as all the inter-faces between MPO and MMO.

The MCSL and MCSC are composed of: MPO (Mercury Planetary Orbiter)

• Spacecraft optimised for its operational mis-sion


4

• 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 dur-ing 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 check-

outs 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 sys-tem and for MPO +MMO

• Separated before capture into Mercury orbit

Figure 3: Composition of MCS, showing module

functions and contributions to MCS

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 fair-ing. 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: 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 communication links in the MPO orbit when the Earth can be located in any direction with respect to


5

the spacecraft body. Both antennas are accommodat-ed 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 ob-structions 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: 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 Device provided by JAXA remains bolted to the MOSIF.

Photo courtesy of JAXA

Figure 6: MMO flight model

The MMO in stowed configuration can be ac-commodated in a cylinder Ø 1800 x 1200 mm.

MOSIF Configuration The MOSIF acronym (MMO SunShield and In-

terFace Structure) very much describes its functional-ity, which then leads to its configuration.

The MOSIF consists of two major assemblies: The Adapter has at its centre the circular inter-

face to the MMO, the arms provide the at-tachments 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 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 7: MOSIF – with MMO installed

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 thrusters being illuminated.


6

Figure 8: MTM deployed configuration

MCS Configuration The MCS configuration is represented by a stack-

ing of the above described modules.

Figure 9: Cross-section through the MCS

Table 2: Overall Mass Budget

Mass Budget MMO [kg]

MOSIF [kg]

MPO [kg]

MCSA [kg]

MTM [kg]

MCSL [kg]

Dry Mass 285 127 1169 1580 1134 2714

Propellant

MOI CPS 602 602 602

MPO CPS 67 67 67

Cruise CPS n/a 157 157

Xenon n/a 581 581

Total Mass 285 127 1838 2249 1872 4121

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.

Figure 10: From launch to MPO orbit showing

gravity assists and separations

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 re-mainder 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 grav-ity assist manoeuvres: 1 x EGA, 2 x VGA and 5 x MGA. During the electric propulsion thrust phas-es the MCS is orientated with the thrusters pointing


7

in the flight direction to provide braking and with thruster steering supporting the attitude control sys-tem. 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 ma-noeuvres 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.

Subsystem Designs and BepiColombo specific solutions Attitude and Orbit Control System (AOCS) For MPO science operations the AOCS must pro-

vide continuous nadir pointing whilst meeting accu-racy and stability requirements. Two Star Trackers (plus a 3rd for redundancy) and an Inertial measure-ment 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 off-loading). The AOCS also controls the thermally criti-cal orientation of the solar array and the 22 N thrust-ers for orbit manoeuvres during Mercury Orbit Inser-tion.

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 con-trolled 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 mo-mentum 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 nomi-nal attitudes in the order of seconds before overheat-ing and damage occurs. In the event of an OBC re-boot, 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 configura-tions.

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 pro-vided 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.

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 recep-tion 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 oper-ations when it supports up- and downlinks in both X- and Ka-band.

The newly developed DST (Deep Space Tran-sponder) 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 X- and 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.


8

MTM Electric Propulsion System (MEPS) The MEPS contains the 4 electric propulsion

(Kaufmann) thrusters, their power processing elec-tronics, 4 thruster pointing mechanisms and the xen-on 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 long-est 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 cre-ated 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.

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 compo-nents 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

tilts the array to a maximum of 75° from the sun. The OSRs occupy 17% of the panel area. The tilted, sin-gle wing configuration is needed to achieve sun illu-mination whilst permanently nadir pointing through-out 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 regu-lated 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 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° 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 in-crease 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.

Chemical Propulsion Systems (CPS) The MPO CPS is tasked with the 15 MOI ma-

noeuvres 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 bi-propellant 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.


9

Thermal Control The MPO TCS must regulate the equipment tem-

peratures (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 mount-

ing panels to collect and transfer the heat the radiator panel

Spreader heatpipes in the radiator panel, ther-mally connected to the equipment panels by 90° linking heatpipes

97 heatpipes are used, of which a few are 3-dimensional hence difficult to test on ground

Fixed louvres are mounted in front of the radi-ator to reflect the planet infrared radiation away from the radiator whilst allowing the ra-diator 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 alumini-um layers. The Nextel layers reach 380°C

Moving inwards to lower temperature, 26 lay-ers 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.

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

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

The MTM TCS must regulate the equipment tem-peratures, 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 mod-ules 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 at-tached 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 connect-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 13: Deployable Thermal Cover to close

separation apertures


10

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 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

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 ac-commodate smaller items. The floors and shear pan-els 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 ele-ments (MTM and MPO structures are separated by 130 mm, with a smaller gap to the MOSIF) are con-nected by Inter Module Hardware (IMH) completing the MCS structure at launch and providing the sepa-ration 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 separa-tion functions and dynamics (both electrical and mechanical), the IMH requires a mass of 47 kg.

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 on-board 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 ena-bles 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 conjunc-tions. By this means, for example, it will be possible to continue MEPS operation during solar conjunc-tions – for which back-up thruster configurations and corresponding attitudes are stored on-board.

In the post-launch, near-Earth phase ground con-tact 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 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


11

new hardware developed for BepiColombo, an exten-sive 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 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 demonstrate adequate workmanship. The modules were transferred to ESTEC in July 2014 for the con-tinuation of assembly and the verification pro-gramme.

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 near-Earth 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 para-mount.

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-

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 suc-cessfully verified in a large scale back-to-back com-parison 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.

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 respec-tive STMs. The avionics and electrical systems have been developed and demonstrated on their respective


12

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

References [ 1 ] Johannes Benkhoff, Jan van Casteren,

Hajime Hayakawa, Masaki Fujimoto, Harri Laakso, Mauro Novara, Paolo Ferri, Helen R. Middleton, Ruth Ziethe, “BepiColombo—Comprehensive explo-ration 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 Pow-er System”, Proc. ‘9th European Space Power Con-ference’, Saint Raphaël, France, 6–10 June 2011 (ESA SP-690, October 2011)

[ 4 ] Eurostar 3000 satellite, Airbus DS http://www.space-airbusds.com/en/programme/eurostar-series-czw.html

[ 5 ] Federica Tessarin, Domenico Battaglia and Tiziano Malosti, Daniele Stramaccioni, Jürgen Schil-ke, “Thermal Control Design of Mercury Planetary Orbiter”, AIAA 2010-6090, 40th International Con-ference on Environmental Systems

[ 6 ] Large Space Simulator (LSS), at ESTEC, Noordwijk, The Netherlands http://www.esa.int/Our_Activities/Space_Engineering_Technology/Large_Space_Simulator_LSS

http://www.space-airbusds.com/en/programme/eurostar-series-czw.html

http://www.space-airbusds.com/en/programme/eurostar-series-czw.html

http://www.esa.int/Our_Activities/Space_Engineering_Technology/Large_Space_Simulator_LSS

http://www.esa.int/Our_Activities/Space_Engineering_Technology/Large_Space_Simulator_LSS

Documents

The BepiColombo Spacecraft, its Mission to Mercury and its … · 2015. 3. 18. · The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification 46th Lunar and Planetary