Flight Operations Preparation for the BepiColombo Mission to Mercury: Concepts and Challenges

Transcription

1 SpaceOps Conferences May 2016, Daejeon, Korea SpaceOps 2016 Conference / Flight Operations Preparation for the BepiColombo Mission to Mercury: Concepts and Challenges Christoph Steiger 1, Elsa Montagnon 2, Andrea Accomazzo 3 ESA/ESOC, Robert-Bosch Str. 5, Darmstadt, Germany The ESA/JAXA BepiColombo mission to Mercury features a complex modular design, with two scientific Mercury orbiters and a cruise module. The spacecraft and mission design lead to a number of challenges for flight operations, including electric propulsion usage during cruise to Mercury, complex module separation and orbit insertion operations upon arrival at Mercury, sophisticated solar array guidance profiles due to thermal constraints, strict attitude constraints requiring frequent updates of critical spacecraft guidance, and the use of Ka-band for science data downlink. This paper presents the concepts, challenges and particularities for BepiColombo operations. I. Introduction epicolombo is an ESA cornerstone mission to Mercury in collaboration with the Japanese Space Agency B(JAXA) with the objective to study the planet and its environment, in particular global characterization of Mercury through the investigation of its interior, surface, exosphere and magnetosphere. A. Mission Overview The BepiColombo mission consists of two scientific spacecraft, ESA s Mercury Planetary Orbiter (MPO) and JAXA s Mercury Magnetospheric Orbiter (MMO), launched together as a single composite, including a dedicated propulsion module (MTM). MPO and MTM are developed under ESA contract by an international consortium led by Airbus Defence and Space Germany. BepiColombo is planned to be launched in April 2018 with Ariane-5 from Kourou. The launch will be followed by a 6.5 years cruise phase, including planetary swingbys at Venus and Mercury, eventually achieving a weak capture by Mercury in late 2024 (Fig. 1). During the cruise phase, electric propulsion will be used for extended periods of time. This is provided by the MTM module, which will be jettisoned at Mercury arrival. A set of complex manoeuvres will deliver the MMO to its operational orbit, and finally the MPO will be put into a 1500x480 km polar orbit (orbital period of about 2.2h) to start its scientific mission, planned to last for one Earth year (1 year extension possible). See Ref. 1 for an in depth overview. Figure 1. Cruise trajectory for April 2018 launch, showing sun distance, electric propulsion usage ( SEP ), and planetary flybys. 1 BepiColombo Spacecraft Operations Engineer, Mission Operations Department, ESA/ESOC Darmstadt, Germany. 2 BepiColombo Spacecraft Operations Manager, Mission Operations Department, ESA/ESOC Darmstadt, Germany. 3 Head of Solar and Planetary Missions Division, Mission Operations Department, ESA/ESOC Darmstadt, Germany. 1 Copyright 2016 by ESA. Published by the, Inc., with permission.

2 B. The BepiColombo Spacecraft Fig. 2 shows an artist s impression of BepiColombo. The combined stack can have the following configurations: - Mercury Composite S/C Cruise (MCSC): MTM, MPO, MMO sunshield (MOSIF) and MMO - Mercury Composite S/C Approach (MCSA): MPO, MOSIF and MMO following separation of the MTM - Mercury Composite S/C Orbit (MCSO): MPO and MOSIF following release of the MMO The JAXA-provided MMO is a passive passenger during cruise, not involved in the control of the composite, which is done centrally within the MPO. The MPO accommodates 11 scientific instruments and has a box-like shape with a size of m, and a dry mass of about 1080 kg. The tremendous heat load at Mercury imposes strong requirements on the spacecraft design, requiring high-temperature multi-layer-insulation and solar array technology. A radiator to dump excess heat into space is mounted on one S/C side, which may not be exposed to sun or Mercury. The MPO AOCS performs 3-axis stabilised attitude and orbit control employing star trackers, inertial measurement units, fine sun sensors, reaction wheels and chemical propulsion. AOCS design is impacted by the challenging environment, requiring special guidance profiles for the MPO solar array (to avoid overheating) and rapid S/C attitude stabilisation in case of contingencies. A separate processing unit, the Failure Control Electronics (FCE), is taking over S/C attitude control in case of transient unavailability of the main on-board computer at safe mode entry, when the on-board computer is rebooting. For communications, a X/Ka-band deep space transponder with moveable high and medium gain antennae (HGA/MGA) is used. The MTM provides propulsion means for cruise. Apart from dual mode bipropellant chemical propulsion, it features electric propulsion with 4 moveable, Kaufman-type thrusters (max thrust 145 mn). The high power demand by the MTM electric propulsion (up to 11 kw) is satisfied with large solar arrays (area of over 40 m 2 in total) using the same hightemperature technology as for the MPO. See Ref. 5 for more details. Figure 2. Artist s impression of BepiColombo in cruise configuration (top), with the various elements of the cruise stack in exploded view (bottom left), and the MPO at Mercury (bottom right). II. BepiColombo Mission Operations Approach and Preparation Status A. The BepiColombo Mission Operations Centre at ESA/ESOC Operations of the composite spacecraft and the MPO will be conducted from ESA s European Space Operations Centre (ESOC) in Darmstadt, Germany. See Fig. 3 for the main elements of the BepiColombo Mission Operations Centre at ESOC. It features the typical setup for ESA/ESOC deep space missions, including a SCOS-2000 based mission control system, a standalone mission planning system, and a SIMSAT-based S/C simulator. The simulator is a key tool for operations preparation, as it is running the platform on-board software on a processor emulator, allowing testing with very high fidelity 3. The Engineering Test Bed (ETB) shown in the figure will only be delivered from industry to ESOC after launch. 2

3 Operations of BepiColombo are performed by the Flight Control Team (FCT), consisting of about 10 engineers and controllers at launch, led by the Spacecraft Operations Manager. The FCT is interfacing with various multi-mission support groups at ESOC, including Flight Dynamics (in charge of orbit determination and command generation for AOCS guidance and control), software support for the mission control system, and ground station operations. For deep space missions, there is a particularly close relation to the Flight Dynamics team due to the Figure 3. Main elements of the BepiColombo ground segment at ESOC. complex navigation and AOCS operations activities. Prior to launch, the FCT deals with all aspects of operations preparation. Key activities include the following: - Specification and acceptance testing of mission control system, planning system and simulator. - Preparation of operational products: writing of the Flight Operations Plan (FOP) using the MOIS suite of tools 4 based on inputs provided by the prime S/C contractor, as well as population of the industry-provided spacecraft database, refining and augmenting it with operations-specific information (e.g. generation of displays, instantiation of specific TCs, TM parameter out of limits, etc.). - Execution of tests with the spacecraft flight model as well as the engineering test bed, with the prime objective of validating the ESOC ground segment and operational products. B. Mission Operations Preparation Status Compared to a typical mission operated by ESA/ESOC, the preparation status for BepiColombo is particularly advanced considering that launch is still two years away. The Ground Segment Implementation Review was passed successfully in The AIT schedule for BepiColombo is such that the MPO is in a more mature state earlier, while the full cruise composite including the MTM will only be completed later. Operations preparation at ESOC hence had to initially focus on MPO stand-alone, prior to shifting the attention to MCSC configuration (which will be operated in the first 7 years of the mission). Status of the main elements of the ground segment: While some preparation activities will only be performed closer to launch, the overall status of the main elements of the ground segment is advanced: - Mission Control System: all functionality available; future work consists of the deployment of the system for the operational configuration at launch, as well as maintenance to keep in line with ESOC infrastructure evolution. - Simulator: functionally almost complete; reception of final delivery and ensuing acceptance testing to be completed within Mission Planning System (MPS): the BepiColombo MPS is a joint development with Rosetta, meaning that the MPS software has already been used operationally during Rosetta comet operations. Work on the operational setup for BepiColombo cruise operations has started and is planned to be completed in Flight Dynamics System: implementation has started, with an initial focus on operations in MPO configuration. Complex testing activities on the flight model have already been supported, providing the relevant guidance products (see below). Preparation of operational products: Mainly driven by the tests with the flight and engineering models, preparation of operational products has started. A first issue of the flight operations plan is planned for May In particular in the area of AOCS, the 3

4 preparation of operational products is well advanced, covering virtually all activities possible in MPO configuration. This also includes validation of the commanding interface from Flight Dynamics to the FCT. Testing with the flight model and engineering model: ESOC has been granted 30 days testing with the flight model (currently located at ESA/ESTEC) prior to launch, the so-called System Validation Tests (SVT). In addition, ESOC can also access the engineering test bed (ETB) located at the premises of the S/C manufacturer for 30 days, the so-called Integrated Ground Space Tests (IGST). Tests in the various slots are carefully prepared and iterated with the S/C manufacturer and the Project team at ESA/ESTEC. ESOC performs testing activities in these slots with the same ground segment systems as used for flight, resulting in highly representative test results. The following tests have been run so far: - IGST-1/-2: first tests with the ETB in 2012/2013, focusing on data handling operations (6 days in total). - IGST-3/-4: test slots in late 2014 and mid 2015, with a main focus on AOCS closed-loop operations in MPO configuration (6 days in total). - SVT-1a: first SVT with the flight model in MPO configuration in early March Apart from performing all AOCS closed-loop operations in MPO configuration, TM/TC validation for all MPO platform subsystems were executed. All MPO payloads were also operated in this slot. SVT-1a comprised 3 days of testing in double shift, corresponding to 6 normal working days of testing time. AOCS closed-loop runs in MCSC configuration have already been performed successfully on the ESOC simulator. The first IGST in MCSC configuration is planned for Q3/2016. III. BepiColombo Operational Challenges A. Spacecraft Modularity A key aspect of BepiColombo is that it is a modular spacecraft. In all configurations outlined in section I.B, the TM/TC interface with ground is with the MPO s on-board computer, hence the overall look and feel for the operator is similar for all S/C configurations. However, the differences in subsystems used as well as the different constraints for the various configurations are significant, such that operations preparation for BepiColombo is in many ways akin to having to prepare more than one mission at the same time. An obvious consequence of S/C modularity is the partial duplication of S/C subsystems on MPO and MTM, with the cruise module possessing its own power subsystem, thermal control, electric propulsion as well as 2 chemical propulsion systems (for S/C attitude control in safe mode, and the execution of axial orbit control maneouvres), all of which leads to a substantial increase in operations complexity. AOCS operations in particular are impacted by S/C modularity. Preparations of attitude slew or orbit control maneouvres have to take into account the vastly different S/C characteristics. Different AOCS guidance contexts need to be maintained depending on the S/C configuration. For instance, the S/C attitude when entering safe mode is configuration-dependent (ref. Fig. 4). AOCS solar array guidance is entirely different between MCSC and MPO/MCSA/MCSO configurations (see section III.B). As a result, the interface between the Flight Control Team Figure 4. Different safe mode attitudes depending on S/C configuration: +Y sun pointing for MCSC, sun close to +X for MCSA/O, sun close to Z for MPO. Safe mode concept is to have a rotation around the sun line, which has to be in synch with the orbital motion around Mercury in MPO and MCSO configurations. 4

5 and the Flight Dynamics team for commanding the S/C is particularly sophisticated. Planning of ground station passes requires careful evaluation of visibility when using the MGA/HGA antennae, as blockage of these anteanne by the S/C body is configuration-dependent, and there are configuration-specific limitations on MGA movement. Albeit only a crude measure for judging S/C complexity, a comparison on the number of telemetry parameters in the operational database (containing all elements used to perform spacecraft operations) between BepiColombo and some previous ESA missions shows the complexity of the former, see Fig. 5 (note: MMO not considered for BepiColombo). Another consequence of S/C modularity is the need for module separation operations upon arrival at Mercury. This includes separation of the MTM about 2 months before arrival, release of the MMO into its target orbit and separation of the MOSIF during the MOI sequence (ref. section III.E). MTM separation itself is performed autonomously by a dedicated on-board control procedure (OBCP), with commissioning of the MPO chemical propulsion system and MPO power subsystem performed by ground thereafter. B. Solar Array Control Because of the intense heat, the single-sided MPO solar array features a mix of solar cells and Optical Surface Reflectors (OSR) to keep its temperature below 200 C. The large MTM solar arrays (area of over 40 m 2 in total) use the same high-temperature technology and can provide up to 13 kw power. During cruise, the entire composite is powered through the MTM SA, while the MPO SA is only required at Mercury (remaining edge on to the sun during cruise to limit degradation). Both arrays can be rotated around their longitudinal axis using dedicated solar array drive mechanisms, under control of the AOCS. To maintain the temperature in the allowed range, both arrays require a special control approach, commanding an offpointing while still achieving sufficiently high power generation. MPO solar array control at Mercury: Due to Mercury albedo and infrared radiation, the maximum exposure of the MPO solar array towards the sun varies over the MPO operational orbit. The MPO SA hence has to be rotated continuously to avoid violation of temperature limits. The allowed sun aspect angle and the corresponding MPO SA angle is calculated by Flight Dynamics based on a thermal model provided by the S/C manufacturer, specifying the MPO SA target operating temperature as input. The commanding products for updating the on-board MPO solar array guidance are then generated accordingly. Fig. 6 shows the complicated AOCS data structure for the MPO SA guidance: each MPO orbit will be covered by an Envelope consisting of 7 segments (specified as Chebychev Polynomials of maximum order 12), which is valid for a specifiable number of orbits. The thermal model used for the MPO solar array will be calibrated in orbit. This activity will already start after MTM separation about 2 months before Mercury capture, when the S/C is not yet in Mercury orbit and hence MPO SA control is still less critical. MTM solar array control in cruise: Down to a sun distance of 0.62 AU, the MTM solar array can be pointed straight at the sun with no thermal limitations. At sun distances smaller than that, the array must also be offpointed to not violate maximum operating temperatures. The operational approach is as follows: Figure 5. Number of TM parameters in the database for missions operated by ESOC. Figure 6. Data structure for MPO solar array guidance. 5

6 - Outside of electric propulsion thrust arcs: S/C power consumption is low, allowing to offpoint the array to reduce degradation effects. Array pointing is commanded such that a fixed amount of power is available to operate the S/C. This is done using a curve provided by the S/C manufacturer, showing the sun distance vs. the sun aspect angle required to obtain the desired amount of power. - During electric propulsion thrust arcs: the array is pointed to supply maximum power during thrust arcs, with the electric propulsion system as the main power consumer. Information on the maximum allowed sun aspect angle depending on sun distances below 0.62 AU has been provided by the manufacturer. As the power available drives the maximum possible thrust level for electric propulsion in turn impacting the trajectory, a dedicated interface will be put in place between the FCT and Flight Dynamics, with the FCT providing the maximum power available based on the latest in-flight S/C power budget. As for the MPO solar array, it will be essential to calibrate MTM solar array performance in flight, to adjust the models and processes for MTM SA control accordingly. The AOCS guidance function for MTM SA control is simpler than for the MPO SA, relying on a table containing 20 segments with time intervals and the corresponding MTM SA angle. It is currently estimated that at most 2 segments per day are needed during cruise, compatible with a weekly MTM SA guidance update. C. Spacecraft Attitude Constraints and Guidance Owing to the harsh thermal environment at Mercury and during parts of the cruise phase, even short S/C mispointings can t be tolerated. This is addressed by the S/C design as follows: - To prevent an offpointing at safe mode entry, the Failure Control Electronics (FCE) running a simplified AOCS software temporarily takes over control, using separate equipment and context information. - The AOCS does not perform an attitude acquisition from lost in space conditions at safe mode entry (e.g. sun search using sun sensors), but is instead relying on attitude context information continuously maintained, allowing a rapid correction of any mispointing discovered at safe mode entry. The following guidance and ephemeris context is required by the AOCS for functioning correctly following a safe/survival mode entry in MPO configuration, with the AOCS firstly entering Sun Acquisition and Survival Mode (SASM) and eventually performing a transition to Safe Hold Mode (SHM): - Last known S/C attitude and rates: context information stored in the non-permanent area of safeguard memory (SGM RAM) at 1Hz when the AOCS is running nominally. - Sun and Earth ephemerides: in the early phases of SASM, sun pointing is established based on last known S/C attitude/rates and the sun direction based on the sun ephemerides. Earth ephemerides is required for pointing the medium and high gain antennae to the Earth. - MPO SASM attitude and solar array guidance: SASM attitude guidance specifies S/C rotation rate and phasing for the rotation around the sun line, which has to be in synch with the orbital motion around Mercury, ensuring the radiator side of the MPO (-Y side, ref. Fig. 4) is never facing Mercury, see Fig. 7. SASM solar array guidance consists of tables with time ranges and commanded sun aspect angles, allowing for a stepwise adjustment of the array position as the S/C rotates around the sun line. - MPO SHM attitude and solar array guidance: SHM is the highest mode reached autonomously after safe/survival mode. For MPO, its attitude and solar array guidance is specified as Chebyshev polynomials in safeguard memory. Apart from the last known S/C attitude and rates maintained autonomously in-flight, ground is in charge of keeping the guidance up to date. While sun and Earth ephemerides don t need to Figure 7. MPO safe mode attitude at Mercury: rotation around the sun line (red arrow) in synch with orbital motion around Mercury, such that the radiator is never exposed. 6

7 be updated frequently, this is not the case for the attitude and solar array guidance in safe mode: the orbit of the MPO around Mercury is left to drift, hence it is expected that weekly safe mode guidance updates are needed to adhere to the strict attitude constraints. Utmost care has to be taken in the preparation and verification of the guidance products uplinked, as any serious mistake could lead to a loss of mission at the next safe mode entry. While this is also true for a typical interplanetary mission, what sets BepiColombo apart is the high frequency of these critical updates in safeguard memory. The absence of a sun search at safe mode entry also complicates the setup of closed-loop runs during ground testing, as one can t rely on the S/C autonomously acquiring a correct attitude at safe mode entry. The typical process for setting up a closed loop run for an Figure 8. Time correlation in SVT-1a. ESOC test on the flight model includes the following steps: 1) Well before the test, a custom scenario in terms of S/C orbit and attitude is agreed with all parties. ESOC Flight Dynamics provides the products for updating the guidance, and AIT prepares the setup of the AOCS SCOE accordingly. 2) On the testing day, AIT configures the S/C with AOCS remaining in Standby Mode (SBM). 3) After TC handover, ESOC loads the context information on S/C attitude and rates into SGM RAM (inflight, this step would not be needed as the AOCS maintains this knowledge autonomously). 4) A UTC time for starting the closed-loop simulation is agreed with AIT. The AOCS guidance products are then shifted accordingly from simulation epoch time (e.g. a 2024 Mercury orbit case) to UTC time by ESOC and uplinked. For this, it is crucial to have a correct time correlation in the mission control system between UTC and S/C on-board time (ref. Fig. 8 for a schematic overview). 5) Coordinating on the voice loop, at the agreed time the dynamic simulation is started by AIT in the AOCS SCOE, and ESOC commands an OBC reboot at the same time, leading to S/C safe mode entry. The AOCS closed-loop run then starts, with all context information correctly retrieved and the S/C in the right attitude. See also Ref. 2 for a more detailed discussion. D. Electric Propulsion Operations in Cruise Following electric propulsion usage on ESOC-controlled missions Smart-1 and GOCE, BepiColombo is ESA s first interplanetary mission using this technology. The Mercury Electric Propulsion System (MEPS) of the MTM contains 4 thrusters, their power processing electronics, 4 thruster pointing mechanisms, and the xenon storage and feed system (see Fig. 9): - The QinetiQ T6 thruster (Ø 22 cm, 145 mn) derived from the T5 flown on the GOCE mission is used. The system typically operates using two thrusters (usage of a single thruster is possible as well). - The thrusters are mounted on individual pointing mechanisms, allowing to adjust the thrust vector. - The xenon storage and feed system includes flow control units, pressure regulators and 3 tanks, and is able to store 580 kg of xenon for provding a deltav of about 5400 m/s. - The MEPS is controlled by the central platform on-board software. The AOCS has a specific Electric Propulsion Control Mode (EPCM) for MEPS operations. In EPCM, the AOCS is also capable of performing reaction wheel offloading by offpointing the thrust vector. For the April 2018 trajectory, 22 thrust arcs are planned with a total duration of 756 days, and the longest arc lasting about 150 days. See Fig. 1 for the distribution of thrust arcs throughout cruise. Interplanetary navigation for a spacecraft Figure 9. MEPS configuration on the MTM (up in the figure is in direction +Z as per Fig. 4). 7

8 employing electric propulsion is significantly different from missions relying on chemical propulsion only. The prelaunch trajectory planning by mission analysis (as depicted in Fig. 1) is relying on conservative assumptions on the available thrust level, in turn dependent on the maximum power available from the MTM solar array. If the available power in-flight should turn out to be higher, corresponding higher thrust levels will be used to shorten a thrust arc, reducing its criticality. This will be assessed by the operations teams prior to each thrust arc, and revised on a weekly basis during a thrust arc, with the Flight Control Team providing the latest figures on the power available to Flight Dynamics for planning of the thrust arc (see also section III.B on MTM solar array control). Electric propulsion implies increased operations complexity compared to traditional chemical propulsion systems. The MEPS will be used over extended periods of time, requiring regular commanding activities (e.g. regular adjustment of thrust level and thrust vector). Routine maintenance activities by ground include the adjustment of pressure regulation set points and neutralizer flowrates, as well as monitoring the possible degradation of the system. For xenon bookkeeping, a combination of a pressure/volume/temperature (PVT) method and a method integrating the measured mass flow will be used. Commissioning operations after launch feature a stepwise checkout of the MEPS, culminating in the firing of the thrusters in EPCM. The maximum non-coverage period that can be accommodated by the mission design is shorter than in the long coasting phases of a chemical propulsion mission. The baseline foresees ground contacts on a weekly basis both during and outside of thrust arcs, which is feasible provided the MEPS works reliably. This approach will be validated and confirmed in the first thrust arc, planned to be supported with daily ground contacts. For some isolated thrust arcs, the specific geometry prevents from having ground contact using either MGA or HGA, requiring to interrupt thrusting once a week to reorient the S/C as needed for ground contact. E. Mercury Orbit Insertion Operations Separation of the MTM is performed about 2 months before the S/C is weakly captured by Mercury in an initial orbit of x674 km (April 2018 launch scenario). The Mercury Orbit Insertion phase (MOI) starts thereafter, including a series of chemical propulsion manoeuvres with the aim of achieving the operational orbit firstly for the MMO (11639x590 km, i=90 deg, RAAN=67.8 deg, w=-2 deg) and eventually for the MPO (1500x480 km, i=90 deg, RAAN=67.8 deg, w=16 deg). Operations in this phase are driven by the following main constraints: - Below a certain altitude, the S/C rotation around the sun line has to be synchronized with the orbital motion around Mercury, to ensure thermal limits are not violated. - Manoeuvres shall not take place around Mercury perihelion ±60 deg due to thermal constraints. - The S/C undergoes eclipse seasons during MOI, which are power-critical in the higher orbits. Special operational measures like boost heating prior to eclipse entry are expected to be required for ensuring a positive power budget. It is imperative to sufficiently lower the orbit prior to the aphelion eclipse season. In particular, a failure to separate the MMO as planned prior to the eclipse season could be missioncritical in case the orbit has not been lowered sufficiently. - Operational constraints on manoeuvre execution: a delta time of at least 3 days is observed between manoeuvres. No manoeuvres are allowed during solar conjunction periods (no Figure 10. Mercury Orbit Insertion (MOI) sequence for launch in April 2018 (depicted in ecliptic J2000 frame). 8

9 ground contact possible) and as from 7 days before (a failed manoeuvre shortly before a solar conjunction may lead to the S/C using incorrect guidance and hence a violation of thermal constraints, with no ground intervention possible). This leads to a rather constrained MOI timeline as shown in Fig. 10. Five initial burns are performed to reduce the apoherm altitude to the MMO target value of km. Following separation of the MMO, the MOSIF is separated shortly after, bringing the S/C into MPO configuration. Another 10 manoeuvres are then required to achieve the MPO operational orbit. Duration of the MOI phase is about 3 months, with a total deltav for the sequence shown in the figure of about 963 m/s. F. Ka-band Operations To increase the scientific return without increasing the duration of the ground station contacts, the BepiColombo transponder includes a Ka-band transmitter in addition to the traditional X-band receiver/transmitter. Use of Kaband on the downlink was technically validated on Smart-1, but this will be the first operational use on a scientific ESA mission. The quality of a Ka-band link is strongly dependent on weather conditions at the ground station. This is addressed both in the S/C design and the operations approach: - A space-to-ground closed-loop file transfer protocol is provided, allowing to automatically recover any lost data due to unpredictable Ka-band link variations, similar to a file transfer protocol used on the uplink for BepiColombo (as well as for previous ESA interplanetary missions). - Selection of an adequate downlink rate for Ka-band operations depending on the expected weather conditions affects the scientific return: if the planning is too conservative, the advantages of Ka-band may not be fully exploited. If it is too optimistic, too much data will need to be retransmitted. While the precise operational concept is yet to be detailed (this will only be relevant for routine operations at Mercury, i.e. not before early 2025), ESOC is currently running studies dedicated to special tools that incorporate local weather forecasts for optimizing the downlink bit rate. Outcome of a first study activity was that by introducing a 1-day weather forecast in the operations concept, a potential advantage of up to 20% in data volume could be achieved as compared to methods based on availability of seasonal or monthly statistics of the attenuation and brightness temperature. IV. Conclusion and Outlook Flight operations preparations at ESA/ESOC for the 2018 ESA/JAXA BepiColombo mission to Mercury are at an advanced state. With the main elements of the ESOC ground segment in place, the particular operational challenges of this interplanetary mission are now being addressed. While sharing some commonalities with previous deep space missions operated at ESOC, the specifics of BepiColombo multi-modularity of the spacecraft, attitude and solar array pointing constraints requiring complex guidance products, first usage of electric propulsion and Kaband science data downlink are a main driver for the preparation activities by the operations team. Tests run by ESOC with the flight model and the engineering model have already taken place, with an initial focus on the ESA mercury orbiter configuration, and the preparation of the corresponding operational products. For future tests, the focus will now shift to the cruise stack, the S/C configuration in the first 7 years of mission. Operations preparation will eventually culminate in the final preparations for launch, including the simulations campaign taking place within the last 6 months before flight. Appendix A Acronym List AIT AOCS APME BMCS EPCM ETB FCE FCT FD Assembly, Integration and Test Attitude and Orbit Control System Antenna Pointing Mechanism Electronics (for High Gain and Medium Gain Antenna) BepiColombo Mission Control System Electric Propulsion Control Mode Engineering Test Bed Failure Control Electronics Flight Control Team Flight Dynamics 9

10 FDIR FOP HGA IGST IMU LEOP MCS MCSA MCSC MCSO MEPS MGA MMO MOI MOSIF MPO MPS MTM NM OBC OBCP OBT OCM OSR PRE RAM RIU RWU SA S/C SADE SASM SBM SCOE SEP SGM SHM SVT TPE Failure Detection, Isolation and Recovery Flight Operations Plan High Gain Antenna Integrated Ground Space Test Inertial Measurement Unit Launch and Early Orbit Phase Mission Control System Mercury Composite Spacecraft Approach (stack consisting of (MPO, MOSIF, MMO) Mercury Composite Spacecraft Cruise (stack consisting of MTM, MPO, MOSIF, MMO) Mercury Composite Spacecraft Orbit (stack consisting of MPO, MOSIF) Mercury Electric Propulsion System Medium Gain Antenna Mercury Magnetospheric Orbiter (JAXA-provided orbiter) Mercury Orbit Injection MMO Sunshade and Interface Structure Mercury Planetary Orbiter Mission Planning Systm Mercury Transfer Module AOCS Normal Mode On-board Computer On-board Control Procedure On-board Time AOCS Orbit Control Mode Optical Surface Reflector Pressure Regulation Electronics (of the electric propulsion) Random Access Memory Remote Interface Unit Reaction Wheel Unit Solar Array Spacecraft Solar Array Drive Electronics AOCS Sun Acquisition and Survival Mode AOCS Standby Mode Special Checkout Equipment Solar Electric Propulsion Safeguard Memory AOCS Safe Hold Mode System Validation Test Thruster Pointing Electronics (of the electric propulsion) References 1 Benkhoff, J., van Casteren, J., Hayakawa, H., Fujimoto, M., Laakso, H., Novara M., et al BepiColombo - Comprehensive Exploration of Mercury: Mission Overview and Science Goals, Planetary and Space Science, Vol. 58, Issues 1-2, Elsevier, pp Steiger, C., Altay, A., Montagnon, E. and V. Companys, AOCS Operations Preparation for the BepiColombo Mission to Mercury, 6 th European Conference for Aeronautics and Space Sciences (EUCASS), Krakow, Poland, Clerigo, I., Montagnon, E. and D. Segneri A Simulated Journey to Mercury: the Challenges of the BepiColombo Simulator Development for the Flight Control Team. SpaceOps 2014, Pasadena, USA, Rhea Group White Paper MOIS. 5 Wilson, R. J., Schelkle, M., The BepiColombo Spacecraft, its Mission to Mercury and its Thermal Verification, 46 th Lunar and Planetary Science Conference, The Woodlands, USA,