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BepiColombo Mission

Spacecraft    MPO    Status   Launch   MMO    MTM    Ground Segment    References

BepiColombo is one of ESA's (European Space Agency) cornerstone missions being conducted in cooperation with Japan, it will explore Mercury, the planet closest to the Sun. Europe's space scientists have identified the mission as one of the most challenging long-term planetary projects, because Mercury's proximity to the Sun makes it difficult for a spacecraft to reach the planet and to survive in the harsh environment found there. The scientific interest in going to Mercury lies in the valuable information that such a mission can provide to enhance our understanding of the planet itself as well as the formation of our Solar System; information which cannot be obtained with observations made from Earth. The overall goal is to study and understand the composition, geophysics, atmosphere, magnetosphere and history of Mercury, the least explored planet in the inner Solar System. In particular, the mission has the following scientific objectives: 1) 2) 3) 4) 5) 6) 7) 8)

- Investigate the origin and evolution of a planet close to the parent star

- Study Mercury as a planet: its form, interior structure, geology, composition and craters

- Examine Mercury's vestigial atmosphere (exosphere): its composition and dynamics

- Probe Mercury's magnetized envelope (magnetosphere): its structure and dynamics

- Determine the origin of Mercury's magnetic field

- Investigate polar deposits: their composition and origin

- Perform a test of Einstein's theory of general relativity

Set to arrive at Mercury in 2024, BepiColombo will investigate properties of the innermost planet of our Solar System that are still mysterious, such as its high density, the fact that it is the only planet with a magnetic field similar to Earth's, the much higher than expected amount of volatile elements detected by NASA's Messenger probe and the nature of water ice that may exists in the permanently shadowed areas at the poles.

The BepiColombo mission is named after Professor Giuseppe (Bepi) Colombo (1920-1984) from the University of Padova, Italy, a mathematician and engineer of astonishing imagination. He was the first to see that an unsuspected resonance is responsible for Mercury's habit of rotating on its axis three times for every two revolutions it makes around the Sun. He also suggested to NASA how to use a gravity-assist swing-by of Venus to place the Mariner 10 spacecraft in a solar orbit that would allow it to fly by Mercury three times in 1974-5.

ESA's Science Program Committee decided at its meeting in Naples in 1999 to name the Mercury cornerstone mission in honor of Giuseppe Colombo's achievements.

Mercury is small compared to the Earth, with a diameter of only 4878 km. It orbits the Sun in an elliptic orbit between 0.3 and 0.47 AU from the Sun. Mercury is difficult to observe from the Earth, due to its close proximity to the very bright Sun. For an in-depth study of the planet and its environment, it is therefore necessary to operate a spacecraft equipped with scientific instrumentation around the planet. It is, however, difficult for a spacecraft to reach Mercury, as even more energy is needed than sending a mission to Pluto. Departing from Earth, a spacecraft needs to decelerate to come closer to the Sun and as the solar gravitational force increases with the square of the distance, the required reverse thrust increases accordingly. Furthermore, the thermal environment close to the Sun and close to the hottest planet in the solar system is extremely aggressive, as the direct solar radiation is 10 times higher than at Earth's distance. 9)

Despite the advances in space flight and the growth in planetary research over the last few decades, enabling detailed investigations of the Earth, Mars, Venus, the outer planets and several moons and asteroids, scientists have not been able to observe much of Mercury. For a very long time the data delivered by NASA's (US National Aeronautics and Space Administration) Mariner 10, which visited Mercury in1974-1975, was among the best available. During these flybys Mariner 10 was able to image about 45% of the planet's surface and to discover its unexpected magnetic field. Further discoveries by Mariner 10 are the existence of gaseous species forming an exosphere and the presence of a unique magnetosphere. However, little to nothing is known about Mercury's interior structure or its elemental and mineralogical composition.

With the launch of the NASA Discovery class mission MESSENGER (Mercury Surface, Space ENvironment, Geochemistry and Ranging) in 2004, the first spacecraft was launched to orbit Mercury. MESSENGER already collected data from two Venus flybys and three Mercury flybys in 2008 and 2009. The orbital phase of the MESSENGER mission started in March 2011. The MESSENGER data will provide valuable discoveries of Mercury and its environment that can be used by the BepiColombo mission in tuning its observations to the most important investigations of planet Mercury.

NASA's MESSENGER mission came to a planned end on May 30, 2015 when it slammed into Mercury's surface at about 14,000 km/h and created a new crater on the planet's surface. MESSENGER ended up exceeding its planned mission timeline by three years, by which time the spacecraft had completely depleted its fuel. The last of the fuel was used to position it within the gravitational pull of Mercury and the Sun, so it could delay as long as possible its inevitable plummet towards the surface – while continuing to beam back images – and go out with a bang. 10)


Figure 1: Messenger's iridescent Mercury (image credit: NASA, JHU/APL, Carnegie Institution, Washington) 11)

Legend to Figure 1: The contrasting colors have been chosen to emphasize the differences in the composition of the landscape across the planet. The darker regions exhibit low-reflectance material, particularly for light at redder wavelengths. As a result, these regions take on a bluer cast.

- The crisscrossing streaks across the disc of the planet show up in shades of light blue, grey and white. These regions take on a light blue hue for a different reason: their youthfulness. As material is exposed to the harsh space environment around Mercury it darkens, but these pale ‘rays' are formed from material excavated from beneath the planet's surface and sent flying during comparatively recent impacts. For this reason, they have retained their youthful glow.

- The yellowish, tan-colored regions are "intermediate terrain". Mercury also hosts brighter and smoother terrain known as high-reflectance red plains. One example can be seen towards the upper right, where there is a prominent patch that is roughly circular. This is the Caloris basin, an impact crater thought to have been created by an asteroid collision during the Solar System's early days.

As the nearest planet to the Sun, Mercury has an important role in showing us how planets form. Mercury, Venus, Earth and Mars make up the family of terrestrial planets; each one carrying essential information to trace the history of the whole group.

The knowledge of how they originated and evolved is key to understanding how conditions supporting life arose in the Solar System, and possibly elsewhere. As long as Earth-like planets orbiting other stars remain inaccessible to astronomers, the Solar System is the only laboratory where scientists can test models applicable to other planetary systems.

Exploring Mercury is therefore fundamental to answering important astrophysical and philosophical questions such as 'Are Earth-like planets common in the Galaxy?'

A European mission to Mercury was first proposed in May 1993. Although an assessment showed it to be too costly for a medium-size mission, ESA made a Mercury orbiter one of its three new cornerstone missions when the Horizon 2000 science program was extended in 1994. Gaia competed with BepiColombo for the fifth cornerstone mission. In October 2000, ESA approved a package of missions for 2008–2013 and both BepiColombo and Gaia were approved.

In February 2007, the mission was approved as part of the Cosmic Vision program. Following an unavoidable increase in the mission's mass during 2008, the launch vehicle was changed from Soyuz-Fregat to Ariane 5. Final approval for the redesigned mission was given by ESA's Science Program Committee in November 2009.

BepiColombo represents the first time ESA and JAXA have joined forces for the implementation of a major space science mission.

BepiColombo's mission is especially challenging because Mercury's orbit is so close to our star, the Sun. The planet is hard to observe from a distance, because the Sun is so bright. Furthermore, it is difficult to reach because a spacecraft must lose a lot of energy to ‘fall' towards the planet from the Earth. The Sun's enormous gravity presents a challenge in placing a spacecraft into a stable orbit around Mercury.

Only NASA's Mariner 10 and Messenger missions have visited Mercury so far. Mariner 10 provided the first-ever close-up images of the planet when it flew past three times in 1974-1975. En route to its final destination in orbit around Mercury in 18 March 2011, Messenger flew past the planet 3 times (14 January 2008, 6 October 2008, and 29 September 2009), providing new data and images. Once BepiColombo arrives in 2024, it will help reveal information on the composition and history of Mercury. It should discover more about the formation and the history of the inner planets in general, including Earth.

Mercury is a major Roman god. He is the patron god of financial gain, commerce, eloquence (and thus poetry), messages/communication (including divination), travelers, boundaries, luck, trickery and thieves; he is also the guide of souls to the underworld. In Greek mythology, Hermes is an Olympian god of transitions and boundaries. In the Roman adaptation of the Greek pantheon, Hermes is identified with the Roman god Mercury. — Hence, in an orbit around the planet Mercury, the point that is closest to Mercury is termed "periherm" while the farthest point of a spacecraft orbit is called "apoherm".

Since Mercury is the closest planet to the sun (0.31 AU to 0.47 AU distant) a peak solar intensity of 11 solar constants (14,500 W/m2) is experienced which imposes enormous thermal challenges on the spacecraft modules and their external equipment.

Table 1: Some history and background 12)

The space segment design is driven essentially by the scientific payload requirements,the launch mass constraints and the harsh thermal and radiation environment at Mercury. Key technologies required for the implementation of this challenging mission include the following:

• High-temperature thermal control materials (coatings, adhesives,resins,MLI,OSR).

• Radiator design for high-infrared environment.

• High-temperature and high-intensity solar cells,diodes and substrates for the solar arrays.

• High-temperature steerable high-gain and medium-gain antennas.

• High specific impulse(Isp=4300 s)and high total impulse (23.7 mNs),to be provided by gridded ion engines.

• Payload technology,such as detectors,filters and laser technology.



Space segment:

The BepiColombo mission is based on two spacecraft:

1) MPO (Mercury Planetary Orbiter) to map the planet. MPO is a three-axis stabilized and a nadir pointing spacecraft with an instrument suite of 11 experiments and instruments. MPO is led by ESA. The MPO will focus on a global characterization of Mercury through the investigation of its interior, surface, exosphere and magnetosphere. In addition, it will test Einstein's theory of general relativity.

2) MMO (Mercury Magnetospheric Orbiter) to investigate its magnetosphere. MMO is a spinning spacecraft carrying a payload of five experiments and instruments. The MMO is led by JAXA.

Among several investigations, BepiColombo will make a complete map of Mercury at different wavelengths. It will chart the planet's mineralogy and elemental composition, determine whether the interior of the planet is molten or not, and investigate the extent and origin of Mercury's magnetic field.

MPO is ESA's scientific contribution to the mission. JAXA/ISAS (Japan Aerospace Exploration Agency/Institute of Space and Astronautical Science) is providing the MMO. ESA is also building the MTM (Mercury Transfer Module), which will carry the two orbiters to their destination, and the MOSIF (MMO Sunshield and Interface Structure), which provides thermal protection and the mechanical and electrical interfaces for the MMO. The MCS (Mercury Composite Spacecraft) consists of the MPO, MMO, MTM and MOSIF. ESA is responsible for the overall mission design, the design, development and test of the MPO, MTM and MOSIF, the integration and test of the MCS and the launch. 13) 14)


MPO (Mercury Planetary Orbiter)

MMO (Mercury Magnetospheric Orbiter)


3-axis stabilized

15 rpm spin-stabilized


Nadir pointing

Spin axis at 90º to Sun

Orbit at Mercury

Polar orbit, period of 2.3 hr, 480 x 1500 km

Polar orbit, period of 9.3 hr, 590 x 11,640 km

Spacecraft mass

4100 kg (total mass at launch)
1150 kg (in Mercury orbit)

275 kg (in Mercury orbit)

Spacecraft size

3.9 x 2.2 x 1.7 m (excluding solar wings)

1.9 m ∅ x 1.1 m

Payload mass, power

80 kg, 100-150 W

45 kg, 90 W

Telemetry band



Data volume (downlink)

1550 Gbit/year

160 Gbit/year

Equivalent average data rate

50 kbit/s

5 kbit/s


High-temperature resistant 1.0 m X/Ka-band high-gain steerable antenna

0.8 m X-band phased array high-gain antenna

Operational lifetime at Mercury

> 1 year

> 1 year

Table 2: Key parameters of the two spacecraft

Launch, journey and orbit:

The BepiColombo trajectory employs a solar electric propulsion system so that a combination of low-thrust arcs and flybys at Earth, Venus and Mercury are used to reach Mercury with low relative velocity. A brief summary of the key stages in the journey to Mercury are given here:

• Launch on Ariane 5 in April 2018 on escape trajectory to reach Venus

• Cruise trajectory with solar electric propulsion stage - the SEPM (Solar Electric Propulsion Module), up to 290 mN thrust - plus six gravity assists: Venus (twice) and Mercury (four times)

• Approximately 6.7 year cruise phase to Mercury

• Ion propulsion stage jettisoned shortly before arrival at Mercury

• Capture and insertion by chemical propulsion engines within the MPO

• On reaching MMO orbit the MMO is released

• MOSIF is released before further descending to the MPO orbit

• MPO is inserted into final orbit using thrust from chemical propulsion engines

• For MPO and MMO: one Earth-year (4 Mercury years) operations in Mercury orbit with optional one year extension.

Key mission dates


April 2018


25 July 2019

First Venus flyby

20 May 2020

Second Venus flyby

09 April 2021

First Mercury flyby

27 March 2022

Second Mercury flyby

16 December 2023

Third Mercury flyby

24 January 2024

Fourth Mercury flyby

18 December 2024

Arrival at Mercury

27 March 2015

MPO in final orbit

01 May 2026

End of nominal mission

01 May 2027

End of extended mission

Table 3: Key mission dates for a 2018 launch into a heliocentric transfer orbit


Figure 2: Artist's rendition of the BepiColombo MCS (Mercury Composite Spacecraft) in cruise configuration heading toward Mercury (image credit: ESA) 15) 16)


Figure 3: Artist's rendition of BepiColombo's MPO and MMO spacecraft in their respective Mercury orbits (image credit: ESA, C. Carreau)


Figure 4: Artist's view of the BepiColombo spacecraft MPO (ESA, foreground) and MMO (JAXA, background) at Mercury (image credit: ESA/ATG medialab. The Mercury image was taken by NASA's Messenger spacecraft, image credit: NASA, JHU/APL, and Carnegie Institution) 17)



Industrial involvement:

• In November 2006, ESA awarded the prime contract for the Implementation Phase to Airbus Defence and Space, former EADS Astrium of Friedrichshafen, Germany. The PDR (Preliminary Design Review) was completed in October 2008.

This included the design and procurement of the 'cruise-composite' spacecraft, including the ESA's MPO (Mercury Planetary Orbiter), the MTM (Mercury Transfer Module), the MMO's sunshield and the interface between the MPO and the MMO. 18) Furthermore, the prime contractor provides the design and development of the data management and attitude and orbit control subsystems, and the integration of the engineering model (Ref. 3).

• In December 2012, TAS-I (Thales Alenia Space-Italia) signed a contract with Astrium GmbH for BepiColombo. TAS-I is part of the industrial Core Team, coordinating 35 European manufacturers within its workpackage. The contract concerns the telecommunications, thermal control and electric power distribution systems, along with satellite integration and testing, plus support during the launch campaign. In addition, TAS-I is developing the X- and Ka-band transponders, onboard computer, mass memory and high-gain antenna, a 1.1 m diameter dish antenna that will enable the satellite to communicate with Earth, while also carrying out a Radio Science experiment during the mission. 19)

• In the UK, Airbus DS (formerly Astrium Ltd.) is the co-prime contractor for the electrical and chemical propulsion systems, for the structure of all modules and for the thermal control of the MTM (Mercury Transfer Module). Airbus DS in France will develop the onboard software.

• The MMO and its scientific payload are designed and developed by JAXA. They are responsible for procuring the spacecraft from an industrial team led by NEC.

MCS (Mercury Composite Spacecraft):

The composite spacecraft, as shown in Figure 5, consists of four modules: the MPO (Mercury Planetary Orbiter), the MMO (Mercury Magnetospheric Orbiter) protected by the MOSIF (MMO SunShield and InterFace Structure) and the MTM (Mercury Transfer Module). The MPO,developed by ESA, is the scientific module characterized by a set of observation instruments to study the surface of Mercury and the gravity field of the planet. The MMO, developed by JAXA, is a spin stabilized scientific module aimed at the study of the magnetic field of Mercury. Both orbiters, the MPO and the MMO, will be carried on top of the transfer module MTM during the cruise phase. 20) 21) 22)


Figure 5: Schematic view of the BepiColombo MCS (Mercury Composite Spacecraft), image credit: Airbus DS

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 (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 optimization 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 (Mercury Magnetospheric Orbiter) 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 stabilized 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 (Mercury Composite Spacecraft -Launch) and MCSC (Mercury Composite Spacecraft-Cruise) are composed of:

1) MPO (Mercury Planetary Orbiter)

- Spacecraft optimized for its operational mission

- Performs command & control for MCS (with only minor hardware modification for MCS configurations, notably the size of the reaction wheels to control the MCS)

- CPS (Chemical Propulsion System) not used during cruise, however after MTM separation the MPO performs approach propulsion and apoherm lowering of Mercury orbit.

2) MMO (Mercury Magnetospheric Orbiter)

- Spins during operational mission

- Is passive during cruise – apart from checkouts.

3) MOSIF (MMO SunShield and InterFace Structure)

- Thermal protection for the MMO

- Mechanical interface for the MMO

- Harness routing between MPO and MMO

4) 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 (Mercury Composite Spacecraft-Approach) is created upon MTM separation. On reaching the MMO orbit, the MMO is released to create the MCSO (Mercury Composite Spacecraft-Orbit). The MOSIF is ejected shortly afterwards to leave the MPO.

Power Subsystem: Each of the 3 modules MPO, MMO and MTM contains power generation, storage and distribution hardware. The MPO and MMO power subsystems supply standard 28 V regulated power within the module, with the MMO being supplied by the MPO, as long as both modules are connected. During the cruise phase the power subsystem is a composite of all 3 modules (controlled from the MPO) whereby, all power is provided by the MTM solar array. The MTM generates both 100 V and 28 V supplies. 23)


Figure 6: Composition of MCS (Mercury Composite Spacecraft), showing module functions and contributions to MCS (image credit: Airbus DS)



MPO (Mercury Planetary Orbiter) Spacecraft

The MPO design is optimized to meet the needs of the payload when the spacecraft is in its operational orbit. The payload components are mounted on the nadir side of the spacecraft, with certain instruments or sensors located directly at the main radiator, to achieve low detector temperatures (Ref. 13).

Structure: The spacecraft structure uses a double-H configuration, designed to harmonize with the single radiator plane necessitated by the Mercury orbit. Heat generated by spacecraft subsystems and payload components, as well as heat that is coming from the Sun and Mercury as it "leaks" through the blankets into the spacecraft, is carried to the radiator by panel-embedded heat pipes. The structural design provides free access to all equipment and instruments during the AIT (Assembly, Integration and Test) program. The design is mass efficient, with the primary structure serving as the mounting surface for all equipment; it will remain permanently assembled during AIT, avoiding the need for connector brackets and preventing alignment disturbances. MPO has four-point bolted interfaces to both the MTM (Mercury Transfer Module) and the MOSIF (MMO Sunshield and Interface Structure), which provides thermal protection and the mechanical and electrical interfaces for the MMO during the journey to Mercury.

The configuration and thermal design provide a classical thermal environment for internally mounted instrument equipment – avoiding costly development programs by re-use of available hardware – while employing dedicated high temperature technologies for external items such as antennas, the solar array, the sun sensors and MLI (Multi-Layer Insulation), which are exposed to the harsh thermal environment around Mercury.


Figure 7: MPO – showing equipment panel perpendicular to radiator (image credit: Airbus DS)


Figure 8: MPO spacecraft in deployed configuration (image credit: Airbus DS)

Power system generation: The MPO provides a 28 V regulated bus which also feeds the MMO during Cruise. The MPO includes a 96 Ah lithium-ion battery. The SA (Solar Array) uses both OSRs (Optical Solar Reflectors) and the control of the sun incidence angle to maintain the temperature below 190ºC. For most of the Mercury year, the solar array requires continuous rotation, in order to generate adequate power while at the same time limiting the temperature. The three-panel array has its rotation axis in an optimized direction, nevertheless, the so-called "artificial eclipses" (a condition in which the Sun vector is along the rotational axis of the solar array, i.e. no power is available) occur for short periods at certain times of the Mercury year due to the solar array mounting geometry. The solar array will provide up to 1000 W of electrical power during full science operations phases. During both the natural and artificial eclipses the battery in the MPO will provide electrical energy to the spacecraft in order to allow the scientific operations to continue without interruption (Ref. 23).

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 m2 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.

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.

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.

AOCS (Attitude and Orbit Control Subsystem): The AOCS equipment consists of:

• Three STRs (Star Trackers), each comprising a star tracker unit housing the optics and electronics, a shutter - which can be closed in the event of a major attitude control anomaly - and a baffle

• Two IMUs (Inertial Measurement Units), including four high-accuracy rate-integrating gyros and four accelerometers in a tetrahedral configuration, together with the processing electronics

• Two redundant sets of two FSS (Fine Sun Sensors)

• Four reaction wheel assemblies, controlled by two sets of wheel drive electronics

• Two redundant sets of four 22 N hydrazine / MON-3 (Mixed Oxides of Nitrogen), a mixture of nitrogen tetroxide and 3% of nitric oxide thrusters to provide the change in velocity (ΔV) needed for orbit capture and orbit lowering to the MMO and MPO operational orbits

• Two redundant sets of four 10 N monopropellant (hydrazine) thrusters for attitude control and reaction wheel momentum off-loading.

The reaction wheels are mounted in a tetrahedral configuration; attitude control can be achieved with four wheels operating simultaneously (the nominal operational scenario) or any combination of three wheels.

During science operations, at least two STRs will be used in combination. In the event of major system anomaly on the spacecraft and consequent loss of attitude control, dedicated shutters will protect the STR optical paths to prevent damage due to accidental sun pointing.

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 IMU are co-mounted with instruments on an optical bench while 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 maneuvers during MOI (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 the cruise phase, 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 minimizing 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.

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.

AOCS operations in particular are impacted by S/C modularity. Preparations of attitude slew or orbit control maneuvers 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 (Figure 9). AOCS solar array guidance is entirely different between MCSC and MPO/MCSA/MCSO configurations. As a result, the interface between the Flight Control Team and the Flight Dynamics team for commanding the S/C is particularly sophisticated.


Figure 9: 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 (image credit: ESA)

TCS (Thermal Control Subsystem): The MPO TCS must regulate the equipment temperatures (achieving standard equipment levels), 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 (Ref. 22):

- 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 3-dimensional hence difficult to test on ground

- Fixed louvers 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 aluminum layers. The Nextel layers reach 380°C

- Moving inwards to lower temperature, 26 layers of aluminized Upilex are followed by 10 layers of aluminized Mylar

- Spacers of glass fiber and AAerofoam are used to separate the layers in the 4 packets, while Kapton rosettes separate the packets

- The installed MLI has a thickness of 65 mm. The total MLI mass is 94 kg.


Figure 10: Photo of a section through the MPO MLI (image credit: Airbus DS)

The MOSIF MLI must shade the MMO and limit the infrared heat load to the MMO. The MOSIF MLI is characterized 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 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:

- While 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 (Figure 112).

- 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 11: Photo of a DTC (Deployable Thermal Cover) to close separation apertures (image credit: Airbus DS, Ref. 22)

DHS (Data Management System): 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 while 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.

The SSMM (Solid State Mass Memory) is a stand-alone unit in the BepiColombo MPO DMS (Data Management System). The SSMM interfaces with the BepiColombo payload instruments and the OBC (On-Board Computer) via main and redundant SpaceWire links. The SSMM stores telemetry packets according to the CCSDS for later downlink via X- or Ka-band. The SSMM also routes telecommand (TC) packets from the OBC to the relevant payload instrument and the returning telemetry reports from instruments to OBC (Ref. 128).

SSMM has a capacity of 384 Gbit. This storage area is organized in packet stores (maximum 50 packet stores active in parallel) for telemetry data storage. There are two types of packet stores that can be created in the SSMM: cyclic packet stores - when the packet store is full, old data is overwritten; and non-cyclic packet store - when the packet store is full the data storage is interrupted (that means new data can not be stored and is lost) and an action from ground is necessary in order to free space by deleting old data via telecommand.

The telemetry science data packets are stored in the SSMM packet stores based on PIDs (Process ID). One PID can only be associated to one SSMM packet store at a time, but several PIDs can be routed to the same SSMM packet store. The instruments will generate low- and/or high-priority science data and store it in different packet stores based on the PIDs.

CPS (Chemical Propulsion Systems): 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 bipropellant while the 5 N thrusters are monopropellant: 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 ΔV plus attitude control.


RF communications: The MPO is equipped with two fixed LGAs (Low Gain Antennas), a 2-axis steerable MGA (Medium Gain Antenna) and a 2-axis steerable 1.1 m diameter HGA (High Gain Antenna). The two X-band LGAs will provide omnidirectional coverage at small distances from Earth and can also be used for emergency commanding at any distance. The X-band MGA will be used primarily during the interplanetary cruise phase and in safe and survival modes. The HGA will provide X-band uplink and downlink and Ka-band downlink communications for spacecraft and science operations. The HGA will also be used during the cruise phase to enhance communications and data dump capabilities whenever needed. 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 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 while the 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.

ESA's Cebreros 35 m ground station (Ávila, Spain) is planned to be the primary ground facility for communications during all mission phases. The ground stations at Kourou (LEOP), New Norcia (critical phases during cruise and Mercury capture), Perth (LEOP), Usuda (backup) and Uchinoura (backup) will be available for backup during critical flight phases and/or for use during special campaigns.

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 Ka-band on the downlink was technically validated on the Smart-1 mission of ESA, 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.


Launch: The ESA-JAXA BepiColombo mission to Mercury was launched on 20 October 2018 (01:45:28 GMT) with Ariane-5 ECA (VA245) from Europe's Spaceport in French Guiana. This was the third and final cornerstone mission of the Horizon 2000+ program. — Signals from the spacecraft, received at ESA's control center in Darmstadt, Germany, via the New Norcia ground tracking station (Australia) at 02:21 GMT confirmed that the launch was successful. 24) 25) 26) 27)

The launch delay decision was made after a major electrical problem was detected during preparations for a thermal test of the MTM (Mercury Transfer Module), one of the major spacecraft elements of BepiColombo. The six-month postponement will have no impact on the science return of the mission. However, the new flight time to Mercury will be 7.2 years, and BepiColombo will now arrive in December 2025, one year later than previously anticipated. The seven-year cruise to the innermost planet of our Solar System will include 9 flybys of Earth, Venus and Mercury. The MTM, MPO and MMO are currently undergoing intensive tests in ESA/ESTEC (European Space Research and Technology Center) in the Netherlands. Everything is going well with the MPO and MMO. The last of the instrument flight models was installed recently on the MPO. 28)


Figure 12: The ESA-JAXA BepiColombo mission to Mercury lifts off from Europe's Spaceport in Kourou (image credit: 2018 ESA-CNES-Arianespace)

Orbit: The launch will be followed by a 7.2 years cruise phase, including planetary swingbys at Venus and Mercury, eventually achieving a weak capture by Mercury in December 2025 (1 year later than previously planned). 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. The seven-year cruise to the innermost planet of our Solar System will include 9 flybys of Earth, Venus and Mercury. — A set of complex maneuvers will deliver the MMO to its operational orbit, and finally the MPO will be put into a 1500 x 480 km polar orbit (orbital period of about 2.2 hr) to start its scientific mission, planned to last for one Earth year (with a 1 year extension capability).


Figure 13: Cruise trajectory for April 2018 launch, showing the sun distance, SEP (Solar Electric Propulsion) usage, and planetary flybys (image credit: ESA)

Figure 14 shows the BepiColombo spacecraft. The combined stack can have the following configurations:

1) MCSC (Mercury Composite Spacecraft - Cruise): MTM, MPO, MMO sunshield (MOSIF) and MMO

2) MCSA (Mercury Composite Spacecraft - Approach): MPO, MOSIF and MMO following separation of the MTM

3) MCSO (Mercury Composite Spacecraft - Orbit): MPO and MOSIF following release of the MMO.


Figure 14: 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), image credit: ESA

MAP (Mercury Approach Phase): The MAP starts after the last electric propulsion maneuver has been completed, approximately two months before the first Mercury orbit insertion maneuver. During this phase, the MTM will separate from the spacecraft stack. The remaining composite of MPO/MMO/MOSIF, the MCSA configuration, will drift into Mercury's sphere of influence, and will need only a small maneuver to get captured in an initial orbit of approximately 590 x 178,000 km ((April 2018 launch scenario). This process is known as a 'weak stability boundary' capture.

MOI (Mercury Orbit Insertion) phase: The MOI starts thereafter, including a series of chemical propulsion maneuvers with the aim of achieving the operational orbit firstly for the MMO (11,639 x 590 km, i=90º, RAAN=67.8º, ω=-2º) and eventually for the MPO (1500 x 480 km, i=90º, RAAN=67.8º, ω=16º).


Figure 15: Schematic of the MOI sequence for launch in April 2018 (depicted in ecliptic J2000 frame), image credit: ESA

Operations in the MOI 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.

- Maneuvers 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 mission-critical in case the orbit has not been lowered sufficiently.

- Operational constraints on maneuver execution: a delta time of at least 3 days is observed between maneuvers. No maneuvers are allowed during solar conjunction periods (no ground contact possible) and as from 7 days before (a failed maneuver 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 Figure 15. Five initial burns are performed to reduce the apoherm altitude to the MMO target value of 11,639 km. Following separation of the MMO, the MOSIF is separated shortly after, bringing the S/C into MPO configuration. Another 10 maneuvers are then required to achieve the MPO operational orbit. Duration of the MOI phase is about 3 months, with a total ΔV for the sequence shown in the figure of about 963 m/s.

Mercury Orbit Phase: Once the MPO mission orbit is reached, the final commissioning of the MPO and its payload is performed; this will last about one month. The MPO attitude follows a continuous nadir-pointing profile, providing optimum viewing conditions for the payload.

All MPO science data will be stored in the spacecraft's solid-state mass memory and downlinked during daily station passes with ESA's Cebreros ground station. Every half Mercury year, about every 44 Earth days, the attitude of the spacecraft will have to be reversed around the nadir direction to keep the radiator pointing away from the Sun.

The MMO will communicate with the JAXA/ISAS Sagamihara Space Operations Center via the Usuda Deep Space Center (UDSC) 64 m antenna in Nagano, Japan.

Nominal mission science operations are scheduled to be performed for one Earth year, with a planned extension of another year.



Mission status

• November 16, 2018: In mid-December, twin discs will begin glowing blue on the underside of a minibus-sized spacecraft in deep space. At that moment Europe and Japan's BepiColombo mission will have just come a crucial step closer to Mercury. 29)

- This week sees the in-flight commissioning and test firing of the four thrusters – with one or two firing at a time – of the Solar Electric Propulsion System that BepiColombo relies on to reach the innermost planet. This marks the first in-flight operation of the most powerful and highest-performance electric propulsion system flown on any space mission to date.

- Each thruster and its associated power processing and propellant flow control units will be tested to full power to check no ill-effects were incurred from launch, culminating in the first twin thruster operations – the configuration to be used throughout most of the mission.

- Their first routine firing is scheduled for the middle of next month, and the propulsion system will operate continuously for three months to optimize the spacecraft's trajectory for the long voyage to Mercury.

- The voyage inward: Like all objects in the Solar System, the spacecraft is in solar orbit, moving perpendicular to the pull of the Sun's gravity. BepiColombo therefore has to slow down through a series of braking maneuvers and flybys, making it more susceptible to the Sun's gravity and letting it spiral closer to the heart of the Solar System.

- The thrust produced by the electric propulsion system serves to decelerate the spacecraft, or in some cases accelerates it to make its braking flybys more effective. No less than nine planetary flybys of Earth (once), Venus (twice) and Mercury itself (six times) are required to place the multi-module spacecraft in orbit around Mercury in seven years' time.


Figure 16: Two T6 gridded ion thrusters undergoing a joint test firing inside a vacuum chamber at QinetiQ in Farnborough, UK. BepiColombo's Solar Electric Propulsion System has four T6 thrusters for redundancy, with one or two operating at one time. The two thrusters needed testing to check they could be operated in close proximity for prolonged periods without any harmful interactions. In space the plumes seen here would not be visible; they occur due to vestigial gases building up inside the chamber. The glow from the thrusters would be visible however (image credit: QinetiQ)

- Space tug: The MTM (Mercury Transfer Module) portion of the spacecraft, containing the propulsion system, is in essence a high performance ‘space tug'. Its task is to perform all the active trajectory control maneuvers needed to convey the other portions of the BepiColombo ‘stack' – ESA's MPO (Mercury Planet Orbiter) and Japan's MMO (Mercury Magnetospheric Orbiter) – to Mercury orbit.

- The high performance of the propulsion system, in terms of the amount of fuel the thrusters require, is critical. Inert xenon gas is fed into the thrusters, where electrons are first stripped off the xenon atoms. The resulting electrically charged atoms, referred to as ions, are then focused and ejected out of the thrusters using a high voltage grid system at a velocity of 50,000 m/s.

- This exhaust velocity is 15 times greater than conventional chemical rocket thrusters, allowing a dramatic reduction in the amount of propellant required to achieve the mission.


Figure 17: Exploded view of the BepiColombo spacecraft components. From bottom to top these are: the Mercury Transfer Module, Mercury Planetary Orbiter, Sunshield and Interface Structure, and Mercury Magnetospheric Orbiter. The spacecraft are shown with solar arrays and instruments deployed (image credit: ESA/ATG medialab)

- "The propulsion system transforms electricity generated by the Mercury Transfer Module's twin 15 m-long solar arrays into thrust," explains ESA electric propulsion engineer Neil Wallace.

- "At full power, a thrust equivalent to the weight of three 1-euro coins is developed, meaning that the thrusters have to keep firing for long periods to be effective, but in the absence of any drag and assuming you are patient, the maneuvers that are possible and the payload that can be carried are dramatic."


Figure 18: T6 gridded ion thruster being test fired in October 2018 inside the LEEP2 vacuum chamber at QinetiQ in Farnborough, UK. In space the plume seen here would not be visible; it occurs due to vestigial gases building up inside the chamber. The glow from the thruster would be visible however (image credit: QinetiQ)

- Electrifying spacecraft propulsion: The four T6 thrusters around which the solar electric propulsion system is designed, have a heritage dating back decades. QinetiQ in the UK – formerly the UK Defence Evaluation and Research Agency and before that the Farnborough Royal Aircraft Establishment – has been researching electric propulsion since the 1960s.

- The first flight of their technology came with the 10 cm-diameter T5 thruster, a key element of ESA's 2009 gravity-mapping GOCE mission, where it allowed the satellite to orbit at the top of Earth's atmosphere for over three years, skimming through the diffuse atmosphere at the unprecedentedly low orbital altitude needed for the mission.

- The scaled-up T6 thrusters are 22 cm in diameter, the increase in size required for the higher thrust and lifetime requirements of the BepiColombo mission. And unlike GOCE's T5, these T6 thrusters are maneuverable, courtesy of gimbal systems developed by RUAG Space in Austria.

Figure 19: Mercury Transfer Module electric propulsion thruster steering test. Video showing a test of the mechanisms steering the four solar electric propulsion thrusters on BepiColombo's Mercury Transfer Module (speeded up by 20 times).The module will use a combination of electric propulsion and multiple gravity assists at Earth, Venus and Mercury to carry BepiColombo's two scientific orbiters – ESA's Mercury Planetary Orbiter and Japan's Mercury Magnetospheric Orbiter – to the innermost planet in our Solar System [image credit: ESA/ D. Tagliafierro (TAS-I)]

- "They are clever mechanisms that complicate the system design a bit – all the electrical cables and pipes have to cross a moving boundary – but add a lot to performance," adds Neil. "They ensure the thrust vector of either a single or double engine firing crosses through the center of gravity of the spacecraft, which changes over time as propellant is used up."

- Thruster operations are controlled using two Power Processing Units, the architecture of which are designed to support the firing of two T6s simultaneously even in the event of any system anomaly, guaranteeing the maximum thrust of 250 mN can be maintained.

- Injecting intelligence: "The intelligence of the system for autonomous thruster operation comes from these Power Processing Units – contributed by Airbus Crisa in Spain," explains Neil, "which supply the regulated voltages and currents to the thrusters based on instructions from ground control via the spacecraft on-board computer."


Figure 20: Along with the four T6 gridded ion thrusters, BepiColombo's Solar Electric Propulsion System also includes gimbal mechanisms for each thruster, three tanks of xenon gas, a high pressure regulator, four flow control units and two power processing units, providing the intelligence of the system. In addition several meters of high-voltage harness and piping are required to connect this complex system together (image credit: ESA)

- The other key elements are propellant FCUs (Flow Control Units), also overseen by the PPUs, and the high-voltage electrical harness. The FCUs ensure the correct flows of xenon gas are supplied to the thrusters and were developed by Bradford Engineering in the Netherlands to provide programmable flow rates.

- The various elements of the propulsion system have undergone individual and extensive performance and qualification testing ultimately concluding in a series of tests performed at QinetiQ's Farnborough site.


Figure 21: Test setup: The T6 gridded ion thruster being set up for test firing inside the LEEP2 vacuum chamber at QinetiQ in Farnborough. The far end of the vacuum chamber is covered with a protective carbon disc to prevent the sandblasting-like ‘sputtering' from being in the path of the beam (image credit: ESA)

- Testing times: The spacecraft configuration and the extreme nature of the BepiColombo mission – needing to function in thermal conditions akin to placing it in a pizza oven – often demanded similarly extreme test scenarios, pushing the solar electric propulsion technology and test facilities to their limits. "One important test early in the program was to ensure that two thrusters could be operated in close proximity for prolonged periods without harmful interactions," adds Neil. "They turned out to be remarkably tolerant of each other with no measurable effects."

- One of the biggest ironies of the thruster qualification for BepiColombo, heading close to the Sun, was the extreme minimum temperatures experienced by its ion thrusters.


Figure 22: BepiColombo plasma simulation: When the Mercury Transfer Module of the BepiColombo mission fires its electric propulsion thrusters an ion beam is extracted. This is created through the ionization of xenon propellant, generating the charged particles that can be accelerated further using an electric field. Together with gravity assist flybys at Earth, Venus and Mercury, the thrust from the ion beam provides the means to travel to the innermost planet (image credit: ESA/Félicien Filleul)

Legend to Figure 22: After escaping the pull of Earth's gravity with the Ariane 5 launcher, the spacecraft is on an orbit around the Sun. The transfer module then has to use its thrusters to brake against the mighty pull of the Sun's gravity. It also has to tune the shape of its orbit in order to make a series of nine gravity assist flybys at the planets before finally delivering the mission's two science spacecraft into Mercury orbit. — This image is an excerpt from a supercomputer simulation that models the flow of plasma around the spacecraft just after the high energy ion beam is switched on. An outline of the composite spacecraft with its extended solar arrays is included for reference.

- Neil explains: "Despite the fact the mission is headed to Mercury, the bulk of the spacecraft shadows the thrusters for very long periods and when not operating they naturally cool to temperatures way lower than ever tested in the past. We needed to prove they would turn-on and operate within specification when cooled to minus 150 C.

- "It was a remarkable testament to the robustness of the technology that even after temperatures sufficient to freeze the xenon in the pipes the thrusters were able to start and operate flawlessly."

- End of the journey: The propulsion system is dependent on the Mercury Planetary Orbiter's onboard computer for its control and command, so by itself it will not be able to function. Its ultimate fate is to be cast off, when the three-module BepiColombo stack separates before entering Mercury orbit, to circle the Sun indefinitely in the vicinity of the planet, letting the two science modules go to work.

- "At one point while planning the BepiColombo mission, the Mercury Transfer Module was planned to impact the planet," Neil comments, "a sort of Viking funeral that seemed fitting to all of us engineers."

- Gridded ion thruster technology will have a life far beyond BepiColombo however, with commercial applications in development, and future, even more ambitious ESA science missions set to rely on the technology.


Figure 23: After a seven year journey through the inner Solar System, BepiColombo will arrive at Mercury. While still on the approach to Mercury, the transfer module will separate and the two science orbiters, still together, will be captured into a polar orbit around the planet. Their altitude will be adjusted using MPO's thrusters until MMO's desired elliptical polar orbit is reached. Then MPO will separate and descend to its own orbit using its thrusters. The fine-tuning of the orbits is then expected to take three months, after which, the main science mission will begin (image credit: ESA, Ref. 29)

• On 4 November 2018, the Mercury Transfer Module (MTM) of the ESA-JAXA BepiColombo mission was commanded to conduct a special one-off activity, called the "solar array drive run-in". This operation consisted of performing several rotations of the MTM solar array over its full movement range, to clean the solar array drive mechanism slip ring from contaminants accumulated during the long ground testing phase before launch. 30)

- The operation saw the back side of the arrays turned towards the Sun, and at the same time into the field of view of one of the MTM's monitoring cameras – M-CAM 1. One image is presented here, with the exquisite details of the cabling and mechanisms on the backside of the array visible. Watch the full sequence here.

- The image has an exposure time of 20 milliseconds and a resolution of 1024 x 1024 pixels. The structure seen in the bottom corner is one of the sun sensor units on the MTM, with the multi-layered insulation visible.

• October 29, 2018: Soon after launch, the two 15 meter-long solar arrays of the BepiColombo MTM (Mercury Transfer Module) were pointed straight at the Sun to maximize the amount of power obtained, as is common practice for spacecraft powered by solar energy. On Saturday 27 October, a week after launch, the solar arrays were rotated to a position off-pointed by 54º with a rotation rate of about 0.5º/s. 31)

Figure 24: One of the MT M's three monitoring cameras, M-CAM 1, captured a sequence of images during the rotation, which took a bit less than two minutes to complete. With a picture taken every six seconds, the array is seen rotated by a further 3º in each subsequent image. Each image has an exposure time of 20 ms and a resolution of 1024 x 1024 pixels. The structure seen in the bottom corner is one of the sun sensor units on the MTM, with the multi-layered insulation visible. In some of the images ‘blooming' artefacts (the horizontal stripes in this orientation) are present due to overexposure (image credit: ESA/BepiColombo/MTM, CC BY-SA 3.0 IGO)

- While off-pointing the solar arrays reduces the power output, it also helps to slow down the degradation of the solar cells, which naturally occurs over time. As the MTM solar arrays are sized to provide up to 11 kW for operating the electric propulsion system – a major power consumer that is currently off – they are capable of providing far more power than is really needed at the moment. Hence, it is possible to rotate the arrays partially away from the Sun, yet still get plenty of power to supply the spacecraft.

- For checking out the electric propulsion system late in November and when starting the first period of continuous electric propulsion operations in mid-December, the arrays will be rotated towards the Sun again. As the spacecraft gets closer to the Sun later in the interplanetary cruise phase, the MTM solar arrays must be off-pointed to avoid overheating, while still ensuring that sufficient power is generated for the electric propulsion system.

- Management of power generation and consequently solar array positioning is one of the major challenges of BepiColombo operations, in the interplanetary cruise phase as well as during routine operations in Mercury orbit.

• October 26, 2018: The 2.5 m long boom carrying the magnetometer sensors onboard ESA's BepiColombo MPO (Mercury Planetary Orbiter) has been successfully deployed. The sensors are now prepared to measure the magnetic field on the way to Mercury. 32)

- Following launch last weekend, and having completed the LEOP on Monday (22 October), confirming the spacecraft and systems were healthy and functioning now they are in space, attention has now turned to checking the suite of scientific instruments on the science orbiters.

- As part of this activity, one more piece of hardware had to be deployed: the magnetometer boom onboard the MPO. The deployment, which took about one minute to complete, was captured in a series of images taken by one of the monitoring cameras onboard the Mercury Transfer Module (MTM).

Figure 25: The MTM is equipped with three monitoring cameras – or ‘M-CAMs' –which provide black-and-white snapshots in 1024 x 1024 pixel resolution. The magnetometer boom is seen in M-CAM 2. The images were taken with an exposure of 40 ms, and a time interval of six seconds between images, starting at 12:40:09 GMT (14:40:09 CEST) on 25 October. Eleven images were taken in the sequence – eight of them capture the motion of the boom, as seen here (image credit: ESA/BepiColombo/MTM , CC BY-SA 3.0 IGO)

- At the same time, the sensors in the boom itself recorded the local magnetic field during the deployment.

- The M-CAMs already returned space ‘selfies' in the days after launch, featuring the MTM's deployed solar wings and MPO's antennas – activities which were confirmed first by telemetry. A portion of the array can be seen towards the right in this orientation, and the cone-shaped medium-gain antenna is in the lower part of the image on the left.

- The monitoring cameras will be used at various occasions during the seven year cruise phase. While the MPO is equipped with a high-resolution scientific camera, this can only be operated after separating from the MTM upon arrival at Mercury in late 2025 because, like several of the 11 instrument suites, it is located on the side of the spacecraft fixed to the MTM during cruise.

- Once at Mercury, the magnetometer will measure the planet's magnetic field, the interaction of the solar wind, and the formation and dynamics of the magnetosphere – the magnetic ‘bubble' around the planet. Together with measurements captured by a similar instrument suite onboard JAXA's Mercury Magnetospheric Orbiter, the spacecraft will provide scientists with data that will help investigate the dynamic environment of the planet, as well as the origin, evolution and current state of the planet's magnetic field and its interior.

• October 22, 2018: A stunning early morning launch lifted the ESA/JAXA BepiColombo spacecraft into space on Saturday, 20 October. This marked the start of intensive, round-the-clock flight control activities to ensure the mission's health and functioning in the harsh environment of space. 33)

- At 13:45CEST on Monday 22 October, just 58 hours into its mission, the critical first segment of the fledgeling satellite's long voyage to Mercury was wrapped up, as teams at ESA's mission control center declared the critical LEOP (Launch and Early Orbit Phase) complete.

- The end of the beginning now beckons months of extensive in-orbit commissioning activities, in which operations teams will work extended hours daily until the end of December, performing tests to ensure the health of BepiColombo's science instruments, its propulsion and other systems.

- "Lots of people think you point a spacecraft in a particular direction and off it goes, making its own way to its final destination," says Elsa Montagnon, Spacecraft Operations Manager for BepiColombo. "In reality, the post-launch period is extremely busy, and so, too, is the long interplanetary cruise. In the next months, teams on ground will be working in 12-hour shifts, including weekends, to get the spacecraft on the right path to the smallest planet of our Solar System."

- In the hours before Saturday's liftoff, the Main Control Room of ESA's ESOC operations center was center stage for the network countdown — a synchronized sequence involving facilities at ESOC, Europe's Spaceport in Kourou, French Guiana, and at ground stations on four continents supporting the launch.

- The final GO/NOGO rollcall saw Flight Director Andrea Accomazzo check in with the flight controllers at ESOC, each confirming they were "go for launch".


Figure 26: BepiColombo Flight Operations Director Andrea Accomazzo seen during launch in the main control room at ESA's ESOC control center on Saturday, 20 October, 2018 (image credit: ESA/J. Mai)


Figure 27: Artist's impression of the upper stage and payload launch adapter separating from the BepiColombo spacecraft stack about 30 minutes after launch. In this view, the Mercury Transfer Module is at the left, the Mercury Planetary Orbiter is in the middle with its folded solar array facing the viewer, and the Mercury Magnetospheric Orbiter is hidden inside the sunshield visible at far right. The solar wings of the spacecraft open at a later stage (image credit: ESA/ATG medialab)

• October 22, 2018: Compilation of images taken by the three monitoring cameras onboard the BepiColombo MTM (Mercury Transfer Module) following the mission's launch at 01:45 GMT on 20 October. The monitoring cameras – or ‘M-CAMs' – returned black-and-white images in 1024 x 1024 pixel resolution of the deployed solar wing and antennas. 34)

- JAXA's MMO (Mercury Magnetospheric Orbiter) sits inside a protective sunshield on ‘top' of the MPO, and cannot be seen in these images.


Figure 28: M-CAM 1 imaged one of the deployed solar wings of the transfer module, while M-CAM 2 and M-CAM 3 captured the medium- and high-gain antennas on the MPO (Mercury Planetary Orbiter), image credit: ESA/BepiColombo/MTM , CC BY-SA 3.0 IGO

• October 20, 2018: The BepiColombo MTM (Mercury Transfer Module) has returned its first image from space. The view looks along one of the extended solar arrays, which was deployed earlier this morning and confirmed by telemetry. The structure in the bottom left corner is one of the sun sensors on the MTM, with the multi-layered insulation clearly visible. 35)

- The MTM is equipped with three monitoring cameras, which provide black-and-white snapshots in 1024 x 1024 pixel resolution. The other two cameras will be activated tomorrow and are expected to capture images of the deployed medium- and high-gain antennas onboard MPO (Mercury Planetary Orbiter).

- The monitoring cameras will be used on various occasions during the cruise phase, notably during the flybys of Earth, Venus and Mercury. While the MPO is equipped with a high-resolution scientific camera, this can only be operated after separating from the MTM upon arrival at Mercury in late 2025 because, like several of the 11 instrument suites, it is located on the side of the spacecraft fixed to the MTM during cruise.


Figure 29: View of the deployed MTM solar array (image credit: ESA/BepiColombo/MTM – CC BY-SA 3.0 IGO)

BepiColombo's monitoring cameras: The MTM of BepiColombo is equipped with three monitoring cameras (M-CAM), which provide black-and-white snapshots in 1024 x 1024 pixel resolution. The positions of the three cameras are indicated with the orange icons, and example fields of views are illustrated. 36)

- M-CAM 1 looks down the extended solar array of the MTM, while M-CAM 2 and 3 are looking towards MPO (Mercury Planetary Orbiter). The MPO's medium-gain antenna and magnetometer boom can be seen in M-CAM 2, once deployed. M-CAM 3 has the possibility to see the MPO's high-gain antenna. Since all deployable parts of the spacecraft are rotatable, a range of orientations may be seen in the actual images.

- The first sets of images are expected to be taken about 12 hours and 1.5 days after launch.


Figure 30: MTM camera locations (image credit: ESA)




MPO will carry a highly sophisticated suite of eleven scientific instruments, ten of which will be provided by Principal Investigators through national funding by ESA Member States and one from Russia (Ref. 3). 37)

BELA (BepiColombo Laser Altimeter)

BELA will characterize the topography and surface morphology of Mercury. It will also provide a digital terrain model that, compared with the data from the MORE instrument, will give information about the internal structure, the geology, the tectonics and the age of the planet's surface. Co-PIs (Principal Investigators): Nicolas Thomas, University of Bern, Switzerland, and Tilman Spohn, DLR, Germany. Further partners are the MPS (Max Planck Institute of Solar System Research) and the IAA (Instituto de Astrofisica de Andalucia). 38) 39)

BELA is the first European laser altimeter to be built for inter-planetary flight. A key element has been the development of a European high-power (50 mJ) pulsed Nd:YAG laser allowing instrument operation at distances of > 1055 km from the target. The nadirpointing geometry of MPO necessitated the use of baffles to reject the incoming sunlight (when MPO is over the nightside hemisphere but still illuminated by the Sun). The strict mass constraints combined with the expected large temperature excursions (arising in large part from the eccentricity of Mercury's orbit) drove the selection of a beryllium telescope as receiver. In addition, the competences within the participating countries led to adoption of a digital rangefinder concept.

Laser: The laser is a fully redundant 1064 nm Nd:YAG with 5 ns pulse duration and a nominal 50 mJ pulse energy. A beam expander collimates the beam to a 60 µrad width. The system can operate at up to 10 Hz, consumes 20 W with a mass of < 5 kg (including MLI, cabling, beam expander, and drive electronics).


Figure 31: Schematic view of the BELA instrument and its components ( BELA team)

Baffles: The receiver baffle follows a Stavroudis concept. It is an aluminium structure combining ellipsoidal and hyperbolic surface machined with 4 nm roughness. The internal diameter is 204 mm and an extremely thin wall thickness has been achieved to minimize mass. Although the transmitter baffle is smaller, it must also hold a thermal filter to prevent the beam expander focussing reflected light from Mercury on to the laser. A narrow band transmits the laser wavelength but rejects light outside a band around this wavelength.

Telescope: The receiver telescope is a two-mirror on-axis design with a 20 cm primary. The telescope is an all-beryllium design with a mass of roughly 600 g. The primary mirror is just 2 mm thick. The telescope surfaces have been produced using diamond-turning of a deposited copper layer followed by gold coating. The aperture at the vertex of the primary mirror is close to the focus of the telescope and supports the instrument straylight rejection concept.


Figure 32: The receiver telescope which is being manufactured out of beryllium (image credit: BELA team)

Rangefinder module: Unlike previous planetary laser altimeters, the rangefinding of BELA is performed using a digital approach where the signal is digitized and the return pulse detected using software in an FPGA. The resolution is limited by the digitization frequency and the bandwidth but tests indicate that in optimum conditions, accuracies of the order of 20 cm over the (typically) 500 km range can be achieved. The rangefinder can also detect fairly low return pulse energies. Testing indicates that a return pulse containing just 6 photons can be detected. The final system will have inferior performance because of noise contributions and the effects of radiation damage in flight. However, from a technological standpoint, the ranging system will meet the requirements foreseen in the original BELA proposal.

BELA operations: The 60 µrad wide beam is reflected from the surface (surface spot size = 20-50 m) and received around 5 ms later at a 20 cm diameter f/5 telescope. The image is refocussed onto a silicon avalanche photodiode through a narrow bandpass interference filter. The signal is then sampled and fed to a digital pulse discrimination electronics. This system determines the time of flight (and therefore range), the integrated pulse intensity, and its width. The data are passed to a digital processing unit which controls the operation and services the spacecraft interface. Onboard data compression and data storage are foreseen. The experiment requires significant baffling and thermal control but can operate over the dayside hemisphere (with only slightly reduced signal to noise) allowing optimum data acquisition over a minimum duration. BELA will provide 2 ns time resolution (30 cm range) which is commensurate with the expected knowledge of the spacecraft position. Optimum data return is expected at altitudes up to at least 1000 km above the surface. Samples will be acquired about every 250 m on ground-tracks separated by 25 km at the equator (crossing at the poles). Over the lifetime of the mission, data points will be 6 km apart (decreasing with latitude). The experiment will provide return pulse intensity and width information allowing an assessment of surface albedo and roughness at 20 m scales including in unilluminated polar craters.


ISA (Italian Spring Accelerometer)

The objectives of the ISA accelerometer are strongly connected with those of the MORE experiment. Together the experiments can give information on Mercury's interior structure as well as test Einstein's theory of the General Relativity. PI: V. Iafolla, CNR-IFSI, Italy. 40) 41) 42)

ISA is a three–axis high-sensitivity accelerometer, characterized by an intrinsic noise level of about 10-10g/√Hz in the frequency band 3 x 10-5 — 1 x 10-1 Hz. The main goal of ISA it to measure the very strong non–gravitational accelerations acting on the MPO (Mercury Planetary Orbiter) spacecraft, which are an important source of error in the RSE (Radio Science Experiments) measurements. The non-gravitational accelerations are proportional to the area-to-mass ratio of the spacecraft, and are very difficult to be properly modeled for a complex in shape and active satellite like the MPO. 43)

The main objectives of RSE are to perform precise measurements of:

• gravitational field of Mercury

• rotation of Mercury

• general relativistic effects, in particular Mercury perihelion precession by state-of-the-art radiometric tracking of the MPO spacecraft.

Indeed, the modeling depends on a set of parameters related with the physical properties of the satellite surface and structure, which will be strongly influenced, and with completely unknown laws, by the strong radiation environment in the surroundings of Mercury. In order to reach the ambitious objectives of the RSE, the a posteriori reconstruction of the MPO orbit should reach the 10-8 m/s2 level in acceleration over a time span of one orbital revolution of the spacecraft around Mercury, i.e. about 2.3 hours. The ISA measurements have to be integrated with the radar tracking measurements from Earth's stations in a very precise orbit determination procedure. The RSE are a complex mix of measurements and scientific objectives, and it is not possible to separate them neatly in independent experiments. These experiments are based, from one side, on a sophisticated and very precise tracking system, both in range and range-rate, that will use a full 5-way frequency link from Earth's ground stations to the MPO (X-band, Ka-band and a mixed mode).

From the other side, a precise orbit determination software and procedure is needed in order to reconstruct the orbit of the MPO around Mercury, and of Mercury center-of-mass around the Sun, while solving in a complex least-squares fit for local and global parameters. Finally, the ISA measurements will also be useful to estimate the speed variations produced by the onboard thrusters firings during the offloading maneuvers of the spacecraft reaction wheels, at least once every 24 hours.


Figure 33: Illustration of the ISA elements (image credit: CNR-IFSI)


MERMAG (Mercury Magnetometer)

The MPO's MERMAG will provide measurements that will lead to the detailed description of Mercury's planetary magnetic field and its source, to better understand the origin, evolution and current state of the planetary interior, as well as the interaction between Mercury's magnetosphere with the planet itself and with the solar wind. PI: Karl-Heinz Glassmeier, Technical University of Braunschweig, Germany Co-PI: C.M. Carr, Imperial College London, UK.

MERMAG consists of magnetometers on board MPO and MMO: MPO-MAG and MMO-MGF. Some measurements are only possible using the magnetometers on both spacecraft. MPO-MAG is a dual digital fluxgate magnetometer, which shall be used to measure DC and low frequency perturbations of the magnetic field. 44)


Figure 34: Schematic of Mercury's magnetic field (image credit: MERMAG team)

The primary objective of the magnetic field investigation on MPO is to provide the magnetic field measurements that will lead to the detailed description of Mercury's planetary magnetic field, and thereby constrain models of the evolution and current state of the planetary interior. This objective will be achieved using accurate magnetic field measurements by MERMAG on MPO. It will be supported by measurements made on the MMO (Mercury Magnetospheric Orbiter), both to distinguish the effects of the Hermean magnetosphere on the MPO measurements and to use the MMO measurements directly to augment the database for the determination of the internal terms. With the data of MERMAG, it will also be possible to determine all the terms associated with the internal field up to the octopole with high accuracy as well as higher order terms, depending on the structure of the internal field.

The secondary objectives of MERMAG relate to the interaction of the solar wind with the Hermean magnetic field and the planet itself, the formation and dynamics of the magnetosphere as well as to the processes that control the interaction of the magnetosphere with the planet. In particular, measurements close to the planet will allow the determination of the conditions for access of the solar wind to the planetary surface and assessing the role and importance of different current systems, including subsurface induction currents and the conductivity of the regolith. These objectives will be greatly assisted by the planned close association with the magnetic field investigation on the MMO.

Measurement principle: MPO-MAG is designed as follows: Two identical magnetometers are used each with their own dedicated electronics. This two sensors technique will be applied in order to help determine the magnetic influence of the spacecraft. The instrument hardware comprises an electronics box, two sensor units with their associated thermal hardware and mechanical fixings, plus an electrical harness which connects the sensors to the electronics box. The sensors are mounted on a deployable boom, whilst the electronics box is located inside the spacecraft structure.

The boom is a critical subsystem both for the MPO-MAG instrument and the spacecraft. It enables both sensors to be slightly removed from the spacecraft; combining the signal from both the inboard and outboard sensors will help determine the magnetic interference from the spacecraft itself.


Figure 35: Schematic of boom mounting technique (image credit: MERMAG team)

The MPO-MAG instrument is largely autonomous in operation, requiring a minimum of commanding only for selecting from a set of science operations modes and corresponding telemetry bit rates. The two sensors measure the magnetic field with a sample rate of 128 Hz. These data will be reduced onboard to a lower temporal resolution depending on the instrument mode: 64, 32, 16, 8, 4, 2, 1, and 0.5 Hz.


MERTIS (Mercury Radiometer and Thermal Infrared Spectrometer)

The objective is to provide high spectral resolution data. MERTIS will return detailed information about the mineralogical composition of Mercury's surface layer. This is crucial for selecting a valid model for the origin and evolution of the planet. PI: Harald Hiesinger, University of Münster, Co-PI: Jorn Helbert, DLR, Germany. MERTIS is a joint project of the University of Münster, two institutes of the German Aerospace Center (DLR) in Berlin-Adlershof and several industrial partners and research institutes. 45) 46) 47)

MERTIS has four scientific goals: the study of Mercury's surface composition, identification of rock-forming minerals, mapping of the surface mineralogy, and the study of the surface temperature variations and thermal inertia. The instrument will provide detailed information about the mineralogical composition of Mercury's surface layer by measuring the spectral emittance in the spectral from 7-14 µm with a high spatial and spectral resolution. Furthermore MERTIS will obtain radiometric measurements in the spectral range from 7-40 µm to study the thermo-physical properties of the surface.

MERTIS is a pushbroom radiometer. The spectrometer employs an uncooled microbolometer array made from amorphous silicon, which yields a short thermal time constant as well as very low NETD (Noise Equivalent Temperature Difference). The array provides spectral separation and spatial resolution according to its two-dimensional shape. The operation concept principle is characterized by intermediate scanning of the planet surface and three different calibration targets: free space and on-board black body sources.

The general instrument architecture comprises two separate parts - the Sensor Head including optics, detector and proximity electronics and the Electronics Unit containing the power supply with an interface to the primary bus, the sensor control and the driving electronics. This highly integrated and low mass (3.3 kg) measurement system is completed by a motor driven Pointing Unit device which orients the optical path to the planet and the calibration targets.



Figure 36: Illustration of the MERTIS instrument configuration and its elements (image credit: MERTIS team)

On Mercury, the spectral radiance at day side shows that the thermal emission starts to dominate the radiance already at wavelengths larger than 1.2 µm (at 725 K) depending on the surface albedo. The range between 0.8 and 2.8 µm is a transition region characterized by the overlapping of the reflected solar radiation and the thermal emission. However, Mercury's thermal flux exceeds the flux reflected from its surface. This enables emittance spectroscopy in the thermal IR range where there is high potential for mineral identification because it is in this region where the major rock-forming minerals (e.g. feldspar) have their fundamental vibration bands.


Figure 37: MERTIS instrument overview (image credit: MERTIS team)

The optics design is is based on an all-reflective optics concept with an off-axis TMA (Three Mirror Anastigmatic) telescope behind an IR entrance window and a pointing mirror for the target selection (planet view or calibration views to deep space and two reference sources). The spectrometer is a derivative of an elegant relay disclosed by Offner in the early 1970s. This combination is free from spherical aberration, coma and distortion and, when the algebraic sum of the powers of the mirror reflecting surfaces utilized is zero, the image produced is free from third order astigmatism and field curvature.



Radiometer (µRAD)

Focal length

50 mm

F number


Microbolometer array
Illuminated pixels

160 x 120 @ 35 µm
100 spatial, 80 spectral
2 x 15 @ 250 µm

Spectral channel width

90 nm / pixel

Spectral resolution

78 – 156 λ/Δλ

Spectral range

7 – 14 µm

7-40 µm


0.95 x109 cm √Hz W-1

7 x 108 cm √Hz W-1

IFOV (Instantaneous Field of View)

0.7 mrad

5 mrad

GSD (Ground Sample Distance)
- Periherm 400 km
- Apoherm 1500 km

280 - 1400 m (M= 1- 5)
1050 m

2000 m
7500 m

FOV (Field of View)

4º ACT (Across Track), 0º ALT (Along Track)

4º ACT, 1º ALT

Swath width

28 km

Table 4: Key optics parameters of MERTIS


Figure 38: Optics configuration of MERTIS (image credit: DLR)

Radiometer: The basic concept of the radiometer channel is to place a 2 x 15 elements thermopile double line array sensor chip with integrated optical slit for the spectrometer at the focal plane of the TMA entrance optics. The small signal voltages of the order of µV to mV generated by the thermopile sensors are transmitted differentially via a starr-flex interface PCB (Printed Circuit Board) to a proximity electronics, where the signals are multiplexed, amplified, converted to digital units and transmitted to the MERTIS ICU (interface Control Unit).

The unusual solution of incorporating the optical slit into the radiometer chip was driven by the very small space available which is mainly a consequence of the MERTIS requirement to minimize the heat input through the entrance optics. This design requires a modification of the standard thermopile design where a self-supporting membrane containing the thermoelectrically active layers is spanned over a surrounding Si frame. Here, an additional central bridge is added to the Si frame which provides additional mechanical support (Figure 39), thereby allowing to cut the slit into the center of the membrane. Furthermore, the thermopile pixels are only weakly coupled to this central Si bridge by very narrow V-shaped bars to reach the maximum possible sensitivity.

Each thermopile pixel of 200 µm x 1100 µm size consists of 14 thermocouples connected in series using Bi0.87Sb0.13/Sb as thermoelectric materials. The pixels are coated by a thin layer of black silver smoke, a material with nearly constant high absorption from the visible to the far infrared. The pixel width of 200 µm is close to the diffraction limit of the optics at 40 µm wavelength, a small gap of 50 µm width between the pixels is necessary for technological reasons and also effectively eliminates thermal crosstalk between neighboring pixels (under vacuum).

One infrared 8-14 µm bandpass filter is mounted directly above the thermopile array on 50 µm high standoffs which are micro machined directly onto the chip. The other array uses only the MERTIS entrance filter to realize a 7-40 µm broadband IR channel which is aimed at measuring low object temperatures down to 100 K.


Figure 39: µRAD detector chip and part of I/F flexboard (without IR filter), image credit: MERTIS team

The electrical interface of the thermopile arrays is provided by direct wire bonding of the sensor signals to a starr-flex PCB, the ends of the two flex wires are equipped with a Nano-connector to feed the signals into the radiometer electronics. Chip and PCB are mounted inside a dedicated Aluminum housing comprising of a baseplate and a cover. The whole detector unit is then fixed to the TMA structure by Shapal spacers which are providing an excellent thermal coupling but electrical insulation.


Figure 40: Photo of the MERTIS flight model, the housing is ~180 x 180 x 130 mm in size, the planet baffle is ~200 mm long, space baffle on top (image credit: MERTIS team, Ref. 46)


MGNS (Mercury Gamma-ray and Neutron Spectrometer)

MGNS will determine the elemental compositions of the surface and subsurface of Mercury, and will identify the regional distribution of volatile depositions on the polar areas which are permanently shadowed from the Sun. PI: I. Mitrofanov, IKI (Institute for Space Research), Moscow, Russia. 48) 49)

Since Mercury lacks a thick atmosphere, its natural nuclear emissions can be detected from orbit, i.e., gamma rays arising from cosmic-ray interactions and those arising form the natural radioactive decay of K (Potassium), Th (Thorium) and U (Uranium). Mercury has a very weak magnetic field with cut-off rigidity near the equator of ~1 MeV. Therefore, the galactic cosmic rays are essentially unimpeded and interact directly with the shallow subsurface producing copious secondary neutrons within the first 1–2 m of the surface (Figure 41). These neutrons interact with the soil nuclei either by in-elastic scattering or capture reactions, producing secondary nuclear gamma rays. Each chemical element has a unique set of nuclear lines, so the data from a gamma-ray spectrometer in near-orbit can, in principle, uniquely identify the elemental composition of the Mercury shallow subsurface.


Figure 41: Galactic cosmic rays produce secondary neutrons which induce gamma-ray line emission from the surface of Mercury. Line emission also results from natural radioactive isotopes in the surface regolith (image credit: IKI)

The intensity of a gamma-ray line of a particular element depends on the spectrum and flux of secondary neutrons and so knowledge of the spectral density of neutrons is also a necessary prerequisite for the determination of the elemental abundance. The energy spectrum of leakage neutrons, in turn, depends on the elemental composition of the soil. A neutron with a mass m, loses a small fraction of energy -m/(M+m) in a collision with heavy nucleus of mass M. However, when m=M, the incident particle will lose half its energy, as is the case when a neutron collides with a hydrogen nucleus. Thus, it can be seen that the addition of even a little hydrogen into a soil will decrease the leakage flux of epithermal and high-energy neutrons while simultaneously increasing the flux of thermal neutrons.

The objective of MGNS is to observe neutron fluxes in wide energy range (from thermal to 10 MeV) and gamma-ray with high energy resolution (approximately 3.5% at the energy of 662 keV) in the energy range from 300 keV to 10 MeV during the interplanetary cruise phase and on the orbit around Mercury.

Physical characteristics of Mercury nuclear emission

Requirements for MGNS measurements

Detectors and initial data products of MGNS instrument

Flux of gamma-ray lines from the Mercury subsurface

To measure the set of the most intense gamma-ray lines, which characterize the content of soil-composing elements and natural radio-isotopes

Scintillation detector of gamma-rays SCD/G with the high spectral resolution and high efficiency for gamma-rays Data product is energy spectrum of counts for gamma-rays with 4096 linear channels at the energy range 0.3–10.0 MeV

Flux of thermal neutrons from the Mercury surface

To measure the flux of thermal neutrons below the threshold of 0.4 eV

Detectors SD1 and SD2 with 3He proportional counters, with and without Cd shielding, respectively Data product is the time profile of counts for thermal neutrons, which is determined, as difference of counts from SD2 and SD1

Flux of epithermal neutrons from the Mercury surface

To measure the flux of epithermal neutrons in two energy ranges 0.4 eV–1 keV and 0.4 eV–500 keV

Detector SD2 with 3He proportional counter and with Cd shielding for energy range 0.4 eV–1 keV. Detector MD with 3He counter and polyethylene moderator inside Cd shield for energy range 0.4 eV–500 keV. Data products are two time profiles of counts for epithermal neutrons from SD2 and MD

Flux of high-energy neutrons from the Mercury surface

To measure the flux of high-energy neutrons in the energy range 0.3–10.0 MeV

Sthylbene scintillator SCD/N within a anti-coincidence plastic scintillator APC Data product is the energy spectrum of counts for high-energy neutrons with 16 energy channels for energy range 0.3–10.0 MeV

Table 5: Measurements, detectors and initial data products of MGNS

The MGNS has one detector (SCD/G) for gamma rays and four detectors (SD1, SD2, MD and (SCD/N) for neutrons. Additionally, the high-energy neutron detector (SCD/N) is surrounded by the APS (Anti-coincidence Plastic Scintillator), to protect the sensitive volume of SCD/N from external charged particles. All these detectors are integrated into a single module, which also contains the electronic boards for analog signal processing, HV and LV provision, data storage, logic and interface support.

MGNS instrument: The MGNS design is based on the heritage of HEND (High Energy Neutron Detector (HEND) flown on NASA's "Mars Odyssey" mission. The HEND has successfully operated for more than 7 years in space and has returned more that 3 GByte of scientific data. The concept of the MGNS design and its sensors are shown in Figures 42 (a)–(d). A schematic of the MGNS electronics is presented in Figure 43. It consists of two detection segments: the MGRS (Mercury Gamma-Ray Spectrometer) and the MNS (Mercury Neutron Spectrometer) supported by the DLS (Digital and Logic Segment), which is based on an FPGA. The overall dimensions of the MGNS correspond to 257 x 342 x 140 mm3.

All three detectors SD1, SD2 and MD [Figure 42 (b)] have identical 3He proportional counters and analog electronics. They are based on the HEND prototype elements with counter LND 2517 having a diameter of 12.7 mm, a length of 94 mm and a pressure of 6 atmospheres. The digitalization of counts allows the project to record the well-known two-peak energy spectrum from 3H and p. The energy peak at 764 keV corresponds to the full energy deposition of both particles, the low-energy peak at 191 keV corresponds to the energy of 3H only, when a proton escapes from the detection volume. If necessary, one may reject the contribution of low-amplitude noise by a commendable lower-energy threshold. The front-end read-out electronics for SD1, SD2 and MD are quite simple and identical for all three sensors.

The MGNS has two scintillation sensors, SC/N and SC/G [Figure 42 (c) and (d)]. The sensor SC/N has a stilbene scintillator. It is also based on the HEND heritage. Recoil protons have randomly distributed energies from 0 up to the total energy of the neutron, En, and produce a scintillation flash in the stilbene. The light is easily detected for energetic protons above energy of -300 keV. The low-energy cut-off of the sensor SC/N is determined by this threshold. The high-energy cut-off is governed by the decreasing cross section of the recoil reaction with increasing neutron energy. We use a cylindrical stylbene crystal of size Ø30×40 cm for detector SC/N. The efficiency curve for neutron detection by the SC/N has a maximum of about a few cm2 around 2.0 MeV.

Electrons, either external, or produced by gamma rays, also generate scintillation light in the stylbene crystal, as well as protons. However, the time profiles of the scintillation flash are quite different for electrons and protons, and a special analog board of the MNS segment separates counts into these two categories. It has a high accuracy for separating electrons and protons with a mis-identification of only 1 case in 2000. External cosmic ray protons have also to be separated from recoil protons. Similarly to the design of HEND, a plastic scintillator surrounds the stilbene crystal for the rejection of external protons. An event in the plastic is used for the generation of a veto signal for rejecting cosmic ray events in the stilbene.


Figure 42: Schematic views of the MGNS (image credit: IKI)

The gamma-ray spectrometer of the MGRS [Figure 42 (d)] is based on a LaBr3 scintillation crystal with a size of about 8 cm (both in diameter and length). For a spectral resolution of 3% at 662 keV, one would like to have about 8 energy channels over the Gaussian profile of the spectral line. Therefore, events from this sensor are converted into an energy spectrum with 4096 linear channels over the energy range 300 keV–10.0 MeV.

The architecture of the DLS (Digital and Logic Segment) is based on a radiation-resistant Actel FPGA (Figure 43). The logic of the FPGA is developed in accordance with project requirements and specifications. Low- and high-voltage supply units are operated by the DLS with the possibility of changing the levels of HV by commands. The DLS unit also provides the interface with the spacecraft systems for power, thermal control, data readout and commanding.


Figure 43: The main elements and units of the MGNS (image credit: IKI)


MIXS (Mercury Imaging X-ray Spectrometer)

MIXS will use X-ray fluorescence analysis to produce a global map of Mercury's surface atomic composition at high spatial resolution. This technique has also been used by the D-CIXS instrument on ESA's Smart-1 mission to the Moon. PI: Emma Bunce, University of Leicester SRC (Space Research Center), Leicester, UK. MIXS is funded in part by the UKSA (UK Space Agency) and will use novel X-ray optics to determine small-scale features on Mercury in order to find out what the planet's surface is made of. It will do this by measuring fluorescent X-rays that come from the planet's surface, excited by high energy X-rays from the Sun, to identify chemical elements. -The MIXS contribution in Germany is funded by the Max-Planck-Society. MPS (Max Planck Institute for Solar System Research) is responsible for the scientific calibration of the detector unit and MPE is providing the whole detector flight hardware.

The MIXS instrument is designed to determine the surface composition of the planet by means of fluorescent X-ray spectroscopy. MIXS is the essential monitor of the solar X-ray flux which, along with solar wind protons, excites the characteristic K-series lines of the elements present in surface material. MIXS is also concerned with the interaction of the solar wind with Mercury's magnetosphere and exosphere which is expected to produce intensive emission of continuum X-rays. 50) 51) 52)

The MIXS-T channel of the instrument features the first true imaging X-ray telescope to be used in planetary science – realized using MCP (Microchannel Plate) optics manufactured by Photonis SAS (Brive, France). MIXS-T will have an angular resolution of 2-4 arcmin, sufficient to resolve surface features, dependent on solar flare state, smaller than ~20 km across. A second, conventional non-imaging collimated channel is provided by MIXS-C with superior grasp to MIXS-T, but no spatial resolution on scales less than the field-of-view. The detectors for both channels are 64 x 64 pixels in a macropixel DEPFET (Depleted P- Channel Field Effect Transistor) active pixel sensor geometry (Max Planck Semiconductor Laboratory, Munich). The energy resolution of these devices is expected to be sufficient to resolve the K series lines of all the elements of interest. An excellent low-energy detector response opens up the possibility of measuring Fe abundance from the L-alpha line at 0.7 keV.


Figure 44: Photo of the MIXS FM (Flight Model) instrument (image credit: University of Leicester) 53)


MORE (Mercury Orbiter Radio science Experiment)

MORE will help to determine the gravity field of Mercury as well as the size and physical state of its core. It will provide crucial experimental constraints to models of the planet's internal structure and will test theories of gravity with unprecedented accuracy. Its data will be used in conjunction with those from BELA and ISA to achieve these goals. PI: L. Iess, University of Rome ‘La Sapienza', Italy. 54) 55)

The MORE hardware, a transponder enabling high accuracy range and range rate measurements, has been designed to determine the planet's gravity field to degree and order 25 and carry out improved tests of relativistic gravity. The measurements undertaken by MORE address three different areas:

- reconstruction of the planet's gravity field, the coefficients of the spherical harmonics expansion up to degree and order 25, and the Love number k2 (gravimetry experiment). The expected accuracies range from signal to noise ratio of 104 for degree 2, to SNR of 10 for degree 20.

- estimation of the rotational state of Mercury by means of obliquity and amplitude of the physical librations in longitude (rotation experiment). These measurements, carried out in collaboration with the camera team, provide the moments of inertia of the whole planet and its mantle.

- determination of the post-Newtonian parameters, the mass and the oblateness of the Sun, and the upper limits to the temporal variation of the gravitational constant G (relativity experiment).

The crucial onboard elements of the MORE experiments are:

- the Ka/Ka transponder (KaT), provided by ASI (Italian Space Agency)

- the TLC/TCM deep space transponder (DST)

- the 1.2 m high gain antenna (HGA)

- the high sensitivity ISA (Italian Spring Accelerometer).

In addition, the final global fit and orbit reconstruction will incorporate also laser-altimetric and optical observables provided by the onboard laser altimeter (BELA) and high-resolution camera (SIMBIO-SYS). The KaT and the DST are particularly important because they enable a multi-frequency radio link at X-band (7.2 GHz uplink/8.4 GHz downlink) and Ka-band (34/32.5 GHz), a configuration already exploited by the Cassini mission. Thanks to this configuration, range rate measurements will attain accuracies of 3 µm/s (at 1000 s integration time) at nearly all solar elongation angles. In the geometric optics limit, the use of multiple frequencies allows a complete cancellation of plasma noise, the dominant noise source in the S- and X-band radio links. A novel WBRS (Wideband Ranging System), based upon a pseudo-noise modulation scheme at 24 Mcps, will provide observables accurate to 20 cm (two-way).

The effects of non-gravitational accelerations on the spacecraft dynamics (quite large in the harsh hermean environment) will be removed to a large extent thanks to the ISA accelerometer. These instrument readouts will be sent to ground in the telemetry stream and referenced to the phase center of the high gain antenna. The orbit determination code will then use a smoothed version of the accelerometer measurements to integrate the equation of motion, effectively realizing a software version of a drag-free system.

Such a complex experimental setup, implemented for the first time in a planetary mission, will be used not only for the reconstruction of the Mercury's gravity field, but also for a precise reconstruction of the spacecraft orbit. Accuracies of 0.1-1 m in the radial position seem attainable. The position of the MPO in the hermean frame (whose origin is defined by zeroing the dipole terms in the harmonic expansion of the gravitational potential) will be used for the appropriate referencing of the laser altimetric measurements and the images from the high-resolution camera. The combination of altimetric and gravity measurements will provide the topographic heights, a crucial information to determine the structure of Mercury's crust and outer mantle.

The along- and across-track position of the spacecraft is crucial for the rotation experiment, aiming to determine the rotational state of the planet by means of optical tracking of surface landmark. The pole position and physical librations in longitude will be obtained from a precise geo-referencing of high-resolution images (5 m pixel size at pericenter). The final accuracy of this experiment rests not only upon an accurate knowledge of the spacecraft position, but also on the quality of the attitude reconstruction. The onboard star trackers and gyroscopes should allow an accuracy of 1-2 arcsec. In addition, the spacecraft design ensures a high stability of the optical alignment between the star trackers and the camera.

Although MORE will make use of laser altimetric and optical images to stabilize the global orbital solution, the crucial observable quantities are range and range rate. These quantities are generated at the ground station after establishing a two-way coherent link. The core element of the tracking system is the reference oscillator, a H-maser with a frequency stability of one part in 1015 over time scales of 1000 s. The orbital solution is obtained from the observable quantities by means of a least squares fit, where the state vector of the spacecraft and the parameters of the dynamical model are jointly derived.

Numerical simulations of the MORE show that a batch-sequential estimation is fully adequate to reach the mission requirements in orbit determination and gravity field reconstruction. The batch-sequential filter provides the advantages of a batch method (i.e. a-posteriori data processing and superior stability), while adding sequential updates of the dynamical model. In our approach, observation data are processed in three steps:

• Batch-sequential estimation (initialization of local and global parameters)

• Multi-arc estimation (improvement of the global parameters)

• Single arc estimation (trajectory reconstruction).

Position errors are consistently found below 10 m in the along and cross-track components, while much better accuracies are obtained for the radial component. The orbital reconstruction is therefore fully adequate to support the laser altimetric observations (accurate to 1 m in the nadir direction). In the remaining two components (along- and across-track) the orbit determination accuracies are significantly larger, but still compatible with the requirements of the libration experiment (2 arcsec for Mercury's librations in longitude). An improvement in the orbit determination is necessary for a measurement of physical librations below the arcsecond level.

Better results are expected if additional observations become available. In the current planning, BepiColombo's MPO will be tracked by two stations, namely ESA's 34 m antenna in Cebreros, supporting mission operations, and NASA's DSN (Deep Space Network) antenna DSS 25 in Goldstone (California) for the radio science experiment. The X-band Doppler data, acquired at the Cebreros antenna, may prove valuable for the estimation of the ΔVs associated to desaturation maneuvers, a major source of uncertainty in the orbital reconstruction.


PHEBUS (Probing of Hermean Exosphere by Ultraviolet Spectroscopy)

The PHEBUS spectrometer is devoted to the characterization of Mercury's exosphere composition and dynamics. It will also search for surface ice layers in permanently shadowed regions of high-latitude craters. PI: Eric Quemerais, LATMOS-IPSL, France. The PHEBUS consortium composed of three main partners: Tohoku University (Japan) will provide the detectors and the main entrance mirror, IKI (Russia) will implement the scanning system, and SA/IPSL (France) will take in charge the design, assembly/ test/ integration, and will also provide three small detectors (zero order monitor, Ca- and K-channels). Finally, a small optional spectro-imager (0.3 kg), implemented under the responsibility of Southwest Research Institute (USA), is proposed for UV mapping purpose.

PHEBUS is a double spectrometer for the EUV (Extreme Ultraviolet) range (55-155 nm) and the FUV (Far Ultraviolet) range (145-315 nm) dedicated to the characterization of Mercury's exosphere composition and dynamics, and surface-exosphere connections. The optical configuration of PHEBUS can be divided into two independent parts. The collecting part consists of a straylight rejection baffle, an off axis parabolic mirror and an entrance slit, allowing to scan Mercury's exosphere thanks to a rotating mechanism. This movable mirror collects the light from the exosphere above thelimb and focuses it to the entrance slit. The parameters of the mirror were calculated so as to have a 170 mm effective focal length and a folding angle of 100º. This part determines the FOV (Field of View) and the LOS (Light of Sight). The spectrometer part determines the spectral resolution of the instrument and is composed by the entrance slit, two holographic gratings and the detectors. 56) 57) 58)

The spectrum detection is based on the photon counting method and is done using MCP (Micro-Channel Plate) detectors with RAE (Resistive Anode Encoder). Photocathodes are CsI for the EUV range, and CsTe for the FUV range. The size of the detectors active area is 40 x 25 mm2 equivalent to a matrix of 1024 x 512 virtual pixels (spectral x spatial). Furthermore Calcium and Potassium lines are selected by the FUV grating. These extra visible lines are monitored using PM (Photomultipliers) with bialkali photocathode also used in photon counting mode.

The main advantage of the MCP+RAE detectors is their very high sensitivity mainly due to a very low dark current. Thus photon counting is easily achievable on typical experiment temperature range (from -20ºC to+40ºC), avoiding mass and power expensive devices to cool the detectors. Five to six orders of magnitude for the detection are then a typical value and offer the monitoring of a wide range of emission.

Optical specifications: The wavelength ranges are 55-155 nm for the EUV and 145-315 nm for the FUV. The spectral resolution is defined in terms of FWHM (Full Width at Half Maximum) and Full Width at 1% of maximum (FW1%). The required spectral resolution is 1 nm for EUV and 1.5 nm for FUV. These values are to be compared with the result of the optical design optimization: the FWHM is about 0.5 nm on EUV, and 0.8 nm on FUV. Furthermore, the FW1% is about 0.9 nm on EUV, and 1.5 nm on FUV. These calculated values do not include any spreading effects due to scattering by gratings.


Figure 45: a: PHEBUS instrument, b: Optical configuration 1 – Entrance pupil; 2 – Entrance mirror; 3 – Slit; 4 – FUV grating; 5 – EUV grating; 6 – FUV detector; 7 – EUV detector (image credit: LATMOS/CNRS)


SERENA (Search for Exosphere Refilling and Emitted Neutral Abundance)

SERENA will study the gaseous interaction between Mercury's surface, exosphere, magnetosphere and the solar wind. PI: S. Orsini, CNR/IFSI (Institute of Space Astrophysics and Planetology), Rome, Italy. The instrument has been developed by an international team: SwRI (San Antonio, TX, USA); Space Research Institute (Austrian Academy of Sciences, Graz, Austria); IRF (Swedish Institute of Space Physics), Kiruna, Sweden; Physicalisches Insitut (Space Research & Planetary Sciences, University, Bern, Switzerland); FMI (Finnish Meteorological Institute), Helsinki, Finland; MPS (Max Planck Institute for Solar System Research), Kathlenburg-Lindau, Germany.

The SERENA instrument suite will study in-situ the composition, the vertical structure and the source of the deposit processes of the exosphere of Mercury. Such an environment is composed by thermal and directional neutral atoms (exosphere) originating via surface release and charge-exchange processes, and by ionized particles originated through photo-ionization and again by surface release processes. In order to accomplish the scientific goals, in-situ analysis of the environmental elements is necessary, and for such a purpose the SERENA instrument shall include four units: 59) 60) 61)

• two Neutral Particle Analyzers:ELENA and STROFIO

• two Ion Spectrometers (MIPA and PICAM).

In particular, ELENA investigates the neutral gas escaping from the surface of Mercury, and the related involved processes; STROFIO investigates the exospheric gas composition; PICAM investigates the exo-ionosphere extension and composition, and the close-to-planet magnetospheric dynamics; MIPA investigates the plasma precipitation toward the surface and ions energized and transported throughout the environment of Mercury.

The main scientific goals of SERENA are:

- Chemical and elemental composition of the exosphere

- Exo-ionosphere composition and distribution

- Surface emission rate and release processes

- Plasma precipitation rate

- Particle loss rate from Mercury's environment

- Gas density profile asymmetries


ELENA (Emitted Low-Energy Neutral Atoms)

The ELENA instrument within the SERENA suite is a new kind of low energetic neutral atoms instrument, devoted to detect neutral atoms from E ~20 eV up to E~5 keV, within a nadir pointing 1-D FOV (perpendicular to the spacecraft orbital plane).

The main ELENA scientific objectives are as follows: 62)

- Surface emission rate and release processes

- Particle loss rate from Mercury's environment

- Remote sensing of the surface properties

- ENA imaging applications for comparative solar-planetary relationship.

ELENA is a TOF (Time-of-Flight) sensor, based on the state-of-the- art of ultra-sonic oscillating shutter (operated at frequencies above 20 kHz and up to 50 kHz), mechanical gratings and MCP (Micro-Channel Plate) detectors. The purpose of the shuttering system is to digitize space and time when tagging the incoming particles without introducing "disturbing" detector elements, which may affect the particle's trajectory or the energy. This is particularly important in this case, in which neutrals of energies of a few tens of eVs must be detected.

Energy range

<0.02-5 keV (mass dependent)

Velocity resolution (ΔV/V)

Down to 10%

Viewing angle

4.5º x 76º

Angular resolution

4.5º x 4.5º (actual), 4.5º x 2.4º (nominal pixel)

Mass resolution M/ΔM

H and heavy species

Optimal temporal resolution

40 s

Geometric factor G

2 x 10-5 cm2 sr

Integral geometric factor

6 x 10-4 cm2 sr

Table 6: Performance parameters of ELENA

The ELENA sensor concept is showed in the Figure 46. The entrance of the start section (an aperture of about 1 cm2) allows the impinging neutral particles to enter through the shuttering system with a definite timing. Particles are then flown into a TOF chamber, and finally detected by a 1-dimensional array composed by MCP and a discrete set of anodes corresponding to a FOV of 4.5° x 76°, allowing the reconstruction of both, velocity and direction of the incoming particle events. The spacecraft footprint track will provide the second dimension.


Figure 46: Photo of the ELENA EQM instrument (image credit: CNT/IF SI)

The composite radiation made by neutrals, ions and light impinges onto the ELENA sensor entrance. A grid system placed between the main ELENA entrance and the shutter minimizes the ion and electron background radiation before the shuttering system. The internal ion deflector is a stack of particle cross-track grids which introduce a transversal E-field able to filter out the bulk of the charge particles of both signs.


STROFIO (Start from a Rotating FIeld mass spectrOmeter)

The STROFIO instrument within the SERENA suite is a mass spectrograph that determines particle mass-per-charge (m/q) by the TOF technique. The name comes from the Greek word "Strofi", which means "to rotate": the phase of a rotating electric field "stamps" a start time on the particles' trajectory, and the detector records the stop time.

STROFIO is the only instrument that can detect all neutral species reaching the spacecraft, not just those with emissions or absorptions at specific wavelengths. In contrast with optical techniques, STROFIO can make measurements with high sensitivity, both in full sunlight and in Mercury's shadow. The main science objectives of the STROFIO instrument are the following:

- Chemical and elemental composition of the exosphere and its variability

- Neutral gas density asymmetries

- Surface release processes

- Relationship between the exospheric composition and the surface composition.

STROFIO is characterized by a high-sensitivity (0.14 counts/s when the density is 1 particle/cm3). The achieved mass resolution is m/Dm = 60. STROFIO is a novel type of mass spectrometer: the start time is imprinted on the trajectory of the particle by a radio frequency electric field, that bends the trajectory in a given plane, and the stop time is the time when the particle reaches the detector.

In particular, the neutral particles enter into the ionization chamber through the entrance in the ram direction (see the figure). The neutral gas is ionized and accelerated into the mass analyzer. Here the ions experience the effects of an electric field, constant in magnitude, but with direction rotating uniformly in space, in a plane perpendicular to the initial ion velocity, at a frequency f. The trajectory of an ion can hit the detector only if the field points to the detector while the ion traverses the dispersing region. At other times, the ion will simply miss the detector. The time difference between the instant when the particle arrives at the detector and the time when the field was pointing in the appropriate direction is equal to the travel time through the field free region. In the table the major STROFIO characteristics are summarized.


Figure 47: Illustration of the STROFIO instrument (image credit: CNR/IFSI)

Energy range

< 1 eV

Viewing angle


Mass resolution M/ΔM


Mass range

1-64 dalton amu (atomic mass unit)


0.14 (counts/s / (particles/cm3)

Temporal resolution

10 s

Table 7: Performance parameters of STROFIO

MIPA (Miniature Ion Precipitation Analyzer)

MIPA is a simple ion mass analyzer optimized to provide monitoring of the precipitating ions using as little spacecraft resources as possible. The analyzer is capable to measure all main groups of ions present in the magnetosphere. The energy range and mass range of the analyzer is optimized to cover accelerated ionospheric ions. The main science objectives of MIPA are the following: 63)

- Plasma precipitation rate

- Planetary response to SW (Solar Wind) variations

- Magnetosphere structure and dynamics.

The ion flux arrival angle and energy are analyzed by an electrostatic deflector, comprising of two 90° cylindrical electrodes, followed by an 128° double focusing cylinder electrostatic analyzer. The ions exiting the energy analyzer are post accelerated up to 1 keV energy by a voltage applied to a TOF cell. Inside the cell, ions hit START and STOP surfaces producing secondary electrons recorded by two ceramic channel electron multipliers giving respective timing. For energies above 4 keV, the post acceleration is switched off. The timing of the event gives the ion velocity and, in combination with known energy, the mass. The Figure shows the MIPA principle elements and the overall view. The MIPA total G-factor can be controlled/decreased by decreasing the post acceleration voltage resulting in lowering the impact energy and thus secondary electron yield from the START and STOP surfaces, and reducing the size of the aperture slits. The table summarizes the MIPA major characteristics.

Energy range

15 eV - 15 keV

Energy resolution ΔE/E


FOV (Field of View)

90º x 360º, 4 polar x 6 azimuth pixels

Angular resolution FWHM (Full Width Half Maximum)

22.5º x 60º (polar x azimuth)

Mass range

1-50 amu (atomic mass unit)

Mass resolution

M/ΔM ~5

Sampling time

7.8125 ms

Time resolution

18 s, 4 polar x 6 azimuth x 96 energy steps


1-10% (adjustable to decrease geometrical factor)

Geometrical factor

~10-5 cm2 sr eV / eV per pixel, w/efficiency

Table 8: Performance parameters of MIPA


Figure 48: Photo of the MIPA instrument (image credit: CNR/IFSI)


PICAM (Planetary Ion CAMera)

PICAM operates as an all-sky camera for charged particles, allowing the determination of the 3D velocity distribution and mass spectrum for ions from thermal up to ~3 keV energies and in a mass range extending up to ~ 132 amu (Xenon). The instantaneous 2p FOV coupled with this mass range and a mass resolution better than ~ 100 is a unique capability, which provides to PICAM superior performances in the frame of the MPO mission. The major PICAM scientific objectives are the following: 64)

- Exo-ionosphere composition and distribution

- Planetary response to SW variations

- Magnetosphere structure and dynamics.

Figure 49shows a general layout of the sensor. The ion optics is based on the principle of a modified pinhole camera. The sensor is symmetric along the Z-axis and its FOV is a hemisphere centered along this axis. Ions enter through an annular slit (figure). After reflection on an ellipsoidal ion mirror (2) the 90° polar angle distribution is folded into a 15° angular range. Here the ions pass a modulated wire gate which defines discrete packets of ions for analysis of the time-of-flight until the particles impact on the MCP. The modulation can be either single-shot or with a pseudo-random sequence which results in higher efficiency. Particles pass through a narrow, toroidal analyzer and through an exit slit and are reflected by a planar mirror. After energy selection in the toroidal analyzer, they enter the TOF and imaging section. A cross section of the ion optics is shown in the figure. UV rejection will be obtained by a striated primary mirror covered by a non-reflecting layer of Cu2S, which decreases the UV reflection by a factor of 1000. Multiple reflections within the instrument, the small entrance slit and the narrow exit slit in front of the mass analyzer provide very strong protection. The outer part of the ion optics is designed for hot conditions. The lower part of the sensor containing the MCPs and the detector electronics is thermally decoupled. The table summarizes the PICAM major characteristics.


Figure 49: Photo of the PICAM device (image credit: CNR/IFSI)

Energy range, energy resolution ΔE/E

1 eV -3 keV, 7%

Viewing angle

3-D, 1.5p

Angular resolution


Mass resolution M/ΔM


Mass range

1 -~132 amu (Xe)

Time resolution

1-32 s

Geometric factor G = SW (Solar Wind)

2.3 x 10-2 cm2 sr

Geometric factor ΔE/E

1.6 x 10-3 cm2 sr eV/eV

Table 9: Performance parameters of PICAM


SYMBIO-SYS (Spectrometers and Imagers for MPO BepiColombo Integrated Observatory System)

SYMBIO-SYS, also written as SIMBIO-SYS, will examine (also in stereo and color) Mercury's surface geology, volcanism, global tectonics, surface age and composition, and geophysics. PI: E. Flamini, ASI, Italy. SIMBIO-SYS, is an international project led by Italy, with main cooperation from France and Switzerland. The STC instrument is a collaboration of CNR-IFN UOS Padova,CISAS University of Padova,, INAF -Osservatorio Astronomico di Padova, Selex-Galileo, University of Padova, and ASI. 65) 66) 67) 68)

SIMBIO-SYS incorporates capabilities to perform 50 - 200 m spatial resolution global mapping in both stereo mode and color imaging, high spatial resolution imaging (5 m/px scale factor at periherm) in panchromatic and broad-band filters, and imaging spectroscopy in the 400 - 2200 nm spectral range. This global performance is reached using three independent channels:

• STC (STereoscopic imaging Channel)

• HRIC (High Resolution Imaging Channel)

• VIHI (Visible and near-Infrared Hyperspectral Imager).


Figure 50: a) SYMBIO-SYS STC stereo configuration; b) position and size of useful filter strips images on the detector area (in black), image credit: SYMBIO-SYS team)

STC (STereoscopic imaging Channel) optical design:

STC is a double wide angle camera designed to image each portion of the Mercury surface from two different perspectives, providing panchromatic stereo image pairs required for reconstructing the DTM (Digital Terrain Model) of the planet's surface. In addition, it has the capability of imaging some portion of the planet in four different spectral bands (Figure 50).

The STC design is composed of two "sub-channels" looking at the desired stereo angles, that share the majority of the optical elements and the detector (Figure 51). With respect to classical two- or single-camera designs, this solution allows to reach good stereo performance with general compactness, saving of mass, volume and power resources.

In general, stereo cameras adopt a pushbroom acquisition mode: the detector is a linear array and the full bidimensional image is reconstructed placing side by side each of the lines successively acquired at a suitable rate determined by the spacecraft velocity. For STC, instead, a push-frame mode has been chosen; in this case the detector is a CMOS APS (Active Pixel Sensor) bidimensional array, so actual 2D images of the planet surface are acquired, then buffered and read while the spacecraft moves. Only when the image on the detector has shifted along track by an amount corresponding to the FoV of each filter, another image is acquired.

The selected APS device has the snapshot option, that is substantially an electronic shutter; for this reason, no mechanical shutter has been foreseen for this instrument. The push-frame acquisition method allows to have some overlap of the imaged regions in the along-track direction, increasing the image matching accuracy and taking into account possible small drifts of the satellite pointing.

STC scientific requirements

Scale factor

50 m/px at periherm


40 km at periherm

Stereoscopic properties

± 21.4° stereo angle with respect to nadir; both images on the same detector

Vertical accuracy

80 m

EE (Encircled Energy)
MTF (Modulation Transfer Function)

> 70% inside 1 pixel
> 60% at Nyquist frequency

Wavelength coverage

410-930 nm (5 filters)


Panchromatic (700±100 nm), 420±10 nm, 550±10 nm, 750±10 nm, 920±10 nm

STC optical characteristics

Optical concept

Catadrioptic: modified Schmidt telescope plus folding mirrors fore-optics

Stereo solution (concept)

2 identical optical channels; detector and most of the optical elements common to both channels

Focal length (on-axis)

95 mm

Pupil size (diameter)

15 mm

Focal ratio


Mean image scale

21.7 arcsec/pixel (105 µrad/pixel)

FOV (cross-track)


FOV (along-track)

2.4º panchromatic, 0.4º color filters


Si_PIN (format: 2048 x 2048; 10 µm squared pixel); 14 bit dynamic range

Table 10: STC scientific requirements and STC optical characteristics


Figure 51: The entire STC optical design is shown: in (a) the configuration is viewed in the plane defined by the along-track and nadir directions; in (b) the projection in the orthogonal plane, the one including across-track and nadir directions, is given. In the inset, an enlarged view of the focal plane region helps to better follow the rays which are focalized on the APS detector (image credit: SYMBIO-SYS team)


HRIC (High Resolution Imaging Channel): 69) 70)

The accommodation of the HRIC channel in the SYMBIO-SYS experiment is shown in Figure 52. The HRIC optical design has been developed at INAF Astronomical Observatory of Capodimonte, with the main objective of characterizing relevant Mercury surface features at very high spatial resolution (pixel scale of about 5 m at 400 km from planet surface) in the visible range. The high resolution images of selected regions will allow the project to identify key surface features (e.g., craters, scarps, lava flows and plains) and to study their relation with internal processes, as well as the effect of external agents, such as meteor bombardment. In addition, HRIC images will be of paramount importance in support of experiments aiming to the identification of Mercury orbital parameters, such as the obliquity and the amplitude of the libration.

The HRIC optical design has been optimized in order to satisfy not only scientific requirements and nominal optical performance, but also dimensional requirements and mechanics constraints of compactness and low mass required for space applications. The optical design is based on a catadioptric concept, with optimization of a Ritchey-Chretien configuration by a dedicated corrector (Figure 53). The instrument has a focal length of 800 mm and is equipped with a dedicated refractive camera, in order to correct the field of view covered by a detector of 2 k x 2 k pixels, with a pixel size of 10 µm. The focal ratio of the instrument is f/8, in order to be diffraction limited at 400 nm and to optimize radiometric flux and overall mechanical dimensions. The main HRIC optical characteristics are reported in Table 12 . The combined (reflective + refractive) solution guarantees a good balance of achieved optical performances and optimization of resources (mainly volume and mass). The curvature of the lenses is spherical, in order to simplify manufacturing.

The quality of the optical design has been optimized and checked by analysis with the Zemax code. The adopted configuration corrects and transmits well over the whole band of observation (400-900 nm). A relative obscuration ratio of 0.3 (between primary and secondary mirror diameters) has been achieved in order to provide a good energy transfer to the telescope exit pupil. The optical design presented in this paper takes also into account the foreseen filters and detector package, in which the detector window acts also as filter.


Figure 52: SYMBIO-SYS instrument including the three integrated channels: HRIC (High Resolution Imaging Channel), STC (STereoscopic imaging Channel) and VIHI (Visual and near-Infrared Hyperspectral Imager), image credit: SYMBIO-SYS team 71)

The main tasks of the HRIC are to provide high resolution images of selected Mercury surface features like craters, scarps, lava flows and plains with a panchromatic filter and to help in geo-mineralogical characterization of local surface features by band-pass filters.

Pixel scale

5 m/pixel at periherm (400 km from Mercury surface)

Pixel resolution

12.5 rad/pixel (with pixel of 10 µm)

Spectral range

400-900 nm

Image quality

Diffraction limited at 400 nm


1 panchromatic (650 nm central wavelength, 500 nm bandwidth),
3 band-pass (550 nm, 700 nm, 880 nm, 40 nm bandwidth)

Table 11: HRIC main scientific requirements


Figure 53: HRIC optical layout (SYMBIO-SYS team)

Optical configuration

Catadioptric: Ritchey-Chretien modified with a dedicated corrector


100 mm

Angular FOV (Field of View)




Focal length

800 mm

Image scale

12.5 µrad/pixel

Focal plane detector

2048 x 2048 (CMOS APS)

Detector pixel size

10 µm x 10 µm

Table 12: HRIC optical parameters

Since the system is diffraction limited, the fraction of diffraction EE (Encircled Energy) curves enclosed in one pixel and the diffraction MTF (Modulation Transfer Function) until the Nyquist frequency have been considered in order to evaluate the image quality. For both diffraction EE and MTF, central obscuration has been included. In addition, the RMS spot radius on the image plane and field curvature and distortion have been evaluated. The field corrector has been optimized in order that images in different filter bands can be compared without distortion and field curvature. Thus, the image quality is high over the entire field of view. The main optical performances are reported in Table 13.

Polychromatic diffraction EE in one pixel


Polychromatic MTF at Nyquist frequency


Polychromatic RMS spot diameter (geometric)

0.8 µm

Field curvature

12 µm



Table 13: HRIC main optical performances


VIHI (Visible and near-Infrared Hyperspectral Imager)

VIHI is one of the three optical channels of the SYMBIO-SYS instrument suite for the BepiColombo mission to Mercury. Its scientific objective is to study the hermean surfaces composition by sensing the photon flux reflected off the planet. VIHI works in the spectral range of 400 -2000 nm with 256 spectral channels (6.25 nm/band sampling). The particularity of this channel is the use of a single detector matrix (264 x 264) for both visible and infrared wavelengths.

The instrument has an instrument FOV (Field of View) of 250 µrad corresponding to a spatial scale of about 100 m/pixel at periherm and 375 m at apoherm. The instrument operates in pushbroom configuration, sampling the surface of Mercury with an FOV of 64 x 0.25 mrad. The main technical challenges of this experiment are the focal-plane design (cadmium-mercury-telluride thinned to improve the efficiency at visible wavelengths), the short dwell time (from about 40 ms at the equator to about 100 ms at the poles), thermal control, mechanical miniaturization, radiation hardening, high data rate, and compression. 72)

The optical layout has been optimised for the minimum volume, with multiple folding of the optical path and adoption of a large number of off-axis elements, whose realisation and alignment required an intensive co-engineering efforts between optical and mechanical design, manufacturing, integration and verification.

All the optical elements of VIHI, including the FPA (Focal Plane Assembly), are mounted on a compact optical bench, with approximate size of 160 x 230 mm.

VIHI detector is a HgCdTe sensor hybridized over a specifically designed CMOS ROIC (Readout Integrated Circuit) of 256 x 256 useful pixels with 40 µm pitch. This was specified as the only possible detector geometry capable to provide the required performance in terms of spectral and spatial coverage and sampling. The detector input circuit is a CTIA (Capacitive Transimpedance Amplifier), the ROIC provides snapshot integration mode which allows read while integrate capability. The repetition time at the equator is 40 ms. The detector is integrated into an FPA which also hosts the TEC (Thermoelectric Cooler), the order sorting filter and a light trap for the spectrometer zero order. A thermal strap is mounted on the back of the FPA, in thermal contact with the hot side of the TEC, while the ROIC is fixed on the TEC cold side. This allows to minimize the mass to be cooled down by the TEC. Moreover, thermal conductance between the FPA and the optical bench is minimized by adoption of a thin interface structure with three legs and a frame around the package.

A shutter mechanism is included in the VIHI design, aimed at measuring the dark current of the detector, as well as the thermal background from the spectrometer, during daylight. This is periodically required during science operation to achieve the required radiometric accuracy. This mechanism consists of a rotating motor actuator, that allows to obscure the slit of the instrument by means of a small opaque cover blade, in such a way to block incoming radiation from the telescope. The mechanism design is based on a two position actuator (with return spring), that implements a ‘fail safe position' concept. When there is no power supply the motor is in the open position, so to allow science observations even in case of failure of the motor and/or the driving circuit. On the other hand, upon a dedicated telecommand, the motor is powered and the mechanism commutates on the closed position. This approach both avoids single point failure and minimizes the power consumption of the shutter, since the motor is powered only when it is in the closes position (i.e. during dark current acquisition). This is the only mechanism in SIMBIO-SYS.

A calibration unit is included in the VIHI design to allow spectral and radiometric performance check of the instrument in the whole spectral range. This is planned to be done every few weeks during the mission to monitor the performance of the optics and the detector against ageing effects, especially due to the harsh radiation environment at 0.3 AU.


Figure 54: VIHI Optical bench drawing (internal harness and stray-light shields are not shown), image credit: SYMBIO-SYS team


SIXS (Solar Intensity X-ray Spectrometer)

SIXS will perform continuous measurements of X-rays and particles of solar origin employing a very wide field of view. PI: J. Huovelin, Observatory University of Helsinki, Finland.

The objective of SIXS on MPO (Mercury Planetary Orbiter) is to investigate the direct solar X-rays, and energetic protons and electrons which pass the spacecraft on their way to the surface of Mercury. These measurements are vitally important for understanding quantitatively the processes that make Mercury's surface glow in X-rays, since all X-rays from Mercury are due to interactions of the surface with incoming highly energetic photons and space particles. The X-ray emission of Mercury's surface will be analyzed to understand its structure and composition. SIXS data will also be utilized for studies of the solar X-ray corona, flares, solar energetic particles, and the magnetosphere of Mercury, and for providing information on solar eruptions to other BepiColombo instruments. 73) 74)

SIXS consists of two detector subsystems. The X-ray detector system includes three identical GaAs PIN detectors which measure the solar spectrum at 1–20 keV energy range, and their combined field-of-view covers ~1/4 of the whole sky. The particle detector system consists of an assembly including a cubic central CsI(Tl) scintillator detector with five of its six surfaces covered by a thin Si detector, which together perform low-resolution particle spectroscopy with a rough angular resolution over a field-of-view covering ~1/4 of the whole sky. The energy range of detected particle spectra is 0.1–3 MeV for electrons and 1–30 MeV for protons.

A major task for the SIXS instrument is the measurement of solar X-rays on the dayside of Mercury's surface to enable modeling of X-ray fluorescence and scattering on the planet's surface. Since highly energetic particles are expected to also induce a significant amount of X-ray emission via PIXE (Particle-Induced X-ray Emission) and bremsstrahlung when they are absorbed by the solid surface of the planet Mercury, SIXS performs measurements of fluxes and spectra of protons and electrons. SIXS performs particle measurement at all orbital phases of the MPO as the particle radiation can occur also on the night side of Mercury.

The energy ranges, resolutions, and timings of X-ray and particle measurements by SIXS have been adjusted to match with the requirements for interpretation of data from Mercury's surface, to be performed by utilizing the data of the MIXS (Mercury Imaging X-ray Spectrometer), which will measure X-ray emission from the surface.


Figure 55: Illustration of the SIXS instrument elements (image credit: SIXS team)

Measurement principle: SIXS is capable of performing measurements of X-ray spectra with time resolution down to 1 second in the energy range 1- 20 keV, and simultaneous proton and electron spectra in the energy range 0.33–30 MeV for protons and 50 keV–3 MeV for electrons. Both X-ray and particle channels are capable of measuring count-rates up to 20,000 cps with very small pile-up, and have a low background. Both channels have a total FOV of at least 180º in diameter. The detectors can be operated in near room temperature, and the detector materials are highly radiation tolerant. The instrument has on-board radioactive sources (Fe55) for spectral calibration of the X-ray detectors.

Science data will be made available in PDS, FITS, and partly in ASCII formats. Scientific analysis of the data will be carried out primarily at the home (PI) institutes of the SIXS and MIXS instruments. Standard science analysis will include determination of the physical energy scale and spectral resolution of the X-ray data using background signal from the Fe55 source, and fitting of the X-ray spectra with publicly available astronomical X-ray spectrum analysis software. The X-ray standard analysis results will be further interpreted for purposes of fluorescence analysis of Hermean surface by MIXS, and also for independent studies of X-ray corona and flares of the Sun.