48th International Conference on Environmental Systems ICES-2018-10 8-12 July 2018, Albuquerque, New Mexico

BepiColombo MTM Thermal Balance/Thermal Vacuum Test

B. Weinert, J. Schilke1 Airbus Defence and Space, Claude Dornierstraße, Friedrichshafen, Germany

S. Tuttle2 Sigma Space Systems, Canberra, Autralia

and

Daniele Stramaccioni3 European Space Agency, Noordwijk, The Netherlands

The BepiColombo satellite is Europe's mission to Mercury. The mission consists of a spinning satellite “Mercury Magnetosphaeric Orbiter” (MMO) provided by JAXA and a 3 axis stabilized “Mercury Polar Orbiter” (MPO) provided by ESA. A dedicated “Mercury Transport Module” (MTM) provides the transfer from Earth to Mercury by means of solar electric propulsion. The Mission is led by Airbus Friedrichshafen who is has taken over thermal design responsibility from Airbus Stevenage.

The MTM TB/TV test is performed with solar simulation up to 8SC in ESA’s large space simulator (LSS) in November 2017. Thermal balance test phases are conducted to verify mainly the heater layout and heat rejection capability of CPS elements (thrusters, pipes). The general thermal design has been already verified by a STM solar sim test performed in 2013. Thermal vacuum test phases include the functional verification of the ion thruster propulsion. This is achieved by activation of the high voltage on the acceleration grids and a separate release of Xenon. Stringent pressure requirements in the chamber and inside the S/C present a challenge to the execution of those test phases.

This paper reports on the test performed. It includes test setup, modelling of the test setup and a quick evaluation of the test results.

Nomenclature AU = Astronomical Unit BC = BepiColombo, name of the mission HP = Heat Pipe LSS = Large Space Simulator (TV chamber) MMO = Mercury Magnetosphere Orbiter MPO = Mercury Planetary Orbiter MOSIF = Magnetospheric Orbiter Sunshade and Interface MTM = Mercury Transfer Module PFM = Proto-Flight Model RCS = Reaction Control System RCT = Reaction Control Thruster S/C = SpaceCraft SC = Solar Constant SEPT = Solar Electric Propulsion Thruster STM = Structural Thermal Model TAS-I = Thales Alenia Space Italia TCS = Thermal Control System TEC = Thermo-Electric Cooler (Peltier Element) UV = UltraViolet VDA = Vapor Deposited Aluminum αS = Solar Absorptance εIR = Infrared Emissivity

1 Thermal Systems Engineer, Airbus Defence and Space, Claude Dornierstraße, 88090, Friedrichshafen, Germany. 2 Manager, Sigma Space Systems, Canberra, 2905, A.C.T., Australia 3 Thermal Engineer, Mechanical Engineering Department, Keplerlaan 1, 2200AG Noordwijk, The Netherlands

Copyright © 2018 J. Schilke, Airbus Defence & Space I. INTRODUCTION

A. The Mission BepiColombo is Europe’s mission to planet Mercury. The mission consists of a Japanese spinning satellite “Mercury Magnetosphaeric Orbiter” (MMO) provided by JAXA and a 3 axis stabilized “Mercury Polar Orbiter” (MPO) provided by ESA. A dedicated “Mercury Transport Module” MMO (MTM) provides the transfer from Earth to Mercury. The Mission is lead by Astrium Friedrichshafen, with Thales Alenia Space Torino responsible for the MOSIF MPO and MOSIF thermal design. The S/C is lifted by an Ariane V rocket into a direct Earth escape. The trajectory from launch to MPO final orbit insertion at Mercury takes about 7 years and comprises several planet fly-bys for fuel saving: 1 at Earth, 2 at Venus and 6 at Mercury. Deceleration against the huge gravity of the sun is MTM done by solar electric propulsion located in a dedicated Mercury Transfer Module (MTM). During this transfer, the whole S/C is in a 3 axis stabilized sun oriented attitude, while the MPO in its final orbit is planet oriented. A sunshade for the MMO completes the Figure 1. Mercury Composite Spacecraft composed of Mercury composite spacecraft (MCS) stack. MTM, MPO and MMO inside its sunshade MOSIF. This paper addresses the TB/TV testing of the MTM.

B. The Thermal Environment The thermal environment of the mission is extremely variable spanning between extremes never experienced by other ESA missions before. During the cruise phase the distance from the Sun spans from 1.16 AU to 0.298 AU giving rise to solar flux intensities from 1,015 W/m2 (0.74 SC) to 15,380 W/m2 (11.3 SC). The composite S/C will cruise with a very limited variation of the solar aspect angle and will experiences eclipses in correspondence of the various planet fly-bys (see Figure 2). Adequate ground testing of the MTM under its extreme thermal conditions represents another large challenge and there are certain aspects which simply cannot be tested.

Figure 2. BepiColombo Cruise Trajectory

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II. MTM THERMAL DESIGN The MTM, being the transfer module, is missing some elements of a selfstanding S/C. The onboard computer, star trackers, reaction wheels, transmission system are all located inside the planetary orbiter MPO. The MTM is limited to chemical propulsion for attitude control and small out of plane maneuvres, solar electric propulsion and the necessary power subsystem and a remote interface unit (RIU). Due to high sun intensity, the sun oriented face of the MTM is covered by high temperature MLI. Another side of the S/C is accommodating the ion engines, a third one is blocked by the MPO. The remaining 3 sides are used as radiators with 9m² total area to remove the dissipations mainly from PCDU and PPUs. When exposed to the 15.4kW/m² solar flux at 0.298AU, it must allow any two of its four ~5kW electric thrusters to run simultaneously, while maintaining internal temperatures suitable for the operation of 24 bi-propellant reaction control thrusters, internal electronic equipment with standard temperature ranges, thruster pointing mechanisms and a Xenon propulsion system, which amount to almost 2kW of dissipation.

Figure 3. The main external features of the MTM’s thermal design

Key features of the thermal design are:

Sun shield and Skirt: the main sun shield consists of a double blanket arrangement of high temperature MLI, along with a rear cavity which enhances cooling of its inner region. More details of the MLI are reported in Ref.3. Along its lower edge it carries a skirt made of a specially white-coated titanium sheet. This improves the shading of the electric thrusters (in what is known as the “Engine Bay”) and reduces the temperatures behind the sun shield cavity. Radiators: the radiators are classical aluminium honeycomb panels painted white. They cover 3 sides of the MTM. The white paint is necessary as during loss of attitude the radiators could be exposed for several seconds to sun with 15.4kW/m² intensity. Tank Cut-Outs: the two side radiators have large cut-outs to aid the radiation of heat behind the sun shield to deep space, thereby keeping the tanks on the sun-side of the MTM within their temperature limits. Heat Pipe Network: for successful performance, the radiators rely on a good thermal contact being achieved between the horizontally-running surface heat pipes and the radiator panels on one side and the dissipating units on the other side. Vertically oriented heat pipes embedded in the radiators further spread the heat. Some surface heat pipes are extended to the anti- sun radiator, linking the side and rear radiator surfaces for increased area for heat rejection of the high dissipation of the PCDU. See Figure 4 and Figure 5.

Also the sun exposed thrusters and sun sensors are Figure 4. Principle sketch of heat pipe network coupled to heat pipes transferring the absorbed heat to the side radiators. 3 International Conference on Environmental Systems

Reaction Control Thrusters (RCTs): the MTM has 12 pairs of 10N thrusters. Four pairs are permanently sun- illuminated and some substantial measures were needed to avoid the onset of vapour lock at 0.3AU. This thermal design was already verified by thruster EQM, exposed to solar simulation and by a pair of EQM RCTs installed on MTM STM solar simulation test (see ref. 5). Solar Electric Propulsion System (SEPS): the MTM has 4 gridded ion electric thrusters (Figure 6 and Figure 7). These consume approximately 11kW of power and dissipate around 800W in their two processing units and lead to a further 750W of waste heat in the MTM’s Power Conditioning and Distribution Unit (PCDU). The SEPS includes various valves, pressure regulators and flow control units to regulate the Xenon flow from tanks to the ion thrusters. As the ion thrusters are kept in the shadow of the sun shield skirt, heat removal is relatively easy. SEPS Power Units: there are two of these Power Processing Units (PPUs) which provide power to the four electric thrusters. Each of them requires numerous heat pipes to distribute its waste heat. Thus, thermal contacts between those units and the heat pipes in these regions are important and must, therefore, be well instrumented in the thermal test to ensure good workmanship. Power Processing Unit (PCDU): this unit handles all of the power coming off the solar arrays (up to 20kW) and must distribute it to the MTM for electric propulsion and heating, to the MPO for all of its power needs and to the MMO for its heating. Its dissipation has a power density of around 5kW/m². Thus, the coupling to the heat pipes is critical in this area and must be confirmed by the testing. Tanks: the MTM carries 3 Xenon tanks with a capacity of 200kg each. Two are located on the sun side (+Y), one on the shadow side (-Y) to limit the CoG shift when the Xenon is consumed. Mainly for attitude control a hydrazine tank on -Y and an oxidizer tank on +Y complete the tank arrangement.

Figure 5. MTM -X and -Y radiators with surface heat pipes

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Figure 6. MTM sketch of +Z side

CPS RCTs compartment

SEPT

SADM

Flow Control Unit

Remote Pressure Interface regulator Unit

RCTs

Figure 7. MTM +Z side, all MLI removed

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III. THERMAL TEST REQUIREMENTS Adequate ground testing of the MTM under its extreme thermal conditions represents another large challenge and there are certain aspects which simply cannot be tested. The Large Space Simulator at ESTEC in the Netherlands was upgraded to enable testing with 10 solar constant sun simulation. However, this required some collimation of the beam, with the result that some of the extremities of the MTM are outside the main calibrated area of the beam and most significantly, the BepiColombo stack cannot be tested with the modules together. The Mercury Planetary Orbiter (MPO), which sits on top of the MTM, is represented by a special simulator structure in the test. This enables some of the important interactions to be verified, such as infra-red heating of the MTM radiators by the illuminated MPO MLI and sun-trapping at the interfaces between both modules. Note that for the thermal analysis, it has been necessary to include the external surfaces of the MPO in the MTM thermal model. Investigation by analysis also showed that the influence of the MOSIF on the MTM is minimal due to the larger distance and may be neglected both from the thermal modeling and from the thermal balance test without significant loss of representativity. Another critical thermal driver which cannot be fully simulated is the firing of the electric thrusters. For thermal balance therefore test heaters at the heat pipes underneath the PPUs compensate for the reduction of dissipation. For functional verification, each thruster was turning on the acceleration grids without Xenon flow, thus avoiding sending Xenon ions into the facility shrouds. Thereby the high voltage charge of several kV is verified front to end. In a second step each thruster is turned on without switching on the acceleration grids. Thereby the generation of Xenon plasma and the related Xenon flow rates are verified. The separation allowed to respect the more stringent pressure requirements during high voltage application, while relaxing those requirements during the release of Xenon. Potential sun trapping at the sun facing 10N RCT thrusters was already checked in a dedicated unit solar simulation test and in the MTM STM solar simulation test. Nevertheless the impact of MLI workmanship and the repeatability of the heat path from nozzle via thermal straps to radiator heat pipes needed a confirmation on the final flight hardware. Due to the need to concentrate the solar beam to get sufficient intensity, the 10N RCTs were outside the beam when the MTM was in nominal illumination attitude. Test phases thus were introduced that turned the MTM such that one side of RCT was illuminated. The 3-dimensional nature of the heat pipe network means that the vertical heat pipes are not expected to function against gravity. Some segments of the heat pipes from hotspots (heat injection) upwards into colder areas were working, while HP with heat injection only at the top end were not. The use of IR cameras to monitor the radiators was mandatory to identify working or non-working HP segments for the correlation.

To summarise, the main objectives of the test are to show or verify the following: a. Performance of the three-dimensional heat pipe network as a whole – primarily in terms of the average radiator panel temperature and average internal temperature of the MTM b. Verification of thermal interface filler performance at the large dissipative units, PCDU and both PPUs c. Sun shield MLI  10N thruster nozzle interactions and the suntrapping which may occur d. The sun side 10N thrusters – the thermal path from valves to heat pipes to radiator e. Tank temperatures and temperature gradients (among the three Xenon tanks and between the oxidant and fuel tanks) f. The temperature drop / heater power demand increase during eclipses around Earth g. Functional verification of all heaters and temperature sensors h. Tuning of heater control parameter i. Functional verification of the MTM, especially the solar electric propulsion system Implicit in the above is also the objective that the test provide adequate data in order to correlate the thermal model.

IV. MTM PFM TEST SETUP The facility used to conduct these tests is the LSS at ESA’S ESTEC in the Netherlands. Figure 8 shows the layout of the facility, in which can be seen the bank of 19 Xenon arc lamps of the solar simulator located underneath the chamber. The light beam is directed onto the test article by an array of collimating mirrors (far left in the figure). The main chamber in which the MTM was installed has the dimensions 9.5m diameter and 10m height. Following the upgrade to the sun simulator to high intensity for the BepiColombo mission tests, a converging, 6 International Conference on Environmental Systems

conical beam resulted and this can be seen in Figure 10, where the MTM would be mounted at the centre of the axes, as shown in the figure. The maximum solar flux which can be produced is equivalent to ~8 solar constants at the centre plane of the MTM. At the sun shield location, it is about 1 solar constant lower. A dedicated thermal model was used to simulate this light characteristic. Furthermore the intensity variations across the beam were measured and by semitransparent filters in the modelled light path also represented in the model.

Figure 8. The Large Space Simulator at ESA’s ESTEC facility

Figure 10. Conical beam

Figure 9. MTM illumination without/with rotation

The necessity of placing the sun facing RCT into the beam requires a rotation of the MTM by 13°. Theoretically the beam collimation angle is 6.5°. With a 20° inclination of the radiators, they should stay just in the shadow (Figure 9). Unfortunately the beam shape is not exactly conical, but more trumpetlike. The test runs revealed that a small portion of light is visible on the radiators (Figure 13). Turning the radiators out of the sun also moves the 10N RCTs partially out of the sun, hence it was agreed that in the rotated test phases the radiator temperature level is not correlated but set as boundary to the measured temperatures. The sun intensity is reduced compared to flight. This was necessary as it turned out during early BepiColombo tests, that at maximum flight intensity the lamp life is reduced significantly to a value below 24h. A stronger collimation would mean that a smaller portion of the S/C is within the sun beam. This would imply that a larger 7 International Conference on Environmental Systems

rotation is needed to bring the RCTs into the sun. But as the central part of the side radiators would be exposed to sun intensities beyond the heat rejection capability, the internal units would exceed their temperature limits. Hence the sun intensity was limited to 8 SC (11000W/m²).

The MTM is almost in flight configuration. A sketch of the test setup is shown in Figure 11. Besides the installation of about 600 thermocouples for additional temperature monitoring, the following deviations were made: • MTM solar arrays not installed. Due to the need to verify the function of the SADM by rotating in TV, also no possibility of installing yoke or simulators was feasible. The omission of the complete SA allowed to mount fix resistors simulating the thermistors present on the SA directly at the SADM connectors. This allowed checking the impact that temperature changes on harness and acquisition circuitry had on the temperature reading. Thereby the precision of those sensors in flight is increased. • MTM hold downs shielded by test MLI. Initially it was foreseen not to have the HDRM present during TB/TV. Due to rearrangement of the AIT schedule, the HRDMs were installed and the alignment with the stowed SA was performed prior to thermal testing. As the beam is collimated, the HDRM and all their harness cables would be in full view of the sun during the rotated test phases. This would bring the harness above qualification temperatures. Thefore local high temperature blankets were wrapped around each HDRM (see Figure 12) • The MTM is powered by harness connecting EGSE directly to the PCDU. As a consequence, the losses inside the SADM which are present in flight needed to be simulated by test heaters. • The CPS tanks were filled with simulant. The CPS function was not checked in TV at S/C level. • The Xenon tanks were filled only with 5bars instead of 200 bars in flight. They remained isolated from the pipework. The Xenon supply was provided by a bottle outside the TV chamber, connected to a test port upstream of the high pressure regulator. The pressure used during the test was 50 bar, representing a fill level close to end of cruise. • No real ion thruster firing was performed. The test sequence split this function into activation of acceleration grids and separate Xenon release and plasma generation • As no MPO was present, the command and control of the MTM was performed by EGSE. Instead of doing direct low-level commanding, a "mini MPO" was built. This was made of the EQM onboard computer and some additional EQM units to allow running the flight onboard software that controls the MTM. • Additional test heaters were installed. They compensated the reduced dissipation in SADM, PPU (due to not operating the ion thrusters at full thrust) and PCDU. The test heaters also acted as additional safeguards against wrong commanding or power failures.

MTM

MPO simulator

Spinbox and test adapter Figure 11. Sketch of MTM test setup 8 International Conference on Environmental Systems

Figure 12. MTM in sun, nominal sun direction

Figure 13. MTM in sun, rotated position

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V. TEST SEQUENCE AND CONDITIONS The test sequence (see Figure 14) started with thermal balance phases at hot temperature plateau (test phases 30 - 60). This allowed outgassing and reduction of residual internal pressure before switching on pressure sensitive equipment like PPUs. The different TB phases were interrupted by a simulation of an eclipse, which comprised not only switch off of sun but also a reduction of heater setpoints for power saving. The TB phases included two rotated phases to verify the sun facing RCTs. This test phase was repeated (as test phase 55, see Figure 15) because during initialization the command link to the MTM was lost on nominal side. It was decided to go to a safe condition, including warming up shrouds to 20°C and repair the failed component instead of continuation of the TB/TV test without redundancy for commanding. As a last hot phase a functional verification phase at maximum temperature (test phase 80), i.e. beyond predicted flight temperatures, was executed. In preparation of this phase, a waiting period at vacuum conditions was necessary for electronics outgassing purpose (test phase 70) before switching on the PPU was allowed. Thruster firing, i.e. plasma generation was planned but not executed, as the time duration to switch on thrusters at high temperatures was exceeding the allocated time slot. This verification thus was done only in cold conditions. After the hot plateau, a cool down into worst case cold survival was performed (test phase 90). Here survival heaters were verified, both in function on nominal and redundant side as well as verification and adaptation of heater setpoints. A thermal balance (equilibrium) phase concluded the thermal control verification. Thereafter functional verification in cold conditions started with the bus units, followed by SEPS high voltage (acceleration grid activation) and thruster "firing" (i.e. plasma generation). The sequence was chosen as the pressure increase inside the chamber during the "firing" due to released Xenon has large uncertainties and might have blocked the high voltage operation. In the end it turned out that the Xenon was trapped by the cryo panel and shrouds very effectively, so that the pressure increase during "firing" was still below 1e-05mbar.

Figure 14. Test profile (specified)

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Figure 15. Test profile (as run)

VI. TEST RESULTS MTM functional performance was successfully demonstrated. The SEPS system could be verified front to end with onboard software in the loop. Even if that meant splitting plasma generation and acceleration grid operation, all potential workmanship issues were checked. All heaters were operating as intended. Small adaptation of local areas of the heater density needed to be performed on 2 CPS pipe heaters. All heat pipes were functioning (some vertical ones only segment wise due to gravity). In addition to thermocouples at both end of the heat pipe, IR pictures were helpful to check operation of non-horizontal heat pipes (see Figure 17) Overall heat rejection and temperature level was as expected. The temperature gradients between PCDU or PPU and the heat pipes they are mounted to are even below predictions. This shows a good performance of the interface filler. The temperature drop in eclipse was slower than expected, leading to about 30% less energy consumption during eclipse. A problem was detected on the Xenon feed line. At a mounting bracket close to the chamber wall and LN2 shrouds, the insulating MLI was forgotten (see Figure 16). This resulted in locally insufficient heating, as the heater control thermistor was about 1m further downstream of the pipe. A two phase loop built up, where at this bracket condensation occurred and the condensation temperature is a function of the pressure in the pipe. The condensed Xenon flows down inside the pipe until sufficient heat for complete evaporation is collected. The heater control thermistor was always in the gas section. Initial attemps to increase the setpoints such that condensation could be avoided were not successful. The effect of the increased power was a slightly increased condenser temperature and a significantly increased length of the evaporator, so that the whole energy of the heater was still dumped to this cold spot.

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Figure 16. Missing MLI at Xenon feed line clamp

Figure 17. IR snapshot of -X radiator (test phase 30)

Figure 18. IR snapshot of +X radiator (test phase 30)

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The modelling of the test setup included the geometric modelling of the test chamber for IR heat exchange calculations and a dedicated solar beam modelling. The correlation is still ongoing due to other analysis priorities. As usual the first step was to include measured dissipations. Already this represented a problem at the PCDU: The dissipation was measured by summarizing the input current and output currents multiplied by the relevant voltages. The accuracy of the input current was sized for a total consumption of nearly 17kW, while during the thermal balance phases only about 2.5kW were consumed as S/C power. In flight about 3kW are distributed to the other modules, which was not simulated in the test. The ion thrusters at full thrust consume another 11kW, which was also not present during test. Missing power was supplied by test heaters via dedicated power supplies. As a consequence the input power at PCDU had an uncertainty of about 200W. The output power had lower uncertainties, as the current measurements were typically accurate to better than 2%. So the uncertainty on the output side was about 50W. The expected dissipation was about 250W. Considering the uncertainty of +/- 100% it was concluded to stay with the calculated dissipation. While almost all units are already correlated well (<3°C deviation), some thermocouples located on CFRP structure were more than 15K off. Looking into the predictions it became clear that the neighboring CFRP nodes were predicted with more than 20K temperature difference. The refinement of modelling is ongoing, but a trial with node sizes reduced to about 4 x 4 cm² (roughly heater sizes) significantly improved the predictions. Checks of the photo documentation to determine the precise location of the temperature sensors and allocate for predicted temperatures the weighted temperatures of adjacent nodes will finalize the correlation.

VII. LESSON LEARNED Instrumentation: • Photo record was very valuable (one survey photo, one detailed photo for each sensor, including sensors on GSE). Make sure that you can identify if some nearby thermal HW was installed / modified after taking the picture (e.g. MLI missing during photo session). MLI performance: • Due to the limited number of sensors in a S/C level test (and the need to keep blankets undisturbed for flight) it is difficult to correlate MLI performance from such test. Better perform a representative MLI test (i.e. typical blanket area, typical overlaps, typical amount of fixations, typical temperature of support structure, blanket grounding). Such test can afford placing thermocouples inside the blanket at different layers (during blanket manufacturing) Pressure requirements: • Vacuum level inside S/C is usually about factor 10 higher than outside • Negotiate with equipment what the real pressure needs are, where they apply and which safety margins are to be used. Check what reference measurements the suppliers have (pressure measured inside the unit, inside a compartment, outside the box). Verify if pressure requirements are valid at start of operation or if they determine abort criteria General: • It was helpful to combine relevant onboard data with thermocouple and test heater data in the same data base, so that those data can be combined in a single graph (temperature telemetry, unit and heater currents, onboard and test heater states). Thereby easy access to the test data even months later for comparison with AIT non-conformances was feasible. • Deficiencies in the test facility can be compensated to a certain degree by extra effort on the modelling side. For example, the spatial non-uniformity of the beam was simulated in the thermal model and this improved the overall outcome of the test in terms of the degree of model correlation achieved.

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VIII. CONCLUSION This paper has shown how the testing of the complex thermal design necessary for the MTM on its mission to Mercury had a wide variety of challenges, arising from the difficulty of reproducing the flight environment in a ground-based test facility and the differences between equipment performances in flight and under gravity. The thermal design needed to be tested to the highest degree practicable and this has been achieved without too much compromise to the constraints of the test set-up and the test facility. All of the key features of the MTM’s thermal design were shown to work as expected. It was confirmed that during the testing, it was possible to achieve the necessary temperature level of the 10N thruster when illuminated. Similarly, some of the important operational aspects could be clearly confirmed, such as the maximum heater power requirement, heater settings, eclipse heating energy and the final temperatures at the end of eclipse. The SEPS subsystem could be verified with all hardware in flight configuration using the flight software. Thus, it can be stated that the MTM PFM thermal balance / thermal vacuum test campaign has been highly successful and the MTM is ready for flight with minor adaptations on two heater lines.

IX. ACKNOWLEDGEMENTS Designing the test setup, performing the thermal testing, analyzing the results and deciding the way forward was a very complex activity, which was possible only thanks to the excellent teamwork realized among the different teams involved: the BepiColombo final customer ESA, the spacecraft contractor Airbus Defence & Space Germany (Friedrichshafen), the AIT subcontractor Thales Alenia Space Italia (Torino), ESA-TEC division who did sample tests and performed a lot of in-situ measurements. Finally, we would like to thank the test facility responsible ETS (European Test Services) who contributed significantly to the test success.

X. LIST OF REFERENCES 1 ESA BepiColombo homepage, URL: http://sci.esa.int/science-e/www/area/index.cfm?fareaid=30 2 Tuttle, S.L. and Cavallo, G. “Thermal Design of the Mercury Transfer Module,” 39th International Conference on Environmental Systems, Savannah, Georgia, 2009. 3 Moser, M. et al., “High Temperature Multilayer Insulation of the BepiColombo Spacecraft – a design at the edge of material capability” Proceedings of the 40th International Conference on Environmental Systems, Barcelona, Spain, 2010, AIAA-2010- 6091 4 Schilke, J., et.al., “BepiColombo MOSIF 10 SC Solar Simulation Test”, 41st International Conference on Environmental Systems, Portland, Oregon, 2011. 5 Schilke, J., et. al., “Development and Qualification of BepiColombo MLI”, Proceedings of the 43th International Conference on Environmental Systems, Vail, U.S., AIAA-2013-3504. 6 Tuttle, S. et. Al., "Thermal Testing of the Mercury Transfer Module", Proceedings of the 44th International Comference one Environmental Systems, Tucson, Arizona, ICES-2014-036

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