Thermal design and thermal verification tests of the Solar ...

45th International Conference on Environmental Systems ICES-2015-76 12-16 July 2015, Bellevue, Washington

Thermal design and thermal verification tests of the Solar Orbiter Heat Shield STM

C.Damasio1 European Space Agency, ESA/ESTEC, Noordwijk ZH, The Netherlands

P. Defilippis2

Thales Alenia Space Italia, Strada Antica di Collegno 253, 10100 Torino, Italy

C. Draper3 and D. Wild4 Airbus Defence & Space Limited,Gunnels Wood Road, Stevenage, SG1 2AS, UK

Solar Orbiter is the next solar-heliospheric mission in the ESA Science Directorate. The mission will provide the next major step forward in the exploration of the Sun and the heliosphere investigating many of the fundamental problems in solar and heliospheric science. One of the main design drivers for Solar Orbiter is the thermal environment, determined by a total irradiance of 13 solar constants (17500 W/m2) due to the proximity with the Sun. The spacecraft is normally in sun-pointing attitude and the main barrier to protect the satellite from severe solar energy is the Heat Shield. The Heat Shield is mainly composed by a panel aimed to provide a support and the interfaces with feedthroughs and doors for the Remote Sensing Payloads. The insulation from the extreme external environment is guaranteed by high and medium temperature MLI blankets. A thermal verification campaign of the Heat Shield STM (Structural Thermal Model) has been performed in the second quarter of 2014 to verify the thermal design and to correlate its thermal mathematical model. A first Thermal Balance test in vacuum was performed in May in the ESTEC Large Space Simulator where the heat Shield was subjected to solar fluxes of 6 and 10 solar constants. A second Thermal balance Test was performed in IABG solar simulation test facility in June where solar fluxes of 1.34 and 0.41 solar constants were used.

This paper will describe the Heat Shield thermal design, the Thermal Balance Tests performed to verify the thermal design and the relevant TMM correlations.

Nomenclature AU = Astronomical Unit BOL = Beginning of Life CDR = Critical Design Review EOL = End of Life FM = Flight Model HTMLI = High Temperature MLI IABG = Industrieanlagen Betriebsgesellschaft mbH LSS = Large Space Simulator LTMLI = Low Temperature MLI MLI = Multi Layer Insulation S/C = Spacecraft 1 Solar Orbiter Thermal System Engineer, Thermal Division (TEC-MT), [email protected], +31 71 5656276. 2 Heat Shield Thermal System Engineer, Thermal Systems, [email protected], +39 011 7180679 3 Solar Orbiter Heat Shield Assembly Manager, TOPME, [email protected], +44 1438 773495 4 Solar Orbiter Thermal Architect, Thermal Architecture, [email protected], +44 1438 778148

mailto:[email protected]



International Conference on Environmental Systems

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Figure 1 The three science windows

STM = Structural Thermal Model TBT = Thermal Balance Test TMM = Thermal Mathematical Model SOLAR ORBITER Instruments EPD SIS = Energetic Particle Detector - Suprathermal Ion Spectrograph EPD STEP = Energetic Particle Detector - Supra Thermal Electron Proton EPD HET-EPT = Energetic Particle Detector - High Energy Telescope - Electron Proton Telescope EUI = Extreme Ultraviolet Imager MAG OBS = Magnetometer Outboard Sensor MAG IBS = Magnetometer Inboard Sensor METIS = Multi Element Telescope for Imaging and Spectroscopy, coronagraph PHI = Polarimetric and Helioseismic Imager RPW ANT ENNA = Radio and Plasma Waves, Antennae RPW SCM = Radio and Plasma Waves, Search-coil magnetometer SOLOHI = Solar Orbiter Heliospheric Imager SPICE = Spectral Imaging of the Coronal Environment STIX = X-ray Spectrometer/Telescope SWA HIS = Solar Wind Plasma Analyser - Heavy Ion sensor SWA PAS = Solar Wind Plasma Analyser - Proton & Alpha sensor SWA EAS = Solar Wind Plasma Analyser - Electron Analyser system

I. Introduction Solar Orbiter satellite, due for launch in 2018, will carry a suite of 10 instruments to observe the Sun from a

vantage point never before achieved with any spacecraft. Solar Orbiter will be launched into an escape trajectory on a NASA-provided Atlas-V or Delta-IV launcher, with

Ariane-V back-up. Subsequent mission phases are: - Launch & Early Orbit Phase of 7 days. - Near-Earth Commissioning Phase of 3 months. - Cruise Phase of about 3 years, during which it will make three gravity assist manoeuvres, first with Venus,

then twice with the Earth. Between the first two manoeuvres, Solar Orbiter will achieve its greatest distance from the Sun, at 1.47 AU. This is a key distance for sizing the spacecraft’s solar array.

- Nominal Mission Phase of 3.8 years, where the spacecraft will enter its operational orbit around the Sun followed by an Extended Mission Phase of 2.5 years. Over these periods, Solar Orbiter performs 14 orbits of the Sun, with perihelions of 0.28AU and aphelions of 1.1 AU, during each of which are three science windows of 10 days each (figure 1). Both in-situ and remote sensing payloads operate within the windows. In-situ data are acquired at all other times. The operational orbit will be in resonance with Venus, providing the means for five gravity assist manoeuvres which will gradually increase the orbital inclination, enabling observations at increasingly higher solar latitude up to 33°.

The primary science goals of Solar Orbiter involve answering fundamental questions on how processes on the Sun


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Figure 2 Instruments of Solar Orbiter

drive the properties and phenomena of the heliosphere. To achieve this, the payload complement consists of 6 remote sensing instruments and 4 in-situ instruments, which are to be operated simultaneously during the main science observation windows, thus providing simultaneous observations of phenomena on the solar surface with measurements at the location of the spacecraft (figure 2). There are three such observation windows of 10 days each per orbit, centered around perihelion and around the points of maximum and minimum solar latitude. For the remainder of the orbit, only the in-situ instruments will be collecting data.

The remote sensing instruments (SPICE, STIX, EUI, METIS, PHI, SOLOHI) cover a range of wavelengths from visible to X-ray and provide simultaneous high resolution spectroscopy and imaging, while the in-situ instruments measure the solar wind constituents in terms of the plasma and electromagnetic fields and waves. The combination of solar disc instruments with coronagraphs also enables simultaneous observation of surface processes with those in the corona.

The spacecraft itself is three-axis

stabilized, with a heat-shield on one face, which is nominally sun-pointed throughout the mission. Figure 3 shows the location of the instruments on-board the spacecraft. The majority of the spacecraft and instruments are protected by this heat-shield such that their environment is relatively benign. Exceptions to this are the spacecraft appendages, i.e. the solar arrays and high gain antenna, along with those parts of the instruments which require a view of the Sun, e.g. the remote sensing instrument apertures. Several in-situ instrument sensors are placed on a boom in order to minimize the electromagnetic interference from the spacecraft and other instruments.

The total spacecraft wet mass at launch is 1800 kg, of which the total mass budget of the payload suite is 190.4 kg. The industrial team is led by Airbus Defense & Space at Stevenage (UK), TAS-I at Turin (Italy), leads the Solar Orbiter’s Heat Shield.

Figure 3 Solar Orbiter in stowed configuration and exploded view


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Figure 4 Schematics of the Heat Shield configuration

II. Description and need of Solar Orbiter Heat Shield

During its mission around the sun, Solar Orbiter will receive a solar flux of about 17500 W/m2 (i.e. 13 solar constants) at the perihelion at 0.28 AU and this is the major thermal design driver. The protect Solar Orbiter, a heat shield has been sized to provide a complete sun shielding of the platform and acts as thermal insulation of the entire spacecraft including all equipment and parts of the structure within an 8º half-cone to cover both the operational modes of the satellite and maximum off-pointing when the satellite enters safe mode (figure 4).

The Heat Shield allows the payload to function at its optimum temperature despite its proximity to the sun. A stable thermal environment is further achieved by conductively isolating the heat shield from the rest of the spacecraft by use of a series of discrete mounting blades and radiatively through the use of MLI.

The protection it offers from large variations in thermal environment has helped simplify Solar Orbiter’s design by allowing conventional thermal control techniques to be implemented in the platform. This allows a greater degree of confidence to be placed in the thermal design by reducing the need to use new and innovative technology

High temperature heat barriers and a Support Panel provide two gaps between the spacecraft and front shield (the front shield is supported by star brackets). Heat is rejected away from the spacecraft and out to space through these two gaps between the panels of the heat shield.

Material selection is made with avoidance of contamination as a major objective learning from problems experienced on the development of the BepiColombo High Temperature MLI. The structural materials used are Carbon Fiber reinforced Polymers and Titanium each of which have a low coefficient of thermal expansion, the heat shield interface is by Titanium blades which will distort under thermal load to avoid loading the satellite structure.

A series of feedthroughs passing through the heat shield to the instrument platform provide a view of the sun to the onboard instruments (Remote Sensing). The heat shield is based on breadboards initially designed for a solar flux nearly twice that of Solar Orbiter. Feasibility tests were performed in 2007 under ESA Technological Development Activities. The proposed heat shield design blends the best features from the two heat shield designs developed independently by TAS-I and Airbus Space and Defense during previous industrial studies.

III. Detailed design of Solar Orbiter Heat Shield STM


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Figure 6 View of the Support Panel from the S/C side (Support Panel MLI) and blades

Figure 5 The Front shield with the feedthroughs with doors open

The Heat Shield STM, which was subjected to thermal balance tests, consists of a series of physical barriers separated by gaps which allow lateral rejection of infrared-radiation to cold space. The total height of the Heat Shield is 400 mm and its main elements are: the High Temperature MLI (HTMLI) facing the sun, the Support Panel, the Low temperature MLI (LTMLI) on the upper side of the Support Panel, facing the HTMLI and the Support Panel MLI on the lower side of the Support Panel (facing the S/C).

The main structural element is the Support Panel (2940 mm by 2270 mm by 54 mm thick): a sandwich panel

made by an Aluminium alloy honeycomb 3/16-5056-0.0007p and two skins of six plies (0.67 mm thick) of quasi-isotropic layup of Carbon laminates made up by the high modulus, high thermal conductive K13D2U carbon fibers, by Mitsubishi, impregnated by Hexcel M18 resin. The composition of these skins was selected with the aim of achieving best thermal load uniformity to minimize the thermo-elastic distortion of the platform to ensure the co-alignment of the instruments.

The Support Panel has Titanium blades on the lower side interfacing the satellite structure and star brackets on the upper face supporting the High Temperature MLI designed to relief the forces at the S/C interfaces due to deformations thermally induced by the Heat Shield itself.

To avoid exceeding its maximum qualification

limit of 160 °C, the Support Panel is completely wrapped by MLI installed by means of non-metallic velcros glued on the carbon fiber skins:

a) on the upper side (+X) by the Low Temperature MLI: it is a 20 layer MLI lying on the Heat Sheild Support Panel on the upper side. All the layers including the external one are double aluminized Upilex. It is separated from the HTMLI by the Heat Shield main gap, 245 mm wide which allow lateral rejection of infrared-radiation to cold space (shown in figure 5, in the left side and in figure 9). Glass fiber is used as separator between the internal layers.

b) on the S/C side (lower side) and on the edges by the Support Panel MLI that is a 20 layer MLI with black kapton as external layer (figure 6). The use of MLI on the Support Panel back side has been introduced mainly to overcome the use of heaters in cold case science mode at the aphelion (1.47 AU). Glass fiber is used as separator between the internal layers. The gap between this Support Panel MLI and the S/C MLI located on the S/C panel create a secondary gap about 100 mm wide that is also helping in rejecting the heat to cold space.

The main insulating element of the Heat

Shield thermal control is the HTMLI (figure 5). It is composed by a Front Shield that will be exposed to sun and is a layer made in Titanium (50 μm thick) black coated using a process called Co-Blast developed by the Irish company Enbio. It has an absorptivity/emissivity ratio of about 1.1 and it is able to stand high temperature when exposed to sun. Its thermo-optical properties were verified in under UV and particles environments at high temperature in the frame of an ESA Technological Development Activity. The rest of the HTMLI is composed by six bare Titanium layers 10 μm thick used as


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Figure 7 LTMLI, cylindrical and corner feedtroughs and star bracket

Figure 8 Feedthrough and its mechanism, the door has been removed

internal foils and a 25 μm thick bare Titanium layer facing the Support Panel. Seven glass fiber separator layers are placed between all the Titanium foils.

The HTMLI is mechanically sustained by the Support Panel by means of the 10 Titanium Star-brackets. The mechanical fixation of the HTMLI on the Star-brackets is achieved by means of metallic velcros riveted directly on the Star-brackets arms. Thermal stand-offs in Torlon are used to insulate the

Support Panel from the hotter Star-brackets thus minimizing the heat flux to the Support Panel. The Heat Shield provides the apertures through which some payloads can observe the sun (see figure 7 and

figure 8 where the Front Shield is not yet installed). The apertures are created by Titanium feedthroughs mounted on the Support Panel and sized to respect the instrument unobstructed field of view requirements, but minimized to limit absorbed solar energy.

The feedthrough design of five remote sensing instruments has been performed by Sener in Bilbao (Spain) and consists of a Titanium cylindrical hollow structure, mounted to the Heat Shield Support Panel and providing the flanges to support the HTMLI. They are manufactured from Titanium alloy with a diffuse coating for compliance to the high temperatures; the heat absorbed from the solar flux is radiated laterally to space through the main gap at the edge of the Heat Shield.

The hollow structure is made by three cylinders 1 mm thick separated in between by two internal vanes that, together with the top and bottom flanges of the feed-throughs, define the unobstructed field of view for the instruments taking into account the thermo-elastic distortions of the S/C panel, the Heat Shield Support Panel and the feed-throughs themselves. Each aperture is protected by a door that swivels round a driving shaft moved by a dedicated motor.

The doors are intended to be open throughout the majority of the mission to allow the instrument to perform science observations. They have been implemented in case of an instrument malfunction to prevent sun-light entering the satellite enclosure and safeguard the operation of the remaining instruments.

The doors are operated via a drive shaft connected to a drive mechanism which is installed on the heat shield Support Panel to benefit from the more benign thermal environment. The electrical interface between the satellite and the doors is provided through a Doors Electronics Unit mounted on the platform.

The two Feedthroughs for the SWA-PAS and SWA-HIS instruments are mounted on two corners of the Support

Panel. They are formed by a monolithic Titanium piece consisting in an “L-Shape” plate with an average wall thickness of 3 mm and flat inner surfaces that present a stepped shape surface in the lower half.


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Figure 9 Main gap of the Heat Shield with feedthroughs and star brackets

Temperature considerations also have a significant effect upon the design of the feedthrough assemblies.

Thermal control of the feedthroughs is achieved by internal and external design features. The internal and external surfaces are black coated with plasma vacuum deposited process using AlTiN deposition made by METAL ESTALKI Company (Bilbao, Spain). It has an absorptivity/emissivity ratio of about 1.2 and it is able to stand high temperature when exposed to sun. Its thermo-optical properties were verified under a confidence test where UV and particles environments were simulated at high temperature.

The black coating on the internal sides is aimed to reduce excess light seen by the instruments, whilst on the external sides is used to maximized the heat rejection towards the Heat Shield gaps and then to space. Table 1 gives more details on the thermo-optical properties.

To avoid exceeding its maximum qualification limit of 160 °C, the Support Panel is thermally insulated from the hotter feedthroughs and the relevant mechanisms by means of Vespel thermal washers. This insulation will also help in minimizing the stresses generated by varying elastic deformation rates between the feedthrough components.

To minimize the impact of the very hot corner SWA feedthroughs that reach temperatures around 300 °C, Aluminium thermal radiators and doublers 2 mm thick have been adopted on the interfaces between the Support Panel and the corner feedthroughs. These radiators are painted with the conductive white paint MAP PCBE.

The performance of the Heat Shield is required to

limit the average radiative heat flux to the spacecraft to within +/-30 W and conductive heat to within +/-15 W throughout the mission, thus limiting the load on the platform radiators for subsequent heat rejection.

In addition the Heat Shield provides radiative decoupling to minimize the heat transfer from the its rear side to the cold payload radiators located on the S/C ±Y walls.

The conductive heat flux requirement is met using ten thin Titanium blades on the lower side of the Support Panel interfacing the satellite structure designed to relief the forces at the S/C (see figure 6).

To further minimize the radiative flux from the Heat Shield to the S/C, the S/C panel supporting the Heat Shield is insulated by MLI provided by the Airbus Defense & Space.

IV. Heat Shield design changes for flight After dedicated shedding test made on a small scale sample of the HTMLI, it was found that the glass fiber used

in both HTMLI and LTMLI was causing a contamination level that cannot be tolerated by some instruments. It was decided to remove all the glass fiber layers and the MLI design for FM was modified in the following

way: a) HTMLI: the internal layers are made in dimpled bare Titanium foils without spacer. b) LTMLI: the internal layers are made in embossed double aluminized Upilex layers without spacer. A first series of tests has been carried out on small, undisturbed blankets representative of HTMLI and LTMLI.

Scope of this test was to define the number of layers necessary, after elimination of separator net, to give the MLI equivalent characteristics of the blankets used in the STM TBTs (with separator net). With that purpose, several


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Figure 10 Front Shield illuminated by LSS bean at 10 solar constants

calorimetric tests have been performed in ESTEC TEC/MTV laboratory using a circular heating plate with 0.6 m diameter suspended inside a thermal vacuum chamber provided with temperature-adjustable shrouds; the heating plate has been covered on both faces by identical MLI blankets. A further annular blanket will protect the edge of the plate to avoid heat leakage through the borders. The heating plate will be controlled at fixed temperature (up to ~ 550 °C for HTMLI, up to ~240 °C for LTMLI).

The validation of the MLI design will be performed by performing large MLI (about 1 m2) tests where the

HTMLI and LTMLI will be tested representing the MLI overlappings and MLI mounting system as in flight configuration. The final confirmation of the MLI design will be achieved in the S/C FM campaign where the Heat Shield FM will be tested together with the S/C. During the TBT, it was also found that the Vespel washers used between the feedthroughs and the Support Panel were reaching temperatures very close to their maximum limit. For FM, it is planned to substitute them by Titanium washers with a very small thickness (1.2 mm). Mechanical and thermal analyses are still under execution to have the final confirmation of the feasibility of this solution. Aside from these changes, the design of the Heat Shield as described in earlier sections remains the same.

V. Thermal verification of the Heat Shield Structural Thermal Model

The Heat Shield STM Thermal Balance test intended to reproduce the operating temperatures of the Heat Shield components (as far as allowed by facility limitations) in several flight conditions.

The specific goals of the Heat Shield STM Thermal Balance Test are: • to demonstrate that the selected heat shield concept is capable to withstand the severe thermal environment

and to minimize the heat leak into the spacecraft, • to demonstrate the feasibility (and applicability to the Flight Model) of manufacturing/integration solutions

and techniques for mechanical and thermal interfaces with the spacecraft, • to validate the design versus potential thermo-elastic effects that could damage the layers of the thermal

barriers, • to validate the Thermal Mathematical Model that will be used for future flight temperature predictions, • to demonstrate that the deflections of one edge of the front shield (both in plane and out of plane) remain

below the value specified by an instrument. This was achieved measuring the movements of special Titanium targets mounted on the front panel edges, by means of a dedicated photogrammetry apparatus available in ESA.

The verification of the Heat Shield STM

thermal performance consisted in measuring the temperatures, in selected representative conditions, to determine the overall heat transfer to the S/C by all paths (conduction through the supports of the Heat Shield and Feedthroughs, radiation through the multilayer insulation) together with the temperatures of local critical interfaces. The acquired data (taking into account the difference in MLI materials between STM and FM) have been correlated with the analytical predictions for the STM test configuration. The validated TMM shall be used for the future flight predictions.

Due to other commitments of the ESA

LSS facility, the STM thermal balance test


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Figure 11 Front Shield as illuminated by the LSS Beam

Figure 12 Location of thermocouples on the Front Shield

was split in two campaigns and carried out firstly in the ESA LSS facility in Noordwijk (hot and intermediate cases, run in the period 12-15 May 2014) and then in the IABG solar simulation test facility in Munich (cold cases, run in the period 18-23 Jun 2014).

Four levels of solar flux intensities were considered adequate for a good TMM correlation: • 10 solar constants (13700 W/m2) in LSS • 6 solar constants (8240 W/m2) in LSS • 1.34 solar constants (1841 W/m2) in IABG • 0.41 solar constants (570 W/m2) in IABG

While for the lowest figure the standard LSS configuration

(with 6 m diameter “cylindrical” solar beam) is more than adequate, for intensities higher than 2 solar constants, a reconfiguration of the LSS mirror is necessary to obtain an higher level of flux by means of a “conical” solar beam on a smaller illuminated area within the test volume (figures 10 and 11).

Unfortunately the illuminated area at 10 solar constants is significantly smaller than the heat shield footprint. On the other hand the minimum diameter necessary to cover the entire Heat Shield (about 4 m) could be only achievable at the price of much lower flux intensity. It was therefore decided to maintain the nominal maximum intensity and to position the illuminated area as shown in Figure 11 to include all central feedthroughs and a corner one.

The beam of the IABG facility was able to illuminate the complete Heat Shield.

In both tests, a S/C simulator was used to support the Heat Shield and to create a boundary condition

representative of the S/C panel (+50 °C in the first three cases and -20 °C in the coldest one). The simulator was equipped with heaters to obtain the needed panel temperature and with MLI on both sides to have the best representation of the S/C MLI and to minimize the heater power.

VI. Test instrumentation A total of 476 type K and type T thermocouples were used to record the temperatures during both thermal

balance tests. 60 Type K thermocouples were used where high temperatures were (from 260 to 520 °C). - HTMLI: 57 sensors (K) - LTMLI: 16 sensors (L) - Support Panel: 70 sensors (L) - Support Panel MLI: 14 sensors (L) - Star-brackets: 12 sensors (L & K) - Blades connecting the Support Panel to S/C simulator: 14 sensors

(L) - SWA radiators and doublers: 10 sensors (L) - SWA feedthroughs: 10 sensors (L & K) each feed-through - Cylindrical feedthroughs: 176 sensors (L & K) (about 20 each feed-

through) - Feedthrough mechanisms: 36 sensors (L & K) (6 each unit) - Sun sensors: 20 sensors (L & K) (10 each unit) - S/C simulator: 18 sensors (L) - Bundles: 13 sensors (L) The high temperature thermocouples type K were installed on the


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Figure 14 Installation of type K thermocouples on the Front Shield

Figure 13 Thermocouples near the corner feedthroughs (blue on radiator and red on Support Panel)

Figure 15 Thermocouples on the feed-throughs and mechanisms

internal sides of the two external layers of the HTMLI using a cover patch that was fixed by two staples as in the figure 14. The MLI venting slot windows have been be used to exit the cable from MLI, to avoid making holes in the MLI.

The type K thermocouples were installed on the star-bracket and on the feedthroughs using RTV 566 glue.

The type T thermocouples installed on the star-bracket were attached with RTV 566. All the other type T thermocouples were attached using adhesive tape. In the few cases where there was the risk to exceed the maximum limit of the adhesive, the tape was secured by an additional sewing thread.

Figure 12 shows the thermocouple location on the HTMLI. The

thermocouples located on the LTMLI, the Support Panel and the Support Panel MLI were installed in the same configuration. In some local areas more thermocouples were used like on the Support Panel and on the doubler close to the SWA HIS corner feed-through (see figure 13).

Figure 15 shows the typical thermocouple instrumentation for the feed-throughs and on the mechanisms.

In addition to the thermocouples to support the correlation activity,

the temperature of the Front Shield was recorded by means of IR cameras in both test facilities.

The infrared camera, embedded in an hermetically closed canister, was installed in the LSS main chamber, hanging on a nitrogen pipe on the top of the aperture of the auxiliary chamber (see figure 16). The distance between the canister and the heat shield in LSS was about 3.5 m.

The fig. 16 shows the temperature distribution at 10 solar constants

measured by the IR camera in LSS, the temperature is ranging from 400 °C in the red areas to 280 °C in the blue ones. The IR camera gave measurements very close to ones obtained by the thermocouples installed just below the Front Shield.


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Figure 17 Location of Heat Shield in LSS together with IR camera and videogrammetry cameras

Figure 16 Temperature distribution measured by the IR camera

VII. Videogrammetry for deflection verification For the operation of SOLOHI instrument, the deflection of a

portion of the edge of the Front Shield is required to deflect within certain limits.

Videogrammetry was selected to measure the deflections of one edge of the front shield (both in plane and out of plane) during the illumination phase at 10 solar constants in LSS

With 2 cameras stereo images were acquired of the targets that were positioned on the edge of the shield (see Figure 17). Additionally 2 scale bars with calibrated distances were used in order to have traceability of the measurement accuracy and to provide dimensional scale to the measurement point cloud.

The images were processed using the videogrammetry post-processing software and the 3D coordinates of the targets were calculated in a least squares bundle adjustment. Finally the 3D coordinates, thus the relative thermo-elastic deformations can be compared between the different phases of the thermal balance test. The results showed a maximum deflection of 3.4 mm at 10 solar constants. Extrapolating to 13 solar constants, it means a figure of 4.4 mm well below the requirement of 15 mm specified for the Heat Sheild.

VIII. Thermal Mathematical Model description To predict the thermal performances of the heat

shield in flight and during the thermal balance tests, a detailed thermal mathematical model was built. It consists of about 7900 nodes (shown in figure 18) to model the HTMLI, LTMLI, Support Panel, Support Panel MLI, the corner feed-through doublers and radiators, the star-brackets, the blades and is incorporating the model built by Sener to represent the thermal behavior of the feed-throughs and mechanisms. The nodes on the Front Shield, HTMLI and LTMLI have dimensions of about 0.23 m x 0.44 m. The Support Panel has been modelled by 567 nodes having dimensions of 0.086 m x 0.142 m.

The main characteristics of the TMM are the thermo-optical properties of the various surfaces. The values used

in the flight analyses and in the test predictions are in table 1.


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Thermo-optical properties

Solar Absorptivity

IR Emissivity

Co-blast coating on Front Shield 0.95 0.81 Bare Titanium on HTMLI @ 550 °C 0.51 0.27 Bare Titanium on HTMLI @ 200 °C 0.51 0.17 Black kapton 0.93 0.84 Aluminized Upilex LTMLI 0.05 0.04 -0.06 MAP PCBE white paint 0.27 BOL

0.45 EOL 0.91 BOL 0.85 EOL

AlTiN PVD black coating for feed-throughs and mechanisms

0.88 0.75

Table 1 Thermo-optical properties

Temperature levels at 0.28 AU

0.28 AU °C

1.47 AU °C

front side of Front Shield 522 42 HTMLI rear side 346 -3 LTMLI 270 -72 Support Panel 148 -78 Support Panel MLI 112 -96 Star-brackets 203 -85 S/C MLI 96 -75 S/C panel 50 -10 SWA radiators 123 -89 SWA doublers 143 -84 Cylindrical feed-throughs 90 to 360 -90 to -10 Corner feed-throughs 300 to 360 -23 to -40 Motors 100 -93

Table 2 Temperature predicted for flight at 0.28 AU

Figure 18 Front Shield nodal breakdown

The HTMLI and the LTMLI efficiencies were derived by extrapolation from test data obtained for similar MLI in BepiColombo project.

The temperature levels reached in flight at 13 solar constants at perihelion including uncertainties are shown in the next table 2.

IX. Correlation steps and first results Due to constraints in the system schedule, it was decided to

postpone the correlation activity after the Heat Shield CDR that was held end of the 2014 year. Therefore only the first results are available for this paper.

The test predictions for the hot cases were performed using the

LSS model with the representation of the beam in the central plane inside LSS as used for the BepiColombo tests.

Due to the fact that the Heat Shield during its TBT was installed 2,5 m farther from the LSS mirrors, the beam LSS inhomogeneity (average 13747 W/m2, maximum 16518 W/m2, minimum 8913 W/m2) was different from the one in the middle position (BepiColombo TBT) and the LSS model was not suitable.

Therefore it was decided to perform the first correlation steps by increasing the number of nodes of the Front Shield using a mesh of 100 mm x 100 mm and assign to each node the measured flux value provided by ESTEC. The fluxes on the feedthrough nodes were calculated considering the local flux in the relevant position.

In the meantime, some local adjustments of TMM were performed to make the Heat Shield TMM better representing the effective hardware subjected to test. The main modifications were the change of the Front Shield emissivity for the hot case from the old values of 0.81 to the new measured value of 0.92 at 500 °C.

Table 3 provides a summary of the first step of the correlation for the 2 hot cases in LSS and the two cold cases

performed in IABG facility. The main results are briefly discussed here below: Front Shield: The average temperature differences between test measurements and TMM results are about 26 °C

and this is already at an acceptable level. In hot cases only the thermocouples located in the illuminated area far from the beam fall down zone have been considered. In cold cases, the differences between the average temperatures are smaller, and this is expected to be due to the uniform beam that was used in the IABG facility.


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Hot PHASE 6 solar constants

Hot PHASE 10 solar constants

Average

Test T [°C]

Average TMM T [°C]

Average Delta T [°C]

Standard Deviation

T [°C]

Average Test

T [°C]

Average Calculated

T [°C]


Standard Deviation

T [°C] Front Shield

HTMLI Illuminated

area

264.6 287.9 -23.3 38.1

337.9 364.4 -26.5 43.2

Illuminated area of

HTMLI rear side

140.4 147.0 -6.6 26.7

196.8 196.3 0.5 30.3

HTMLI rear side – All TC 115.5 122.4 -4.3 25.7 168.4 169.2 2.2 28.8

LTMLI 52.4 48.9 3.5 12.5 98.7 87.3 11.4 13.9 Support Panel 26.6 16.0 10.6 13.3 63.3 44.6 18.7 15.5

Support Panel MLI -34.9 -51.4 16.6 15.2 -11.4 -36.3 24.9 17.4

Illuminated feed-throughs 89.0 56.1 32.9 31.4 139.4 95.3 44.1 19.1

feed-through mechanisms 68.5 51.6 17.0 29.7 115.1 88.4 26.7 16.1

Cold PHASE

0.41 solar constants Hot PHASE

1.34 solar constants Average

Test T [°C]

Average TMM T [°C]


Standard Deviation

T [°C]

Average Test

T [°C]

Average Calculated

T [°C]


Standard Deviation

T [°C] Front Shield

HTMLI 47.1 46.9 0.2 15.1 144.2 154.2 -10.0 38.0

HTMLI rear side -25.5 17.5 -43.1 17.9 48.0 87.2 -39.2 20.0

LTMLI -70.8 -77.8 7.0 4.7 -12.9 -17.7 4.9 6.7 Support Panel -74.5 -84.1 9.6 8.6 -22.6 -32.4 9.9 10.4

Support Panel MLI -100.0 -106.8 6.9 6.2 -62.4 -75.4 13.0 7.5

All feed-throughs -67.2 -81.5 14.4 3.3 -10.4 -29.3 19.0 5.2

feed-through mechanisms -59.7 -71.2 11.5 3.5 -0.9 -14.6 13.8 4.4

HTMLI rear side, LTMLI: The maximum average temperature differences between test measurements and TMM results are about 10 °C in the two hot cases; this shows a good prediction of the radiative performances of the Heat Shield main gap. The difference on the rear side of HTMLI is much bigger in the cold cases. The reason is still to be investigated and could be due to the variation of the Titanium emissivity versus temperature.

Support Panel: the model predicts colder temperatures for the Support Panel and this discrepancy is important at 10 solar constants (about 19 °C). This may be due to an overestimation of the LTMLI insulating performances.

Support Panel MLI: the model predicts colder temperatures for the Support Panel MLI and this discrepancy is important at 10 solar constants (about 25 °C). As for the Support Panel, this may be due to an overestimation of the LTMLI insulating performances.

Feed-throughs: the model predicts colder temperatures for the feedthroughs and this discrepancy is more important at 10 solar constants (about 44 °C) and much less in the cold cases. This may be due to the model of LSS beam in the hot cases. To increase the flux and reach the high solar flux values needed to perform the Heat Shield test, the beam is focused by mirrors and had a convergence of 8° half-cone. This is not yet modelled in the TMM and the flux is still considered as a parallel beam. This effect could explain the discrepancy of the feedthrough: the convergent beam will impinge the cylindrical sides of the feed-throughs making them hotter than expected. In cold cases the discrepancy can be due to the variation of feed-through emissivity versus temperature not yet taken into account (the PVD coating emissivity is varying from 0.64 at 27 °C to 0.84 at 500 °C). The cold cases have been run considering a rough TMM of the IABG test facility.


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To improve the correlation, the next main steps will be: • The use of the LSS model to simulate the beam distribution on the position of the Heat Shield on the plane

at -2.5 m that was made available by ESTEC in the summer 2014. • The change of the emissivity of feedthroughs from 0.75 to the values that will be interpolated from the

measurement results • Sensitivities to assess the performances of the HTMLI, LTMLI and Support Panel MLI. • Sensitivities to assess the conductance of the bolted interfaces between the feedthroughs and the Support

Panel. • Use of IABG facility detailed TMM

All these steps are expected to improve mainly the correlation of the feedthroughs and, consequently, of the

Support Panel nodes in the area.

Acknowledgments The authors would like to acknowledge the Engineering Services Sections (TEC/MXE) of ESTEC for the effort

and time spent to provide timely support and information helping in progress in their work during the test execution and then during the correlation.

Thermal design and thermal verification tests of the Solar ...

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