46th International Conference on Environmental Systems ICES-2016-120 10-14 July 2016, Vienna, Austria

Development of Passive Thermal Control for Surface Missions

S. Herndler1 and C. Ranzenberger2 RUAG Space GmbH, Stachegasse 16, 1120 Vienna, Austria

and

S. Lapensée3 ESA - European Space Agency, Keplerlaan 1, 2200 Nordwijk, The Netherlands

The extreme thermal environment on Mars asks for an effective thermal insulation to keep equipment within allowable temperature limits and to limit power consumption. Furthermore the thermal insulation should be light weight, flexible, adaptable and consume minimum volume. As a consequence of the Mars atmosphere, conventional vacuum based multilayer insulation offers low efficiency. Therefore, a novel thermal insulation making use of the existing Mars atmosphere was developed. This paper describes in detail the entire process and its results. Potential materials were identified based on literature, samples and manufacturer data and underwent material testing for characterization. Concepts and designs of three-dimensional demonstrators incorporating attachment, grounding and venting provisions were developed. Thermal performance of the demonstrators was validated by measurements in representative environments.

Nomenclature A = area, m2 Ar = Argon avg. = average β = volume expansion coefficient, 1/K CO = carbon monoxide CO2 = carbon dioxide CVCM = collected volatile condensable material d = thickness, m DHMR = dry heat microbial reduction ε = emissivity, - ESA = European space agency ESTEC = European space research and technology centre FS = fumed silica g = gravitational acceleration, m/s2 GG = gas gap GL = conductive heat exchange factor, W/m2K GR = radiative heat exchange factor, - h = heat transfer coefficient, W/m2K HDPIF = high-density polyimide foam k = thermal conductivity, W/mK λ = thermal conductivity of present fluid, W/mK L = characteristic length, m LAVAF = large vacuum facility

1 Systems Engineer, Thermal Systems, RUAG Space GmbH, Stachegasse 16, A-1120 Vienna, Austria 2 Head of Systems Engineering, Thermal Systems, RUAG Space GmbH, Stachegasse 16, A-1120 Vienna, Austria 3 Thermal Engineer, ESA-ESTEC, TEC-MTT, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands LDMRF = low-density melamine resin foam LDPIF = low-density polyimide foam MER = MLI = multi layer insulation MPF = MSL = ν = kinematic viscosity, m2/s N2 = nitrogen Nu = Nusselt number, - O2 = oxygen P = power, W Pr = Prandtl number, - Ra = Rayleigh number, - RML = recovered mass loss RTG = radioisotope thermoelectric generator σ = Stefan-Boltzmann constant, W/m2K4 T = temperature, K TC = thermocouple TGA = thermogravimetric analysis VDA = vacuum deposited aluminum WEB = warm electronics box

I. Introduction ars surface missions put challenging requirements on thermal insulation. It needs to be effective to limit M power consumption and to protect equipment from the harsh Martian environment at a minimum volume required. It needs to be light weight and flexible to be applied directly on the lander/rover external/internal walls, fitted to an electronic box or instrument, wrapped around electrical cables or applied to mechanisms. Furthermore, it needs to be compliant to stringent planetary protection requirements resulting in demanding contamination control measures.

A. Mars Environment and its Implications on Thermal Insulation Mars temperature averages -53°C with variation from -128°C during polar night to +27°C on equator during 1 midday at closest point in orbit to Sun. The chemical composition of Mars atmosphere is dominated by CO2, accounting for 95.3%, with low-percentages of N2 (2.7%), Ar (1.6%), O2 (0.13%) and CO (0.08%). The average surface pressure on Mars is 6.36 mbar at mean radius, variable from 4.0 to 8.7 mbar depending on season.2 Generally heat transfer comprises of three fundamental modes: 1. Conduction, transfer of energy via physical contact, 2. Convection, transfer of energy via physical movement of e.g. gas particles and 3. Radiation, transfer of energy via emission or absorption of electromagnetic radiation. In absence of a transporting media i.e. in vacuum, radiation and solid conduction are the only modes of heat transfer. Conventional vacuum based multi layer insulation (MLI) aims at controlling and adjusting radiative heat transfer and at the same time tries to keep involved conductive heat transfer at a minimum, which works fine in vacuum. On Mars however atmosphere is present, thus gas conduction and convection of the present atmosphere have to be considered as modes of heat transfer in thermal insulation solutions for surface based equipment. As conventional vacuum-based MLI does not offer any measures to suppress conduction of a present gaseous media and is very thin in total, performance will be limited under aforementioned conditions. Thus other means of insulation have to be implemented for planetary landing missions.

B. Thermal Insulation for Mars Surface Missions Previous Mars Landers and Rovers implemented aerogels, fumed silica, different foams and fibrous materials for thermal insulation, just to name a few. For the Viking landers foam, fiber, powder and multi layer insulations were evaluated for thermal insulation. The material selection was mainly driven by the requirements for planetary protection and resulted in the use of several inches of foam insulation and some MLI. Other studies assumed a foam insulation of 3 to 4 inches thickness for potential Mars missions.3

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The same approach was used for MPF lander, which was insulated by a total of 4 inches of 32 kg/m³ polyurethane foam covered with aluminized Kapton film as thermal control surface material.3 The Lander payload electronics box was insulated by a flexible fiberglass insulation blanket with a density of about 16 kg/m3.3 For the Netlanders, although officially stopped in May 2003, three preselected insulation materials were tested in a simulated Mars environment. Melamine resin foam, polyimide foam and fumed silica were investigated, with fumed silica performing 4-5 times better than the foams in Mars environment.4 As thermal insulation for the rover foams, vacuum jacketed enclosures and opacified powders were investigated. Finally a sheet and spar design for the warm electronics box (WEB) was selected with integrated structural and thermal design, composed of a set of E-glass/epoxy structural members, with each volume filled with 25 to 32 mm of 15 to 20 kg/m³ monolithic silica aerogel. Due to the fact that the used aerogel was translucent, a gold-coated 5 mil Kapton film was placed in the middle of the aerogel insulation. For the insulation of cable tunnels polyurethane foam was used.5,6 Both Mars rovers MER-A and MER-B, also known as and , relied on a thermal insulation similar to the one of the Sojourner rover, improving passive thermal control by using carbon-opacified aerogel of 20 to 25 mm thickness.7 For the MSL, also known as , a passive thermal control relying on 1 inch of Martian atmosphere, also known as gas gap, was found suitable with huge benefits for weight and cost.8 But for the majority of the thermal control of the rover during surface operations a mechanically pumped fluid loop is utilized. The main impetus behind this is to use, as far as possible, the waste heat from the radioisotope thermoelectric generator (RTG) to provide heat to the rover in cold conditions.9 The combination of the RTG waste heat and the fluid loop greatly simplifies the rover thermal design in terms of the level of thermal insulation required to maintain the rover and payload at allowable temperatures during cold conditions.9 The planned by JAXA has a gas gap between the component panel and external panels of approximately 60 mm as the baseline, wherein a gas gap separation layer is also inserted.10

C. Reason for a Novel Thermal Insulation for Mars Surface Missions Over the course of previous Mars surface missions a development of thermal insulation from foams and fibrous materials to aerogels, fumed silica and gas gaps is observable. These developments lead to improved thermal performance, but at the same time, made thermal insulation solutions more bulky, less flexible, less adaptable and changed its nature into a structural subsystem. Therefore a novel thermal insulation has been developed, that delivers good thermal performance without the drawbacks of currently known solutions. The developed thermal insulation offers optimal thermal performance but still is of light weight, flexible, able to be applied directly on the lander/rover external/internal walls, fitted to an electronic box or instrument, wrapped around electrical cables or applied to mechanisms. Therefore ESA initiated a study with the objective to develop and characterise high efficiency thermal insulations suitable for Mars surface missions. RUAG Space identified and characterized novel potential insulation materials, developed adequate concepts and designs for thermal insulation utilizing novel materials and validated the thermal performance of application oriented demonstrators in a representative environment.

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II. Development program The development of a novel thermal insulation system for Mars surface Identification of missions followed a multistep approach starting with a literature research potential materials leading to material testing and finally to performance testing of representative demonstrators – see Figure 1. Literature research, A. Identification of potential materials assessment & trade off A literature research was performed, gathering existing knowledge and identifying potential materials and insulation principles. This research yielded fifteen potential materials and insulation principles, which were Testing of potential materials assessed and compared to each other based on literature, material samples and manufacturer data sheets. At the end of this trade off, five out of fifteen potential materials and insulation principles were selected for material Particulate Mechanical testing. cleanliness properties

B. Testing of potential materials Molecular Thermo-optical Evaluated properties during the material testing of the five potential cleanliness properties materials and insulation principles were mechanical and thermo-optical properties, temperature capability, particulate and molecular cleanliness as Thermal Temperature well as thermal performance in 10 mbar CO2 at +25°C. Based on the performance capability evaluated properties three out of five potential materials and insulation principles underwent further thermal performance testing in vacuum, Assessment & 10 mbar N2 and 10 mbar CO2 at +25°C, -15°C and -130°C (not performed trade-off for CO2 due to temperature below freezing point). At the end of the material testing two out of three potential materials and insulation principles were selected. Demonstrators development C. Demonstrators Development For the final two potential materials and insulation principles designs of application oriented three-dimensional demonstrators incorporating Design & manufacturing attachment, grounding and venting provisions were developed and demonstrators manufactured. Demonstrators thermal D. Demonstrators Thermal Performance Test performance test The thermal performances of these two demonstrators were validated by measurements in vacuum, 10 mbar N2 and 10 mbar CO2 at +60°C, +20°C (not performed for CO2 due to test conditions), +5°C, -35°C and -95°C (not Demonstrators thermal performed for CO2 due to cold boundary temperature below freezing point). performance test correlation

E. Demonstrators Thermal Performance Test Correlation A simple mathematical model of the demonstrators thermal Selection of performance test was created and correlated to the test results. Based on the system mathematical model it was possible to estimate the demonstrators thermal performances in a reduced gravitational acceleration environment. Figure 1. Development program.

Details on this development program are presented in the following chapters.

III. Identification of potential materials

A. Literature Research A literature research was performed, gathering existing knowledge and identifying potential materials and insulation principles. This research yielded fifteen potential materials and insulation principles which can be subdivided into the following categories: 4 International Conference on Environmental Systems

1) foams (three candidates with density of 6 – 8 kg/m3 and three candidates with density of 28 – 32 kg/m3), 2) glassy materials (two candidates), 3) nano-structured materials (three aerogels and one fumed silica), 4) stationary gas or vacuum (two candidates) and 5) nano-spaced MLI (one candidate)

B. Assessment and Trade off These fifteen potential materials and insulation principles were assessed and compared to each other based on literature, material samples and manufacturer data sheets. The main assessment criteria were cleanliness, density, flexibility and expected thermal performance. At the end of this trade off the following five out of fifteen potential materials and insulation principles were selected for material testing. 1) High density polyimide foam (HDPIF) was selected due to expected thermal performance in Mars atmosphere. 2) Low density polyimide foam (LDPIF) was selected due to expected low outgassing behavior. 3) Low density melamine resin foam (LDMRF) was selected due to low density and expected good mechanical properties. 4) fumed silica (FS) was selected due to expected excellent thermal performance in Mars atmosphere. 5) concept using stationary gas also known as gas gap (GG), implemented by foils kept at a distance too one another, was selected due to simplicity and expected consequential benefits. Despite the expected exceptionally low thermal conductivity of aerogels in Mars atmosphere, they were dismissed due to poor mechanical properties requiring a secondary structure and problematic cleanliness. Both drawbacks could be taken care of, but this would lead to additional mass as well as increase the overall thermal conductivity of a final thermal insulation. A detailed investigation revealed that the coating planned to be used for the nano-spaced MLI was not structured finely enough to show a similar effect in 10 mbar CO2 as aerogels would do, so this option was dropped. Glassy materials were generally performing below average in the trade off, especially regarding cleanliness, mechanical properties as well as expected thermal performance, and were thus not further investigated on.

IV. Testing of potential materials

A. Material Testing 1. Mechanical Properties Flexible materials HDPIF, LDPIF and LDMRF were investigated regarding compatibility to bending and handling. Measured bending forces over multiple bending cycles and for different bending angles were satisfactorily low for the low-density foams as well as the amount of shed particles during the bending operations. The test revealed that HDPIF is not suitable for bending around edges in the anticipated field of application. FS was not tested as it is not capable of being bent. GG was not tested as it is a concept and the capability of a used thermal insulation to be bent has been sufficiently proven. Densities of all materials were measured and compared to available information from literature and manufacturer data sheets. 2. Operating Temperature Range Temperature capability evaluation was performed on all materials by TGA and thermal shock test by dipping in liquid nitrogen. No material showed considerable mass loss up to +125°C and materials flexible at room temperature remained flexible even in liquid nitrogen, furthermore no influence of the liquid nitrogen exposure cycles on the material structure was observed. Compatibility of the materials with the DHMR process was investigated by heating to +125°C for more than 30 hours, which was withstood without detectable influences, like ruptures, deformation, flaking or discoloration. 3. Molecular Cleanliness Outgassing tests according to ECSS-Q-ST-70-02C11 were performed and all materials showed RML values below 1.0% and CVCM values below 0.1%. 4. Particulate Cleanliness The limit for particulate contamination was 300 ppm with no particle larger than 300 micron valid of the final thermal insulation. Evaluation of cleanliness was performed by tape-lift test directly on the materials which gives an indication on potential contamination levels of the raw material samples compared to each other. Due to the fact that the measured obscured areas of the samples are all separated by roughly one order of magnitude to each other, a qualitative ranking from lowest to highest particulate contamination was possible as follows: GG, LDMRF, LDPIF, 5 International Conference on Environmental Systems

HDPIF, FS. It has to be noted that FS showed problematic cleanliness not only due to the amount of particles, but also due to their size. 5. Thermo-optical Properties Thermo-optical property measurements were performed on all materials, showing emissivity ε between 0.84 and 0.94, only samples LDPIF and LDMRF showed solar transmittance τ of 0.06 and 0.07 respectively all other samples were opaque.

B. Thermal Performance Measurements of thermal conductivity were performed on disc shaped insulation samples using a guarded hot plate calorimeter based on EN 12667:200112. For the measurements five insulation samples were manufactured each based on one material or insulation concept, designed discs with a diameter of 280 mm and a thickness of 10 mm. In the guarded hot plate calorimeter a circular central plate is heated – see Figure 2. The plate is enclosed on either side by two identical samples of known thickness. The outer sides of the samples are coupled to thermally controlled heat sinks. The heating power is regulated in such a way that the temperature on both sides of the sample is constant. The electrical energy in the hot plate flows symmetrically as heat through the two samples. To ensure a one-dimensional heat flow, the central hot plate is enclosed by two concentric guard rings that are kept at the same temperature as the central plate. In order to establish a sufficient thermal contact between the parts and the specimens an external pressure load of about 6900 N/m2 was applied. The disc-shaped samples were measured while in horizontal position. The heat transfer coefficient h (W/m2K) is calculated from the electrical power P (W) fed into the central hot plate, the temperature difference Tint (K) to Text (K) across the two specimens in the stationary state and from 2 the area A (m ) of the central hot plate – see equation 1. The factor 2 accounts for the two specimens. Figure 2. Diagram of a two-plate calorimeter. Using the measured thickness d (m) of the samples, (1) vacuum chamber, (2) hotplate with two guard rings, the thermal conductivity k (W/mK) can be calculated – (3) and (4) cold plate, plate, (5) samples, (6) heat sinks, see equation 2. (7) insulation, (8) ceramic supports, (9) vacuum-tight guided stamp and (10) three thickness sensors. © ZAE Bayern13

h  P 2 ATint Text  (1)

k  Pd 2 ATint Text  (2)

P … main heater power (W) Tint … average temperature of demonstrator inner side (K) h … heat transfer coefficient (W/m²K) Text … average temperature of demonstrator outer side (K) k … thermal conductivity (W/mK) A … measurement area (m²) d … thickness of the insulation (m)

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The thermal conductivities of all disc shaped Table 1. Thermal conductivity of disc shaped samples. Thermal conductivity (10-3 W/mK) insulation samples were screened in 10 mbar CO2 at Parameters +25°C – see Table 1. The thermal conductivity of HDPIF LDPIF LDMRF FS GG GG in vacuum is a result of the distance-keeping 10 mbar CO2, +25°C 18.0 25.3 23.6 7.6 19.8 method chosen to maintain the gap. At this point in 10 mbar CO2, -15°C 15.5 - 18.8 - 17.1 time adequate data was generated by the material 10 mbar N2, +25°C 23.1 - 32.9 - 30.1 tests to exclude the materials LDPIF and FS from 10 mbar N2, -15°C 21.5 - 28.3 - 27.0 further testing. The remaining materials underwent 10 mbar N2, -130°C 12.6 - 17.0 - 17.7 an extensive campaign of thermal conductivity and Vacuum, +25°C 3.2 - 3.4 - 1.7 environmental capability tests in vacuum, 10 mbar Vacuum, -15°C 2.8 - 2.5 - 1.5 N2 and 10 mbar CO2 at +25°C, -15°C and -130°C Vacuum, -130°C 0.9 - 1.0 - 0.6 (not performed for CO2).

C. Assessment and Trade Off of Materials and Performance Testing Based on the performed material tests and the extensive campaign of thermal conductivity and environmental capability tests further trade offs were performed. In a first trade off, after thermal performance screening tests, the field of samples was reduced by LDPIF and FS. LDPIF was dismissed due to overall weaker performance in direct comparison to the other remaining low density foam. Although showing, as expected, excellent thermal performance in 10 mbar CO2, FS was dismissed due to high density, missing flexibility and problematic cleanliness. It has to be kept in mind that implementation of measures against cleanliness issues would not only increase mass and needed volume, but also would reduce overall thermal performance of a final thermal insulation. In a second trade off after test campaign on thermal conductivity and environmental capability, LDMRF was dismissed from further testing due to expected weak thermal performance of a thermal insulation based on it. The remaining samples, HDPIF and GG, were selected for further investigation by three-dimensional thermal performance measurements.

V. Demonstrators Development

A. Design

Based on the two finally selected samples, designs of 1mil VDA/ Mylar/ VDA

Thickness 30 mm Thickness application oriented three-dimensional demonstrators were HDPIF, 10mm thick developed, incorporating attachment, grounding and venting provisions. Tests on these demonstrators were performed to 1mil VDA/ Mylar/ VDA provide relevant design performance for typical configurations. HDPIF, 10mm thick In addition, the demonstrators should prove the capability to 1mil VDA/ Mylar/ VDA manufacture this kind of thermal insulation in a way it would

HDPIF, 10mm thick also be done for a later application case. The shape and the size 1mil VDA/ Mylar/ VDA of the demonstrators should represent a typical configuration, VDA - coating which would be built in a similar manner, and should inside realistically simulate the majority of applications. Figure 3. Composition of HDPIF-3D. Based on this philosophy the demonstrator dimensions have been chosen to be representative for a thermal insulation of a 1mil VDA/ Mylar/ VDA

Thickness 30 mm Thickness lander or rover, being a cube with an internal side length of 300 GG, 10mm thick mm and an insulation thickness of 30 mm. In order to be as 1mil VDA/ Mylar/ VDA representative as possible the following design details were incorporated into the demonstrators: GG, 10mm thick 1) grounding of thermal insulation, 1mil VDA/ Mylar/ VDA

2) attachment techniques, GG, 10mm thick 3) venting provision and 1mil VDA/ Mylar/ VDA 4) closure of inter-thermal insulation contact areas. VDA - coating inside B. Composition Figure 4. Composition of GG-3D. The general composition of the thermal insulation of the demonstrators is shown in Figure 3 and Figure 4.

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The main difference between demonstrators HDPIF-3D and Top GG-3D is the blanket-like character of GG-3D and the tile-like character of HDPIF-3D, which had to be chosen due to the high bending forces of HDPIF. This resulted in more discontinuities and more needed attachment provisions for HDPIF-3D. For the implementation of the GG concept in the corresponding demonstrator, lessons learnt from the thermal conductivity and environmental capability tests were included, which were expected to significantly improve thermal performance. The composition of the thermal insulation of the TC (typ.) demonstrators results in a weight per area of about 1.5 kg/m2 for 2 Bottom HDPIF-3D and about 0.75 kg/m for GG-3D, excluding attachments. Figure 5. 2D-view of TC positions.

C. Temperature Monitoring TCs Each demonstrator was equipped with 26 thermocouples (typ.) (TC) placed at various positions on and within the thermal insulation in order to gain as much information as possible about temperature levels and temperature gradients. For identification of TC positions see Figure 5 and Figure 6. Every position is at least equipped with two TCs, one on the inner and one on the outer side. One position at a face center and one position at an edge were equipped with two additional TCs within the thermal insulation.

Heater

Figure 6. 3D-view of TC positions.

VI. Demonstrators Thermal Performance Test Description

A. Test Setup Tests were conducted at the Mechanical shroud Systems Laboratory of ESTEC using thermal vacuum chamber LAVAF. Heating was demonstrator performed by infrared radiation from a flat plate heater assembly located in the center of suspension the demonstrator. The heater was completely supporting surrounded by the thermal insulation to be m m m tested and the applied power could only be wire frame 0

m heating

0

transmitted through it. The harness of the 0 3

source 5 . 8 heater and the applied thermocouples leaving p thermal p the demonstrator was thermally insulated and a insulation equipped with a guard heater. The test specimens were mounted into the vacuum heater chamber by suspending them to the shroud app. 300 mm with glass yarn – see Figure 7. suspension The cold boundary was provided by the harness temperature controlled shroud of the vacuum chamber and the hot boundary was provided Figure 7. Test setup of demonstrator thermal performance test. by the heating source inside the demonstrator. The measurement uncertainties of the described test setup add up to a total of about 5 %. All surfaces of the demonstrators were thermo-optically representative for the final application. 8 International Conference on Environmental Systems

B. Test Specimens The three-dimensional thermal insulations were supported by a wire frame inside the demonstrator. For the thermal insulation no direct fixation to the supporting wire frame was necessary as the insulation forms a closed volume. Nevertheless, attachment provisions in form of stand-offs were implemented in the demonstrators as well as grounding and venting provisions. Figure 8 and Figure 9 show photos of demonstrators HDPIF-3D and GG-3D before testing, respectively.

C. Test Cases Thermal performance tests were conducted in three atmospheres and at up to five average insulation temperatures, see Table 2. Target temperatures of internal and external faces were also given, but final temperatures varied in the range of ± 10°C depending on thermal insulation performance.

Figure 8. Photo of HDPIF-3D. Table 2. Demonstrator thermal performance test cases.

Tint Text Tavg Atmosphere (°C) (°C) (°C) vacuum 10 mbar N2 10 mbar CO2 +45 +75 +60    +35 +5 +20   - +25 -15 +5    -20 -50 -35    -75 -115 -95   -

For the presentation in the graphs of chapter VII the average temperature Tavg is used, which is the calculated arithmetic mean temperature of the demonstrator average outer temperature Text and average inner temperature Tint – see equation 3. Text and Tint are outside and inside average temperatures of face centers respectively. A weighting factor is needed in the calculation of the average temperatures, as only three face center temperatures are measured but in total they account for four faces – see equations 4 and 5. Figure 9. Photo of GG-3D.

Tavg  Tint  Text  2 (3)

 3  Tint  1 6 4 3 Tint,side Tint,top Tint,bot  (4)  side1 

 3  Text  1 6 4 3 Text,side Text,top Text,bot  (5)  side1 

Tavg … avg. temperature of demonstrator (K) Tint … avg. temperature of demonstrator inner side (K) Text … avg. temperature of demonstrator outer side (K)

D. Measurement Concept The measurement principle applied in these tests is a calorimetric method. A well-known amount of power is applied to a heater inside the demonstrator, while a cold boundary seen from the outside of the demonstrator is set to a fixed temperature. The set-up is then allowed to reach thermal equilibrium. Thermocouples installed on the heater and on the demonstrators external and internal sides allow a measurement of the corresponding temperature 9 International Conference on Environmental Systems

difference. Using a variation of Fourier’s law of thermal conduction, the heat transfer coefficient h (W/m²K) is calculated – see equation 6. The thermal conductivity k (W/mK) is calculated by taking into account the thermal insulation thickness d (m) – see equation 7.

h  P A Tint Text  (6)

k  Pd ATint Text  (7)

h … heat transfer coefficient (W/m²K) P … main heater power (W) k … thermal conductivity (W/mK) d … thickness of the insulation (m) A … measurement area (m²) Tint … avg. temperature of demonstrator inner side (K) Text … avg. temperature of demonstrator outer side (K)

VII. Demonstrators Thermal Performance Test Results & Discussion

A. Thermal Performance in Vacuum Thermal performance of demonstrators in vacuum is very different for HDPIF-3D and GG-3D due to different densities of the thermal insulations. GG-3D showing only half the density of HDPIF-3D and delivering almost half the thermal conductivity of HDPIF-3D – see Figure 10.

B. Thermal Performance in 10 mbar Nitrogen Thermal performance of demonstrators in 10 mbar N2 is comparable for HDPIF-3D and GG-3D – see Figure 11. However, the underlaying mechanism is different. The high density foam of HDPIF-3D reduces heat transport of present atmosphere at the cost of higher conduction due to higher density. The working principle of GG-3D reduces heat transport of solid matter at the cost of offering the present atmosphere a maximum field of activity for conduction, while suppressing the formation of convection. Figure 10. Thermal performance of demonstrators The results clearly show that there is a certain in vacuum. temperature where performance of both thermal insulations is equal. Above this temperature it is preferable to reduce heat transport through the present atmosphere and below this temperature it is preferable to reduce heat transport through solid matter. It can be deduced, that the lower the thermal conductivity of the present fluid, the higher the temperature below which solid conduction of the thermal insulation starts to be of importance. As CO2 gas has a lower thermal conductivity than N2 gas, the temperature below which solid conduction is of importance will be higher in CO2 than in N2. It is easily understood that GG is therefore preferable over HDPIF in CO2. The results clearly show the linear behavior of thermal conductivity over average temperatures.

Figure 11. Thermal performance of demonstrators in 10 mbar N2.

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C. Thermal Performance in 10 mbar Carbon Dioxide In 10 mbar CO2 GG-3D performs clearly better than HDPIF-3D – see Figure 12. Knowing the different underlaying machanisms of HDPIF and GG as well as the results of thermal performance testing in 10 mbar N2 atmosphere this circumstance is as expected. The results clearly show the linear behavior of thermal conductivity over average temperatures.

Figure 12. Thermal performance of demonstrators

in 10 mbar CO . 2

VIII. Demonstrator Thermal Performance Test Correlation

A. Correlation of Vacuum Demonstrator Thermal Performance Tests The thermal performance of the demonstrators in vacuum can be approximated by a superposition of radiative and conductive heat exchange – see equation 8.

4 4 P  GR ATint Text GL ATint Text  (8)

P … main heater power (W) σ … Stefan-Boltzmann constant (5.67E-8 W/m2K4) GR … radiative heat exchange factor (-) Tint … avg. temperature of demonstrator inner side (K) 2 GL … conductive heat exchange factor (W/m K) Text … avg. temperature of demonstrator outer side (K) Tint … avg. temperature of demonstrator inner side (K) A … measurement area (m²) Text … avg. temperature of demonstrator outer side (K) Table 3. GR and GL values for This approach results in an iterative calculation of GR and GL factors, demonstrators in vacuum. determined by approximation of the thermal performance of the HDPIF-3D GG-3D demonstrators in vacuum – see Table 3. Thermal performance of GG-3D in vacuum is almost comparable to GR, - 0.00486 0.00295 the one of a standard MLI containing the same amount of VDA-coated GL, W/m2K 0.00747 0.00416 layers and clearly outperforms HDPIF-3D.

B. Correlation of Atmospheric Demonstrator Thermal Performance Tests To be able to perform an estimation on what the results of the demonstrator thermal performance tests would be in a reduced gravitational acceleration environment as a first step a simple mathematical model is established. In a second step the model is the correlated to the actual test results. Input values for the model are the measured external temperatures of the centers of the different faces of the demonstrators and the temperature of the LAVAF shroud, as well as additional parameters such as pressure, fluid parameters, gravitational acceleration, face orientation and demonstrator dimensions. The value to be correlated to is P (W), which is the power dissipated by the heating source inside the demonstrator at thermal equilibrium. The mathematical model takes into account radiative and convective heat exchange, which is generally expressed in equation 9.

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P  Prad,top  Prad,side  Prad,bot  Pconv,top  Pconv,side  Pconv,bot (9)

P … main heater power (W) rad … radiative conv … convective side … side face top … top face bot … bottom face

The convective heat transfer Pconv (W) is calculated by equation 10.

Pconv  h AText Tfluid  (10)

2 Pconv … convective heat transfer (W) h … heat transfer coefficient (W/m K) 2 Text … avg. temperature of demonstrator outer side (K) A … area of the corresponding face (m ) Tfluid … temperature of fluid participating in convective heat exchange (K)

The heat transfer coefficient h (W/m2K) needs to be calculated via empirical correlations for external free convection flows – see equation 11.14

h  Nu  L (11)

L … characteristic length (m) Nu … Nusselt number (-) λ … thermal conductivity of present fluid (W/mK)

Nusselt number Nu must be calculated for each test case and face orientation based on Rayleigh- and Prandtl- number, for vertical face see equation 12, for top face see equation 13 and for bottom face see equation 14.14

1 2  9 16 8 27  Nu  0.825 0.387  Ra 6 1 0.492 Pr    (12)  

1 4 Nu  0.54  Ra (13)

1 5 Nu  0.52  Ra (14)

Pr … Prandtl number (-) Ra … Rayleigh number (-)

Rayleigh number Ra is the product of multiplying Grashof- and Prandtl-number, whereby the Grashof number Gr is calculated by equation 15.14

3 2 Gr   g Text Tfluid  L  (15)

2 β … volume expansion coefficient (1/K) ν … kinematic viscosity (m /s) 2 Text … avg. temperature of demonstrator outer side (K) g … gravitational acceleration (m/s ) Tfluid… temperature of fluid participating in convective heat exchange (K) L … characteristic length (m)

The radiative heat transfer Prad (W) is calculated by equation 16.

4 4 Prad   AText Tshroud 1  face 1 shroud 1 (16)

2 4 σ … Stefan-Boltzmann constant (5.67E-8 W/m K ) Prad … radiative heat transfer (W) Tshroud … temperature of the LAVAF shroud (K) A … measurement area (m²) Text … avg. temperature of demonstrator outer side (K) εface … emissivity of the corresponding demonstrator face (-) εshroud … emissivity of the LAVAF shroud (-) 12 International Conference on Environmental Systems

The emissivity of the corresponding demonstrator face εface was calculated based on the measurements of the demonstrator thermal performance in vacuum. Based on the formulas above, measured P (W) of the different test cases in 10 mbar N2 and 10 mbar CO2 were correlated by adapting Tfluid (K), the temperature of the fluid participating in convective heat exchange, under the condition that temperatures show a linear behavior over mean temperatures of Text (K) and Tshroud (K) for the different test cases of an atmosphere. Linear approximations of measured and calculated thermal conductivities reveal a very good correlation and bulk fluid temperatures fulfill the requested condition. Based on the correlation performed with gravitational 2 2 acceleration gEarth (m/s ) of 9.81 m/s corresponding to the one of Earth it is possible to shift the performed test to Figure 13. Prediction of thermal performance of Mars gravitational environment by using gravitational HDPIF-3D in 10 mbar CO under reduced acceleration g (m/s2) of 3.71 m/s2 – see Figure 13 and 2 Mars gravitational acceleration. Figure 14.

The predictions of thermal conductivities for Mars environment were generated by changing gravitational acceleration of the correlated mathematical model from 9.81 m/s2 to 3.71 m/s2. Predicted values are not to be considered assured knowledge, but a reduction of thermal conductivity under lower gravitational acceleration is demonstrated.

IX. Conclusion Within an ESA study RUAG Space developed and characterized a high efficiency thermal insulation system suitable for landers and rovers of Mars surface missions. Subsequent to a literature research, a trade off was performed and high-, low-density foams, fumed silica and a gas gap concept were selected for further investigation. Material testing revealed that high density foam and a gas Figure 14. Prediction of thermal performance of gap concept were the most promising candidates for GG-3D in 10 mbar CO2 under reduced demonstrator performance testing. gravitational acceleration. Designs of application oriented three-dimensional demonstrators incorporating attachment, grounding and venting provisions were developed. These demonstrators were then manufactured and their thermal performance validated by measurements in representative environments. Results of these tests reveal an insulation method making use of Mars atmosphere as the best performing option. The developed thermal insulation is very efficient, light weight, flexible and can be applied directly on the lander/rover external/internal walls, fitted to an electronic box or instrument, wrapped around electrical cables or to mechanisms. Performed tests prove maturity of the thermal insulation as a component, being able to be implemented in a system as a viable solution for future Mars surface missions. Based on these results guidelines for thermal insulation useful for other atmospheric conditions can be derived.

Acknowledgments The authors wish to thank the European Space Agency for funding this development within the study “Development of Thermal Insulation for Planetary Landers and Rover”. We would also like to thank the test facility responsible MSL (Mechanical Systems Laboratory) who concurred significantly to the test success. RUAG Space GmbH has been developing and manufacturing MLI for spacecraft and cryogenic application since 1991. 13 International Conference on Environmental Systems

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