Thermal Control of a Light-Weight Rover System in the Permanently Shadowed Regions of the Lunar South Pole

ICES-2020-209

Thermal control of a light-weight rover system in the permanently shadowed regions of the lunar south pole

D. Ivanov1 and D. Fernandes2 Ispace Europe, 5 Rue de l’Industrie, Luxembourg, L-1811

The exploration of the Earth’s Moon has become a topic of great interest in recent years to both private and governmental entities. ispace is aiming to be the enabler for private industry to access new business opportunities on the Moon by capitalizing on lunar resources and expanding our presence in space. The Polar Ice Explorer (PIE) is an in-situ resource utilization (ISRU) exploration mission that aims to find and characterise potential water ice deposits in the lunar polar regions. In the scope of this project, the rover thermal control system development will be discussed. PIE leverages on ispace developed and flight qualified Team HAKUTO’s SORATO rover. The paper explores findings from three key areas: operations in the permanently shadowed regions (PSR) of lunar poles, thermal control design of the rover system and modelling of lunar environment. Thermal modelling of the lunar polar region was conducted with a particular attention towards identification of surface properties, lunar regolith characteristics and modelling of environmental fluxes. Operational mission constraints, such as cooling rates and heater power requirements, were investigated. Thermal design philosophy aimed to maximise passive control means through decoupling of the rover from the ground, reduction of heat losses and conductive path management. Mechanical issues induced by larger temperature swings were investigated. Active control means were considered for elements with tighter operational ranges, such as batteries, motors and externally mounted elements. The rover thermal design challenges and preliminary findings enabling operations in the PSR were outlined.

Nomenclature α = Solar absorptivity β = Infrared emissivity a = surface albedo MCp = System heat capacitance Cp = Specific Heat Capacity GR = radiative coupling GL = linear conductive coupling k = thermal conductivity dt = time step dT/dx = temperature gradient Fij = view factor from surface i to surface j Tsink = Sink temperature h = height g = gravitational acceleration SSM = Second Surface Mirror OSR = optical solar reflector

1 Thermal Engineer, Integrated Mechanical Systems Department, [email protected]. 2 Mechanical Engineer, Integrated Mechanical Systems Department, [email protected].

I. Introduction

The lunar environment, especially polar regions, is characterized by large surface temperature swings throughout each synodic day1 equal to 29.5 Earth days or 708 h. Typical non-illuminated portion of the lunar cycle lasts 354 h. Long night period creates a significant design challenge for surface operations. Surface exploration assets cannot rely solely on solar power to sustain themselves throughout the full lunar night. In addition, The South Pole contains rougher terrain and topographical features with significant heights which cast large shadows2 due to low elevation angle of the sun. Permanently shadowed regions (PSR) are particularly interesting areas for commercial exploration of the lunar surface as they contain trapped volatiles which can be extracted and converted into fuel to enable sustainable mission operations3. However, PSR temperatures as low as 50 to 70 K make it a very challenging thermal task.

Thermal design of the polar rover presents several challenges4. Large swings in temperature cause materials to expand and contract thus creating stresses on mechanical assemblies and interfaces. Minimization of thermal strains requires careful selection of materials with low coefficient of thermal expansion. Operations in the PSR cause embrittlement issues due to reduction of ductility. On the other part of the spectrum, the lunar day may cause severe outgassing issues, which promote rapid degradation of lubricants and enhances abrasive effects of the lunar dust at seals. The lunar dust readily adheres to equipment and outer coatings 5. This not only causes mechanical failures of bearings, wheels and motors but also degraded performance of radiative coatings and reduced power generation of solar cells. These challenges coupled with a large power requirement to survive the lunar night significantly constrained thermal design of the polar rover.

Several lunar landings and operations that lasted longer than one lunar day include Chang’e 4 and Lunokhod missions. Both utilised Radioistope thermoelectric generators (RTG) or Radioisotope heater units (RHU) to survive the lunar night6. On the opposite side, ESA’s Heracles EL3 mission and Blue Origin’s Lander are aiming to bring enough hydrogen and oxygen with them to allow rover or other payloads to take advantage of the regenerative fuel cell system7 and extend mission lifetime to more than 3 lunar cycles by supplying both heat and electricity throughout the non-illuminated portion of the lunar month. Both solutions have their advantages and disadvantages, but fuel cell based systems enable more commercial entities to be involved in the future lunar economy. Overall thermal design trend is to insulate and isolate critical components with tight operational ranges, as well as, add active thermal control solutions like heaters, pumped fluid loops and heat pipes to manage internal environment. Solar panels and RTGs were the two most preferred power sources for shorter and longer missions respectively. The following paper will dive deeper into the thermal design of the rover exploring non- RTG and non-RHU solutions to allow for a feasible commercial alternative. Figure 1: Rover overview

Rover dimensions are 529 x 602 x 372 mm. Total mass of the rover is approximately 10.7 kg excluding the payload. All instruments were accommodated on the internal side of the platform with exception for stereo cameras and antennas. Key objectives of the mission are mapping of the lunar surface and prospecting the area for water ice deposits. Nominal rover mission is expected to last for 1 lunar day or 12 earth days assuming landing two days after sunrise, with no night survival capabilities. However, subsequent design upgrade would be done to improve mission duration and attempt night survival. Overall rover platform is designed to be compact, low mass and energy conservatives. Further chapters of this paper will describe the modelling of lunar environment, explore thermal hardware trade-offs and outline operational requirements for shadows.

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II. Modelling of lunar thermal environment

A. Mission thermal environment

Thermal environment (Infra-red, albedo and solar flux) depends on the mission phase. The mission was therefore divided into several segments: Pre-launch, Leop, Earth Orbit, Transit, Landing and lunar surface operations. The lander ensures sustainable thermal environment compatibility of the rover up-until lunar surface operations phase. The exploration rover will rely on its own thermal control system while operating on the lunar surface. Therefore, key thermal drivers would be moon IR, solar flux at landing latitude and dissipation modes. The rover thermal design utilised a Solar Power Flux with an average value of 1367 W/m2 ± 3.5 %. In addition, ± 1% variation during 11 year solar cycle is also considered, with minimum value of 1321.6 W/m2 (Summer Solstice) and a maximum value of 1412.9 W/m2 (Winter Solstice)8. This environment is applicable at all times except when the Sun is blocked (i.e. during eclipse). Sink temperature to space was taken as 2.53K. However, this mainly applies to the top surface with radiators. Surfaces with a view factor to lunar ground used an adjusted sink temperature.

Several important orbital parameters were included in order to accurately model the lunar environment. The lunar equatorial plane is inclined by 1.53o to the sun’s ecliptic plane and by 6.68 o to the orbital plane around the earth. The ecliptic plane is inclined by 5.15 o to the lunar orbital plane. In comparison, the earth is inclined by by 23.28 o to sun’s ecliptic plane. These parameters determine general motion and view conditions between the moon, the earth and the sun in the solar system9. Distance to sun was chosen to be 1 A.U, which is approximately true due to close vicinity to Earth.

Figure 2: Mission thermal environment definition

B. Topographical Elevation

Topographical map of the moon, taken from the Kaguya and LRO data, shows a typical distribution of highlands, craters and mares on the surface. The sun elevation angle does not exceed 10 degrees at poles and would therefore cast large shadows when illuminating any boulder, crater rim or mountains. Identification of these key topographical features was critical in defining the mission trajectory and designing the thermal control system of the rover9. Further assessment of these features is shown in the next chapter.

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Figure 3: Lunar topographical elevation a) global 2D map, b) global 3D map

C. Solar Elevation Angle

Solar elevation angle was investigated for a proposed landing site in the Amundsen Region. One full revolution with respect to stars (sidereal month) was taken as 27.32 Earth Days. The time between two new Moons (synodic month or lunation) was taken as 29.53 days. Longest solar day in 2023 was estimated to be 12 days, 1 hour, 10 minutes starting at 2023-01-25 00:12:27 (± 45 minutes). Solid red lines represent solar elevation angles over a period of 1 year. Solid blue line represents topographical elevation angles 10. Blue background represents the estimation error.

Figure 4: Solar elevation angle plotted against topographical elevation in the Amundsen Region

D. Environmental Fluxes

The rover design considered a Lunar Albedo with a minimum value of 0.076 and a maximum value of 0.297, with an average value of 0.18. The thermal infrared input from the Moon was factored into the design of the rover by considering its view factor to the surroundings with a cold/hot range of 0 to 140 W/m2. Total flux was still driven by the direct solar radiation during the illuminated portion of the mission, and by lunar infra-red radiation during the eclipse. Albedo flux ranged from 0 to 30 W/m2. Earthshine was taken as 0.111 W/m2 at equator and 0.019 - 0.071 W/m2 at poles. Micrometeorite flux from impact is approximately 4.1E10-6 mW/m2 (negligible) 8.

The thermal control system has to be designed to allow the rover to survive extended charging times, as well as, operate in any orientation or any illuminating condition for at least two hours (troubleshooting operation time in safe mode). During the transportation phase from Earth to the Lunar Surface, the rover’s temperature is managed by the lander which guarantees operational thermal control of the asset within the internal payload bay. However, upon 4 International Conference on Environmental Systems

opening the payload bay the rover temperature would swing between +6 oC to +48 oC at mid-latitude, and –13 oC to +26 oC in polar regions.

The polar ice explorer mission is intended to land on the Moon two days after the local sunrise, which would allow for the maximum amount of sun illumination to be present during deployment and initial operations. Externally mounted instruments (such as ground penetrating radars) with a temperature range between –90 oC and +115 oC would require a careful selection of mounting interface coupling during the hot phase of the mission to avoid excessive heat exchanges with the rover body. On the other hand, internal instruments with a tighter temperature range between -20 oC and +25 oC would benefit from bottom mounting position and decoupling them from high dissipative electronic units to keep their temperature stable during all operational phases. Similarly, during the lunar night, it is critical to keep sensitive instrument within the internal compartment of the insulated rover body.

E. Lunar Surface temperature model

Lunar surface was represented using a two-layer Thermal Desktop model 11 as shown in Figure 4. Heat flux from lunar interior was taken as 0.031 W/m2 ± 20%. The top layer represents lunar dust with a depth of 0.02 m. The bottom layer repsents lunar soil with a depth of 0.60 m. Temperature dependent conductivity model was applied to each layer with values valid from -173 oC to 125 oC. Dust layer conductivity varied from 0.0000954 to 0.00296 W/m/K. Soil layer conductivity varied from 0.00946 to 0.0194 W/m/K. Similar thermal heat capacity was applied to both layer with a value of 1050 J/kg/K. Density was taken as 1000 kg/m3 for dust and 2000 kg/m3 for soil layer. Only top surface was made radiatively active with 0.82 solar absorptivity and 0.97 IR emissivity parameters. The interface between two layers was modelled using a linear conductor (GL) with an equivalent conductivity and path length of 0.01 m. Nodal distribution was selected to represent central (N10000), external (N30000) and middle (N20000) section of the surface. Total surface radius was taken as 30.0 m. The inner section has a radius of 1.0 m and was made with a finer mesh to capture interaction of rover 12 with the surface Figure 5: Lunar Subsurface Model

Figure 4 displays the subsurface temperature model correlation results with respect to Apollo 15 data. The peak temperatures measured during Apollo 17 mission 13 were between 375 to 384 K. In comparison, Thermal Desktop produced a peak value of 379.8 K, which corresponds to a 4.8 K delta with respect to peak Apollo measurement. The minimum temperatures measured during Apollo 17 mission were between 94 to 102 K. Thermal Desktop model produced a minimum temperature of 101.2 K, which corresponds to a 7.2 K delta with respect to a colder Apollo measurement. Overall, Thermal Desktop model performed adequately and generated results within acceptable accuracy.

Equatorial region of the moon is characterized by average surface temperatures of 94.3 ± 2 K during non- illuminated portion of the lunar cycle and 392.3 ± 5 K during illuminated portion. Lunar equator extreme minimum and maximum temperatures range from 92 K to 397 K. Average temperature swing is approximately 290 K. For thermal modelling purposes a boundary condition of -180 oC can be used for shadowed region and 123 oC can be used for illuminated region 14.

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Figure 6: Correlation of surface model with Apollo 17 Data

Polar region of the moon is characterized by average surface temperatures of 75 ± 5 K during non-illuminated portion of the lunar cycle and 230 ± 20 K during illuminated portion. Lunar south pole extreme minimum and maximum temperatures range from 50 K to 221 K. Average temperature swing is approximately 120 K. For thermal modelling purposes a boundary condition of -205 oC can be used for shadowed region and -20 oC can be used for illuminated region9. From Figure 6 it can be seen that the temperature below the lunar surface is almost isothermal below the initial 2 cm of dust and 8 cm of soil. While surface experiences nearly 120 K swings at the South Pole, the subsurface temperature remains constant between 165 K and 180 K. Temperatures vary depending on the illumination conditions throughout each synodic month.

Figure 7: Amundsen region (-82S, 66E) surface and subsurface temperature variation 6 International Conference on Environmental Systems

F. Crater, boulder and rock IR impact on Rover design

Topographical impact assessment 15 was performed to understand how IR flux emitted by boulders and rocks affects the operational temperature of the rover. Assessment was performed by initially varying the size of the boulder from 0.25 m radius to 4.00 m radius at a fixed distance from the rover. Optical properties of the boulder were similar to lunar regolith: absorptivity of 0.82 and emissivity of 0.97. Second test fixed the boulder size to 2.00 m radius and varied the distance to the rover. The distance was varied from 0.50 m to 20.0 m. Average albedo value of 0.18 was taken as a reference. Separate assessment was performed to see the thermal difference between lunar mare albedo of 0.076 and lunar highlands albedo of 0.297.

Figure 8: Boulder and Rock IR Assessment

Construction of geometry was performed in Thermal Desktop. Thermal capacity was taken from lunar mining handbook with an average value of 1050 J/kg/K (940 to 1120). The conductivity of material was taken as an average value of 1.7 W/m/K (0.9 to 2.5). Density was dependent on the geology of a particular landing site. Generally, it varied from Lunar Mare Basalts (3010 ± 40 to 3460 ± 30 kg/m3), to Feldspathic Highlands (2510 ± 40 to 2840 ± 40 kg/m3) and Impact breccias (2360 ± 40 to 2520 ± Figure 9: Modelling of Lunar craters 30 kg/m3). Due to large elevation profile in the Amundsen Region, the Feldspathic highlands density value was applied to each feature during thermal assessment 16. Table 1 below outlines key findings with respect to boulder, rock and crater IR impact on exploration rover temperatures. Reference temperature was taken from a flat surface simulation. Reduced rover thermal model was used to perform analysis. The frame was covered with an MLI and AgTeflon coating was used for radiator.

Table 1: IR impact of lunar features

Lunar feature Delta = Trover (reference) – Trover (feature) T min (oC) T max (oC) Boulder (R = 4.0 m) +3.2 +3.4 Rock (R = 0.25 m) +0.1 +0.4 Boulder (s = 0.50 m away) +3.8 +4.8 Boulder (s = 20.0 m away) +0.2 +0.3 Albedo (0.076 to 0.293) +0.10 average +0.60 average or +3.50 local Crater -0.60 -3.50

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III. Operational requirements for permanently shadowed regions (PSR)

Key thermal factors associated with survival in the shadowed regions include operational limits of electronic units, available heater power, insulation and transient behavior of the system17. Three use cases were evaluated to determine initial requirements for operations in the shadowed regions. Case one included operational traverses and data collection within PSR assuming only battery power. Case two was focused on extension of useful mission time using regenerative fuel cell system as an electrical and heat supplement. Case three focused on night survival capabilities and the use of RFCS as an emergency backup system. Table 2 provides a summary of rover’s transient behavior:

Table 2: System transient behaviour modelling

Parameters Value System MCp (excl. payloads), J/K 9,430 System MCp (incl. payloads), J/K 15,280 Thermal Time Constant, h 3.3 3*Tau (Time to reach 95% of Tss), h 9.9 Average flux, W/m2 187.0 Approximate temperature gradient, K / h 8.10

Current battery is a lithium-ion cell with a capacity of 86 Wh. Total discharge time is taken from a saturated state of 70 ± 2 Wh to DOD at 44 Wh. Assuming a 40% depth of discharge limit (DOD), nominal case without heaters allows for 139 minutes of operations. With heaters, the operational time drops to 65 minutes (15.0 W of heater power). Charging cycle is roughly 143 minutes at 11.0 W. Regenerative fuel cell system has an energy density of 800 Wh/ kg with a 25% electrical output efficiency. Reference RFCS system can supply 5 Wh of capacity for 40 hours.

Table 3: Mission impact with and without RCFS system

Case Description Time to Operate until DOD Combined system with RFCS Case 1: Eclipse operations 139 min (assumes pre-conditioning in the 223 min sun with 5.0 W battery heater) Case 2: Operational time extension 143 min (charging during a sun-pointing 98 min (reduction of charge (reduction of battery charge time) mode) time by 45 minutes) Case 3: Night Survival 61 min (operational mode) 720 min (safe mode)

Night survival is a challenging endeavor without a considerably large power source like an RTG or RHU. Around 15 to 40 W of continuous heating power is needed throughout 354h of night. Nevertheless, RFCS can be used as a backup power source during emergencies and in the cold regions of the moon 18. For initial mission only a one-way fuel cell would be feasible for the rover platform, due to weight and size restrictions. However, in future the hydrogen and oxygen source could be transported by the lander and recharged via solar powered electrolyser.

Figure 10: Electronics thermal assessment (lunar night) 8 International Conference on Environmental Systems

Thermal assessment was performed with a variety of hardware options. Mainly the trade-off consisted of several static radiator surfaces, heat switches, deployable panels and louvres. Thermal case with no radiator was also included for reference. Majority of components on-board the rover, like infrastructure boards, communication modules and autonomy units have an operational temperature range between -40 oC to +60 oC. Battery has a tighter range from -5 oC to + 45 oC. However, battery’s actual operating range should be between -5 oC to +10 oC and it would therefore require a more refined thermal control solution. For the following assessment, the rover was kept above -20 oC during heater assessment and an upper limit of +40 oC was used to represent the desired hot limit14. Table 4 below summarises all hardware cases which were evaluated during the study. Blue colours represent cold temperatures, yellow colours represent nominal temperatures and red colours represente hot temperatures.

Table 4: Thermal hardware assessment results

Thermal Thermo-optical finish No Internal w/o night with night Power Hardware power heaters heaters Requirement Assessment Generation (o C) (o C) (o C) Solar IR T min T max T T max T min T max W Absorptivity, Emissivity, min α ε No Radiator 0.050 0.050 -190.0 30.7 -44.6 84.5 -19.3 84.6 5.7 AgTeflon 0.099 0.750 -203.3 -11.7 -77.1 34.5 -19.2 34.6 20.1 SSM 0.100 0.900 -205.1 -17.7 -81.6 27.2 -19.1 27.2 23.2 White Paint 0.240 0.830 -204.3 -11.4 -79.6 32.6 -19.1 32.6 21.7 Heat Switch 0.099 0.750 -203.1 -0.8 -73.0 34.4 -19.2 34.4 15.9 Deployable Radiator 0.050 / 0.050 / 0.970 -187.0 -17.7 -44.6 15.4 -19.8 17.5 5.6 (MLI/Black 0.910 Paint) Deployable Radiator 0.050 / 0.050 / 0.820 -187.0 -18.5 -44.6 15.4 -19.8 15.8 5.6 (MLI/White 0.850 Paint) Deployable Radiator 0.050 / 0.050 / 0.240 -187.0 -31.5 -44.6 7.2 -19.8 7.2 5.7 (MLI/Solar 0.830 Cell) Deployable Radiator (MLI/White 0.050 / 0.240 0.050 /0.830 -186.9 -12.8 -44.5 7.5 -19.4 7.5 5.6 Paint + Heat Switch) Louvre 0.170 0.100 -191.4 0.0 -47.4 37.6 -19.3 37.7 6.5 Louvre + Heat (close) / 0.170 -191.3 0.0 -47.2 37.8 -19.3 37.8 6.4 Switch 0.750 (open)

The on/off temperature set-point for active thermal control system was set at –5.0 oC. Static radiator coupling to the rover body was modelled using a large GL value of 1500 W/m2/K since this value was not driving during thermal assessment. Louver open/close effective emissivity was varied from 0.750 (open) to 0.100 (close) state. The dynamic property was simulated using Bij*A*e coupling to a space node with temperature dependent logic in thermal desktop18. Heat switch conduction was set to 8.50 W/m2/K open and 700.00 W/m2/K close state19. Deployable radiator was set to open during eclipse entry and close during eclipse exit. Motion was simulated using articulating bodies14.

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The power dissipation during operational mode (excl. motors) was set to 32.6 W and safe mode power dissipation was set to 11.2 W. Each motor of the four-wheeled mobility system dissipated 4.0 W of heat during dynamic operations 17. Heat dissipation during operational, or dynamic, mode included the infrastructure layer (housekeeping, power, communications), high level functional layer (autonomy), payloads and motors. Safe mode dissipation mainly includes the infrastructure layer while keeping the rest of components in the OFF state and rover in static position.

Figure 11: Temperature profiles of rover with various hardware configurations

IV. Thermal Control of the Polar Rover

Thermal architecture philosophy of the rover was to maintain battery, electronic units and payload temperatures within their operational limits throughout the entire mission lifetime maximizing passive means of thermal control. Average temperature within the rover was kept between -20 oC to +50 oC. Active thermal control measures were implemented to regulate the battery with tighter operational temperature range. Further heaters were implemented for electronic components (motor controller, communication board, main computer, FPGA and autonomy unit). Payloads were mounted on the -Z plate of the rover with addition of a thermal filler. Electronic unit thermal control was achieved through conductive decoupling measures, louvre assembly and a heated warm electronic box. Internal compartment was painted black to facilitate radiative thermal exchange between different surfaces and reduce localized thermal gradients. Based on the initial thermal assessment and cold environment at the South Pole of the moon, it is not feasible to rely on a purely passive radiator solution without RHU or RTG power sources. Decoupling of the radiator during non-illuminated portion is needed. Therefore, a louvre solution scaled to the rover form-factor would provide the necessary change in IR emissivity and reduce effective radiative surface to deep space. Selection of structural materials and its bulk properties was based on the rover and lunar surface decoupling concept. Driving motors dissipated significant amount of heat during operations and were therefore thermally insulated from the main body.

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Figure 12: Polar rover thermal modelling

Thermally actuating passive mechanical actuators, such as bimetallic springs, are being investigated. No more than 70% of maximum installed heater power is expected to be used at any point of time to avoid saturation. Rover size fuel cell system7 is currently being researched and developed as an attractive commercial solution to increase effective mission duration during commercial lunar operations.

Figure 13: Preliminary thermal architecture of the Polar Rover

All exposed areas of the rover were covered with multi-layer insulation blankets. Specifically, surfaces would be covered either with Aluminium Polyimide outer layer MLI (10 – 15 layers) or with Beta Cloth, to achieve uniform heat exchange environment and keep the rover temperature stable in non-illuminated conditions. Thermal decoupling measures were considered to limit heat loss to the lunar surface. Hence, the wheels were made out of an insulating material called ULTEM offering a conductivity value of 0.22 W/m/K. Further design and analysis work was conducted to assess the rover architecture and would be presented in the next phase of the project.

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Acknowledgments We thank ispace Europe and ispace Japan team for making Polar Ice Explorer project possible; We thank Dr. Abigail Calzada, Daniel Bolan, Dr Carlos Espejel and other members for contributing material towards lunar modelling activities; In addition, we thank Hyun-Ung Oh and Tae-Yong Park (Space Technology Synthesis Laboratory) for their inputs on night survival modelling in Thermal Desktop.

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