Advanced technologies are sought for thermal management of Earth-orbiting spacecraft, the human lunar habitat, landers, and rovers, for Martian transit spacecraft, as well as planetary expeditions to Jupiter, Venus, and their moons. Future spacecraft will require more sophisticated thermal control systems that can dissipate or reject greater heat loads at higher input heat fluxes while using fewer of the limited spacecraft mass, volume, and power resources.
The thermal control designs also must accommodate the harsh environments associated with these missions including dust and high sink temperatures. Modular, reconfigurable designs could limit the number of required spares.
The lunar environment presents several challenges to the design and operation of active thermal control systems. During the Apollo program, landings were located and timed to occur at lunar twilight, resulting in a benign thermal environment. The long-duration polar lunar bases that are foreseen in 15 years will see extremely cold thermal environments, as will the radiators for Martian transit spacecraft. Long sojourns remote from low-Earth orbit will require lightweight, but robust and reliable systems.
Innovative thermal management components and systems are needed to accomplish the rejection of heat from lunar bases. Advances are sought in the general areas of radiators, thermal control loops, and equipment. Radiators on the Moon’s poles and on a Martian transit vehicle are required that will operate and survive in very cold environments. Variable emissivity coatings, clever working fluid selection, or robust design could be used to prevent radiator damage from freezing at times of low heat load. Also, the dusty environment of an active lunar base may require dust mitigation and removal techniques to maintain radiator performance over the long term.
The lunar base and Martian transit spacecraft active thermal control systems will include high-efficiency, long-life mechanical pumps. Part of the thermal control system in the lunar base is likely to be a condensing heat exchanger, which should be designed to preclude microbial growth. Small heat pumps could be used to provide cold fluid to the heat exchanger, increasing the average heat rejection temperature and reducing the size of the radiators.
Thermal Management in Space
The vehicles and habitats associated with space industrialization and the exploitation of non-terrestrial resources will inevitably require energy systems far exceeding the current requirements of scientific and exploratory missions. Because of the extended duration of these missions, it is not possible to consider systems involving expendables such as non-regeneratable fuel cells. Therefore, these missions become hostages to the capability of continuous-power energy systems. These systems will need to provide hundreds of kilowatts to tens of megawatts of electrical power to a product fabrication system, whether it uses terrestrial or non-terrestrial raw materials.
EVA Thermal Environment
For past manned missions, the EVA thermal environment has been a major design driver for the EVA spacesuit, starting with the thermal protection requirements of the anthropomorphic spacesuit garment and extending to the primary life support system (PLSS) on the spacesuit. The thermal environment influences passive system design, such as material selections and infrared/solar coating selections, as well as active system design, such as sublimators, radiators, liquid cooling systems, and crewmember metabolic expenditure for performing critical EVA tasks.
Along with the thermal environment, metabolic heat load and metabolic rate predictions also affect equipment designs and consumables. Determining accurate metabolic heat loads is a challenge since it is closely tied to the outside thermal environment. An understanding of the impact of metabolic heat load first requires a review of historic space program metabolic expenditures. Also affected by the thermal environment is spacesuit heat leak, which in tum affects insulation requirements and thermal consumables of the spacesuit. A review of past mission heat leak effects as a function of the environmental heat load is important to quantify expected heat leaks on upcoming lunar EVAs. Thermal environments can also affect and degrade spacesuit radiator performance in the presence of both infrared and solar heat fluxes. In evaluating methods to improve spacesuit radiator performance on the lunar surface, the steady-state heat rejection performance is presented for multiple spacesuit-mounted radiator concepts at various solar angles and operating temperatures. The radiator concepts use shades to provide varying degrees of protection from the hot lunar dayside surface.
Thermal management technology is an essential requirement for all spacesuits, spacecraft, and space habitats. During extravehicular activity (EVA), spacesuits must remove metabolic heat produced by the astronaut, residual heat from the suit’s electronics, and absorbed heat from the external environment. The contribution of each of these heat sources can vary considerably, meaning a thermal control system must be able to actively adapt to changing heat loads to keep the suit’s internal temperature within the narrow band safe for human occupation. Currently, both NASA’s Extravehicular Mobility Unit (EMU) and the Russian Orlan spacesuit use water-ice sublimation to provide cooling. The ISS EMU utilizes a base layer called the Liquid Cooling and Ventilation Garment (LCVG) made of nylon (i.e. a synthetic thermoplastic linear polyamide; a large molecule whose components are bound by a particular type of bond) and spandex (i.e. a type of stretchy polyurethane fabric) interlaced with flexible rubber tubing for pumping cool water across an astronaut’s skin to absorb heat. Liquid in a feedwater circuit is exposed to a heat exchanger in the Portable Life Support System (PLSS) backpack, creating a thin sheet of ice that sublimates directly into the vacuum of space. Thermal energy is dissipated as the latent heat of sublimation from the water mass, cooling the water in the separate Liquid Transport Circuit that flows through the LCVG. This method has proven reliable, but there are many problems and opportunities for optimization.
Planetary EVA Environments
Next-generation spacesuits will need to function in a dynamic range of environmental conditions. Heat sources include environmental radiation, possible atmospheric effects, human metabolic heat, and residual heat from suit electronics. The lunar environment is characterized by approximately ⅙ Earth g, and surface temperatures ranging from -143°C (-225°F) to +127°C (+261°F) due to direct sunlight or shading in an ambient vacuum. In direct sunlight, radiation can be reflected from the lunar regolith in the infrared (IR) range, amplifying the radiative heating. In deep shadowed craters or during the extended night of the lunar 28-day day-night cycle, temperatures can approach absolute zero. Temperatures are also dependent on latitude and local lighting conditions. During a lunar EVA, a crew member will experience sharp transient effects as they move in and out of shadows and change their orientation with respect to the sun.
Human metabolic heat can fluctuate dramatically depending on activity level. The metabolic heat from a single lunar EVA can vary from 70 to 730 W depending on activity, with extreme peaks of 880 W in emergency contingencies. This heat can be broken down into two categories: sensible heat and latent heat. Sensible heat is transmitted through the skin to the environment via conduction and convection and generally remains stable over varying activity levels. Latent heat refers to the heat rejected from perspiration and expired water vapor. Latent heat increases rapidly with activity level. The current LCVG and EMU alter this relationship by adjusting to keep skin temperature constant even at higher work rates so that sensible heat continues to be the dominant heat transfer mechanism. A system that relies on the body’s natural transition to latent heat rejection at higher workloads may be more efficient and reduce consumables but will also require more drinking water. It may also require a new mechanism for collecting and filtering latent moisture condensate. PLSS electronics heat will likely decrease with technological advancements but generally runs at a steady 120 W. The current requirement for NASA’s next EVA spacesuit, the Exploration EMU (xEMU) is 100 W. In total, this means that a spacesuit must be able to reject between 190 and 850 W. The current EMU has a requirement baseline value of approximately 250 W in steady-state performance for an 8-hour EVA. The Apollo EMU requirement included about 270 W of cooling for an 8-hour EVA with a 19°C internal temperature maintenance, and a 250K heat sink temperature. This represents a conservative baseline for future suit designs.
Spacesuit Water Membrane Evaporators
The PLSS built into the xEMU will include several updates from the original Apollo and Shuttle era suits, including the integration of a Spacesuit Water Membrane Evaporator (SWME). The SWME utilizes 27,000 thin-walled, hollow, 300 μm diameter polypropylene fibers (PPF; a kind of linear polymer synthetic fiber obtained from propylene polymerization) that provide about 1.1 m2 open pore. The small pores in the fibers prevent liquid from leaking out while presenting very low resistance to evaporation and vapor flow. The circulating water evaporates due to the low water vapor pressure on the outside of the SWME. The evaporation cools the remaining liquid as it returns to the LCVG. This keeps all contaminants in the cooling line and prevents them from clogging the sublimator, but a periodic scrubbing of the water circuit will be required instead. The heat flux is controllable, and the SWME is pressure independent at start-up, operating over the whole range of suit pressures. The SWME can also degas the thermal loop water to prevent pressure build-up. Testing began in 1999 at the Johnson Space Center, with the first full sheet demonstration in 2009. The Gen2 SWME was created in 2010 and has demonstrated a capacity of greater than 800 W of cooling capacity with over 200 hours of testing. The SWME remains a mass consumable life support system, a limitation for its extensibility to Moon or Mars missions.
A critical advancement has been the development of the Space Evaporator-Absorber-Radiator (SEAR), which utilizes a Lithium-Chloride Absorber Radiator (LCAR) to condense and absorb water vapor evaporated through the hollow fibers of the SWME. The lithium-chloride solution is a strong desiccant (i.e. a hygroscopic substance used as a drying agent) that maintains a very low vapor pressure even at relatively high temperatures. The heat generated in this process drives up the temperature of a connected exterior surface radiator panel that rejects heat into space. Lithium-chloride solution maintains a low vapor pressure, even at temperatures that are typically 30°C higher than the SWME temperature. This enables the LCAR to operate as a heat pump, rejecting heat at temperatures of about 50°C-60°C which significantly reduces the size of the radiator. The LCAR designers at Creare LLC expect the system to be able to reject 150 to 300 W depending on available radiator surface area and environmental conditions. The lithium-chloride concentration begins at about 95% and slowly dilutes to 45% over the course of an 8-hour EVA, at which point the LCAR ceases to be effective. The absorbed water can be recovered after the EVA by baking the LCAR at 100°C-120°C, which is a high enough temperature to evaporate the water from solution. This could be done using a regeneration oven inside the spacecraft or habitation module. A comparison of the independent SWME and SWME with LCAR systems.
The most common proposal for solid-state heat exchangers is thermoelectric coolers (TECs). A TEC is a solid-state heat pump made of thermocouples (i.e. sensors for measuring temperature) that creates a temperature difference across the device when a current runs through it in a process called the Peltier Effect (the Peltier effect is the reverse phenomenon of the Seebeck effect). When a current flows through a junction between two different conductors (from a battery), heating and cooling occur because charge carriers diffuse to opposite sides of the device. Changing the polarity of the current reverses the flow direction and switches the temperature gradient. A TEC also has no moving parts, giving it the advantage of simplicity.
Traditionally, TECs have been used for cooling small, energy-dense electronic components due to scaling challenges associated with power and mass. Most research involving heat exchangers for PLSS has focused on rejecting the residual heat from suit electronics, including a new additively manufactured titanium-based liquid-gas heat exchanger for cooling the breathing gas ventilation loop in the next generation EMU. At only 0.39 kg, this small component is less than half the mass of previous prototypes and can reject a higher heat load. Mass reductions will be a key component of future PLSS components, and this demonstrates the potential for future optimization of hardware components. While heat exchanger efforts have focused on small components, there have been several concept studies for using TECs as the primary cooling mechanism for a spacesuit. Preliminary research using analog EVAs on Earth has shown the potential to provide superior cooling with a lower system mass by using TECs and swappable batteries rather than replaceable ice block coolers in analog spacesuits. Benchtop tests indicate that using five lithium-ion batteries totaling 3.4 kg to power a TEC cooling system would greatly reduce the 60 kg of replacement ice reservoirs needed to keep an analog suit cool over the course of a 175 W, 6-hour EVA. The applications of this work have so far been limited to terrestrial analog EVAs, but future work will examine the feasibility of using the Peltier cooling concept for EVAs on Mars.
Radiators have been a component of each of the candidate technologies presented so far, but it may be possible to provide cooling using only a passive process. Spacecraft and robotic explorers already use radiators for heat rejection. Radiators reject residual heat via infrared (IR) radiation emitted by a surface exposed to a heat sink with a much lower temperature. These systems are “closed-loop” with respect to mass. Radiators typically provide a steady level of cooling, making them suboptimal for adapting to the dynamic heat loads of a spacesuit application. Additionally, surface area constraints on rigid structures such as the PLSS backpack limit the amount of cooling possible. However, recent technological developments have reopened the possibility of using a radiator-based approach for EVA thermal control. Early work investigating the application of passive radiators for zero-consumable heat rejection led to the Chameleon Suit architecture, proposed in the early 2000s.
With this approach, the thermal control system was envisioned to allow for variable conductance (i.e. the amount/speed of heat transmitted through a material; heat transfer occurs at a higher rate across materials of high thermal conductivity than those of low thermal conductivity) between the skin surface and the surface of the suit using different lofting techniques. The surface of the suit allowed for further adjustment of the heat flow using variable electrochromic radiators (electrochromism is the phenomenon where the color or opacity of a material changes when a voltage is applied) along with mechanical modulating of the radiators using MEMS (microelectromechanical system) louvers. Technological advancements in the succeeding decades have made such a concept more viable, but it is still limited by an inability to remove heat when the sink temperature is close to or greater than the desired skin temperature, and it still requires battery power to operate. Electrochromic materials may help the problem of environmental temperature variation by allowing for a variably emissive surface. The optical properties of electrochromic films can be controlled electrically using an applied voltage. New technology has also allowed these devices to be installed on a flexible substrate such as Kapton tape or polymer film. When a small bias voltage is applied, the material undergoes an oxidation-reduction reaction, and the surface’s IR emission properties change. The layers of an electrochromic device (ECD) function similarly to the anode, cathode, and electrolyte in a battery.
Heat Rejection Systems
Essentially the system is a small vapor cycle (a thermodynamic cycle, operating as a heat engine or a heat pump, during which the working substance is in, or passes through, the vapor state) that uses the temperature difference between the hot and cold ends of the tube as a pump to transport heat, taking full advantage of the heat of vaporization of the particular fluid.
The fluid must be carefully selected to match the temperature range of operation. For example, at very high temperatures a metallic substance with a relatively high vaporization temperature, such as sodium or potassium, may be used. However, this choice puts a constraint on the low-temperature end since, if the fluid freezes into a solid at the low-temperature end, the operation would cease until the relatively inefficient conduction of heat along the walls could melt it. At low temperatures, a fluid with a low vaporization temperature, such as ammonia, might well be used, with similar constraints. The temperature may not be so high as to dissociate the ammonia at the hot end or so low as to freeze the ammonia at the cold end.
With proper design, heat pipes are an appropriate and convenient tool for thermal management in space systems. For example, at modest temperatures, the heat pipe could be made of aluminum, because of its relatively low density and high strength. Fins could be added to the heat pipe to increase its heat dissipation area. The aluminum, in order to be useful, must be thin enough to reduce the mass carried into space yet thick enough to offer reasonable resistance to meteoroid strikes.
A very carefully designed solid surface radiator made out of aluminum has the following capabilities in principle: The mass is approximately 5 kg/m2 with an emissivity (i.e. the ratio of the energy radiated from a material’s surface to that radiated from a perfect emitter, known as a blackbody, at the same temperature and wavelength and under the same viewing conditions) of 0.86; the usable temperature range is limited by the softening point of aluminum (about 700 K). At higher temperatures, where refractory metals (i.e. a group of metallic elements that are highly resistant to heat and wear) are needed, it would be necessary to multiply the mass of the radiator per square meter by at least a factor of 3. Nevertheless, from 700 K up to perhaps 900 K, the heat pipe radiator is still a very efficient method of rejecting heat.
A further advantage is that each heat pipe unit is a self-contained machine. Thus, the puncture of one unit does not constitute a single-point failure that would affect the performance of the whole system. Failures tend to be slow and graceful, provided sufficient redundancy.
Pump Loop System
The pump loop system has many of the same advantages and is bounded by many of the same limitations associated with the heat pipe radiator. Here heat is collected through a system of fluid loops and pumped into a radiator system similar to conventional radiators used on Earth. It should be pointed out that in the Earth’s environment the radiator actually radiates very little heat; it is designed to convect its heat, i.e. transport heat by convection. The best-known examples of the pump loop system currently used in space are the heat rejection radiators used in the Shuttle. These are the inner structure of the clamshell doors which are deployed when the doors are opened.
Pump loop systems have a unique advantage in that the thermal control system can easily be integrated into a spacecraft or space factory. The heat is picked up by conventional heat exchangers within the spacecraft, the carrier fluid is pumped through a complex system of pipes extended by fins (i.e. surfaces that extend from an object to increase the rate of heat transfer to or from the environment by increasing convection) when deemed effective, and finally, the carrier is returned in liquid phase through the spacecraft. In the case of the Space Shuttle, where the missions are short, additional thermal control is obtained by deliberately dumping fluid.
Since the system is designed to operate at low temperatures, a low-density fluid, such as ammonia, may on occasion, depending on heat loading, undergo a phase change. Boiling heat transfer in a low gravity environment is a complex phenomenon, which is not well understood at the present time. Because the system is subjected to meteoroid impact, the basic primary pump loops must be strongly protected.
Despite these drawbacks, pump loop systems will probably be used in conjunction with heat pipe systems as thermal control engineers create a viable space environment. These armored (closed) systems are rather highly developed and amenable to engineering analysis. They have already found an application on Earth and in space. A strong technology base has been built up, and there exists a rich literature for the scientist-engineer to draw on in deriving new concepts.
Advanced Radiator Concepts
The very nature of the problems just discussed has led to increased efforts on the part of the thermal management community to examine innovative approaches which offer the potential of increased performance and, in many cases, relative invulnerability to meteoroid strikes. Although I cannot discuss all of these new approaches, I will briefly describe some of the approaches under study as examples of the direction of current thinking.
Improved Conventional Approaches
The continuing search for ways to improve the performance of heat pipes has already shown that significant improvements in the heat pumping capacity of the heat pipe can be made by clever modifications to the return wick loop (a capillary structure, also called wick, can be placed inside the heat pipe body to allow the condensed liquid phase of the working fluid wicking against the vapor flow due the capillary action; such a heat pipe is called a wick heat pipe).
Looking further down at the problem of deployability, people are exploring flexible heat pipes and using innovative thinking. For example, a recent design has the heat pipes collapsing into a sheet as they are rolled up, the same way a toothpaste tube does. Thus, the whole ensemble may be rolled up into a relatively tight bundle for storing and deploying. However, because the thin-walled pipes are relatively fragile and easily punctured by meteoroids, more redundancy must be provided. The same principles, of course, can be applied to a pump loop system and may be of particular importance when storage limits must be considered. These are only examples of the various approaches taken, and we may confidently expect a steady improvement in the capability of conventional thermal management systems.
The Liquid Droplet Radiator
The basic concept of the liquid droplet radiator is to replace a solid surface radiator with a controlled stream of droplets. The droplets are sprayed across a region in which they radiate their heat; then they are recycled to the hotter part of the system.
It was demonstrated some time ago that liquid droplets with very small diameters (about 100 micrometers) are easily manufactured and offer a power-to-mass advantage over solid surface radiators of between 10 and 100. In effect, large, very thin radiator sheets can be produced by the proper dispersion of the droplets. This system offers the potential of being developed into an ultra-lightweight radiator that, since the liquid can be stored in bulk, is also very compact.
The potential advantages of the liquid droplet radiator can be seen if we consider again the problem that was discussed at the end of the section on heat pipe radiators. We found that a very good aluminum radiator would require 256 m2 and have a mass of nearly 1300 kg to radiate the low-temperature waste heat from lunar processing. Using the properties of a liquid droplet radiator and a low density, low vapor pressure fluid such as Dow Corning 705 (a diffusion pump fluid designed for high vacuum and for fast pumping of large volumes of gas or vapor in production operations), a common vacuum oil (i.e. a liquid with a low vapor pressure at room temperature; it is one of the vacuum materials; it is used chiefly as a pressure fluid in oil-diffusion pumps, as a sealant for mechanical pumps, as a lubricating material for the rubbing parts of vacuum devices, and for filling liquid vacuum meters) we find that, for the same area (which implies the same emissivity), the mass of the radiating fluid is only 24 kg. Even allowing a factor of 4 for the ancillary equipment (i.e. any devices including, but not limited to, such devices as piping, fittings, flanges, valves, and pumps used to distribute, meter, or control the flow of regulated substances; e.g. circulating pumps, cooling fans and defrost heaters) required to operate this system, the mass of the radiator is still less than 100 kg.
To achieve efficiency, the designer is required to frame the radiator in a lightweight deployable structure and to provide a means of aiming the droplets precisely so that they can be captured and returned to the system. However, present indications are that the droplet accuracies required (milliradians; a unit of angle also known as a mil; a unit of measurement dividing radians in a circle) are easily met by available technology. Recently, successful droplet capture in simulated 0 g conditions has been adequately demonstrated. An advantage of a liquid droplet radiator is that even a relatively large sheet of such droplets is essentially invulnerable to micrometeoroids since a striking micrometeoroid can remove at most only a few drops.
Some may be concerned that the very large surface area of the liquid will lead to immediate evaporation. However, liquids have recently been found that in the range of 300 to 900 K have a vapor pressure so low that the evaporation loss during the normal lifetime of a space system (possibly as long as 30 years) will be only a small fraction of the total mass of the radiator. Thus, the liquid droplet radiator appears promising, particularly as a low-temperature system where a large radiator is required.
Liquid droplet radiators for applications other than 0 g have been suggested. For example, in the lunar environment fluids with low vapor pressures can be used effectively as large-area heat dissipation systems for relatively large-scale power plants. We may well imagine that such a system will take on the appearance of a decorative fountain, in which the fluid is sprayed upward and outward to cover as large an area as possible. It would be collected by a simple pool beneath and returned to the system. Such a system would be of particular advantage in the lunar environment if low mass, low vapor pressure fluids could be obtained from indigenous materials. Droplet control and aiming would no longer be as critical as in the space environment; however, the system would need to be shaded from the Sun when it is in operation.
While this system is far less developed, its promise is so high that it warrants serious consideration for future use, particularly in response to our growing needs for improved power management.
Belt Radiator Concepts
The belt radiator concept is a modification of the liquid droplet concept in which an ultrathin solid surface is coated with a very low vapor pressure liquid. While the surface-to-volume ratio is not limited in the same fashion as for cylindrical heat pipe, it does not quite match that of the liquid drop radiator. However, this system avoids the problem of droplet capture by carrying the liquid along a continuous belt by surface tension. The liquid plays a double role in this system by acting not only as the radiator but also as the thermal contact which picks up the heat directly from a heat transfer drum. Variations on this scheme, in which the belt is replaced by a thin rotating disk, are also feasible but have yet to be fully assessed.
Thermal Management System of the MARE (Moon Age and Regolith Explorer)
The MARE (Moon Age and Regolith Explorer) Discovery Mission concept targets the delivery of a science payload to the lunar surface for sample collection and dating with the objective of determining the impact history of the inner solar system and the evolution and differentiation of the interiors of one-plate planets. The mission science is within a 100-meter radius region of smooth lunar maria terrain near the Aristarchus crater, 23.7°N 47.4°W. The lunar lander proposed for this mission was based on the NASA JSC Morpheus lander vehicle.
As part of this work, the NASA JSC and MSFC thermal teams proposed a thermal management system. This thermal management concept was intended to not only meet the driving requirements below, but also leverage existing SBIR work with the objective of developing technologies that have the potential of reducing mass, power, cost, and complexity while maintaining thermal performance for future architectures.
- The science payload should be able to survive approximately 2 lunar days and one lunar night.
- The thermal control system should be able to dissipate 323 watts of waste heat during science payload operations.
- Thermal control should be able to keep all components within operational and survival limits for science payloads and lander components for extended periods of time.
- The thermal control system should be able to operate during different mission phases.
- The thermal control system should be able to operate at a 25° slope.
The design approach took into consideration the following thermal control strategy in order to propose a simplified thermal management solution for this mission:
- Passive thermal control will be pursued to the extent possible to reduce complexity, mass, and cost.
- Package components to be thermally conditioned in one location if possible.
- Assume payload structure is thermally optimized to be controlled by a common heat spreader.
- Passive heat spreader plate (i.e. a plate or block of material having high thermal conductivity, such as copper, aluminum, or diamond),
- Controlled to a set temperature range (0°C to 50°C) via heat pipes and heaters.
- Directly mount multiple subsystem components to the top side of the heat spreader plate (a two-sided heat spreader could be an option).
- Structure must be thermally isolated from the heat spreader to control heat leaks (the use of heat switches can be evaluated).
- Use of Variable Conductance Heat Pipes (VCHPs) to thermally connect to the radiator.
- Multi-Layered Insulation (MLI) will be used to cover the exterior surface of the thermal control system to protect it from the environment.
- Remote units will be evaluated separately and proper thermal controls will be placed.
- Heater will be used to survive the lunar night.
The MARE thermal control architecture proposed consists of a simple passive thermal control system, making this a reliable, cost and mass-effective approach. This system is composed of two thermally controlled zones on the top deck of the spacecraft to maintain science and vehicle components within operational and survival limits. Cooling is achieved with two high conductivity heat spreaders connected via hybrid Variable Conductance Heat Pipes (VCHPs) to two high conductivity radiators. In the proposal phase, two high conductivity aluminum encapsulated Annealed Pyrolytic Graphite (APG; also known as Thermally Annealed Pyrolytic Graphite (TPG), is a form of synthetic graphite that offers excellent in-plane thermal conductivity) heat spreaders were assumed as the heat spreader plate and the radiators. Driven by cost reduction, and the opportunity to leverage on existing SBIR contracts, these were replaced for testing for 2 HiK™ plate heat spreaders, one HiK™ plate radiator, and one Gas Charged Heat Pipe (GCHP) radiator.
The thermal links provide a means to control the heat leak from the heat spreaders, depending on the set-point. They conduct heat to the radiators when the heat spreader surface temperature is above the upper-temperature limit and turn off by heating the non-condensable gas reservoir, disconnecting the ammonia flow within the VCHP when the heat spreaders are below the lower temperature limit.
In order to survive the lunar night, thermostatically controlled Kapton heaters, directly fed from the primary battery input, provide the necessary heat to keep electrical components within operational and survival thermal limits. The survival heater power needed is approximately 55 watts. Multi-Layered Insulation (MLI) blankets will shroud all top deck components to reduce the amount of heat leak paths to the cold lunar environment and space environment. A thermal math model was developed for the proposed MARE Thermal Control System (TCS) architecture, and thermal analysis was performed for the worst cold and hot mission environments.
The results show that the TCS subsystem will protect the MARE payload from the worst-case hot mission environment with at least a 13°C margin for all components. Results also showed at least a 12°C margin for all components for a worst-case cold mission environment.
Shape-Shifting Technology to Adjust Thermal Control Systems Automatically
With lunar explorations on the horizon, including putting astronauts back on the Moon by 2025, NASA is investing $2 million in cutting-edge thermal technology to make regulating temperatures during missions possible. This technology will be developed by a team of researchers from Texas A&M University, the Boeing Company, and the Paragon Space Development Corporation. The team is focused on creating shape-shifting technology to adjust thermal control systems automatically. The proposed solutions incorporate shape-shifting metals that adjust their own heat rejection based on how hot or cold they are, solving the problem. Prototypes of the morphing radiator were developed and successfully tested the prototypes in a small thermal vacuum chamber at NASA’s Johnson Space Center.
Carbon Fiber Thermal Management Solutions by KULR Technology
KULR Technology, a subsidiary of KT High-Tech Marketing Inc. announced that its carbon fiber thermal management solutions, in particular, custom-designed phase change heat sinks, will be used on NASA JPL missions – the CubeSat/SmallSat Lunar Flashlight mission and the Mars mission as part of the Mars Rover SHERLOC (Scanning Habitable Environment with Raman & Luminescence for Organics & Chemicals) equipment. For both missions, the KULR Technology heat sinks will keep critical and sensitive components such as lasers and corresponding sensors at a cool and consistent temperature throughout their use, avoiding signal distortion or other complications that can arise from overheating.
The KULR design included in the “Lunar Flashlight” and “SHERLOC” projects is a unique and highly effective phase-change system that incorporates KULR’s proprietary, highly conductive vertical carbon fiber architecture with a material similar to wax that can change from solid to liquid while absorbing high amounts of heat energy. The combination of materials designed and assembled by KULR to exact specifications will draw heat safely away from sensors and other components needed to efficiently study lunar ice formations or scan for signs of life on Mars.
For the Lunar mission, if the Flashlight laser gets above 24 Celsius the data can degrade – jeopardizing the entire point of the mission. So, keeping it below 24 Celsius while the laser is spewing out heat at more than 100 Celsius is the trick. Among the more promising uses for KULR’s carbon fiber is dramatically improving battery safety. KULR, in development and testing with NASA, has developed a thermal shield that can prevent dangerous lithium-ion battery fires and explosions due to thermal runaway. KULR also announced an agreement with the National Renewable Energy Laboratory, funded by the U.S. Department of Energy, to be the exclusive manufacturing partner of the Internal Short-Circuit (ISC) device that can cause predictable lithium-ion cell failures in controlled conditions.
Flight Thermal Control System for the Volatiles Investigating Polar Exploration Rover, or VIPER
NASA awarded Advanced Cooling Technologies, Inc. (ACT) a $3.7 million contract to finalize the design and fabricate the flight thermal control system for the Volatiles Investigating Polar Exploration Rover, or VIPER.
VIPER, an important part of the Artemis program, is a golf cart-sized rover that will roam several miles across the South Pole of the Moon to get a close-up view of the location and concentration of ice during its approximately 100-day mission. The Artemis program will accomplish many “firsts” for NASA and the global space community. VIPER is NASA’s first lunar robotic rover and is the first resource-mapping mission on the surface of another celestial body. Additionally, the thermal control system developed jointly by ACT and NASA will be the first to maintain operation for a 100-day mission while in shadowed conditions. “It’s exciting to demonstrate some of the best passive thermal management solutions on such an important mission,” said Bill Anderson, Chief Engineer at ACT.
For this contract, Advanced Cooling Technologies (ACT) will deliver a combination of high-performance, passive thermal technologies specifically designed to meet the challenges associated with a long-term lunar mission.
- Loop Heat Pipes (LHPs) with integral Thermal Control Valves (TCVs)
- External Ammonia Constant Conductance Heat Pipes (CCHPs)
- Honeycomb Radiator Panels with embedded Constant Conductance Heat Pipes (CCHPs)
In addition to the flight articles (i.e. components; also assemblies and subsystems), ACT will be providing system and subsystem level thermal analysis (i.e. a technique used to analyze the time and temperature at which physical changes occur when a substance is heated or cooled), verification, and thermal testing in coordination with the teams at NASA’s Johnson Space Center in Houston and NASA’s Marshall Space Flight Center in Huntsville, Alabama. All flight articles were delivered by June 2022, meeting NASA’s critical launch schedule.
One of the most challenging aspects of VIPER’s mission is operating in shadowed conditions. Due to the slow rotation of the lunar surface relative to the sun, the environmental temperature drops to below 100K (-280F) for approximately 14 earth days. This condition has been avoided in the past, including the Apollo mission which landed in the lunar morning. To enable long-duration science missions on the lunar surface it is necessary to maintain the electronics above survival temperature while in shadows. The hardware ACT is developing reduces the amount of survival heat needed by an order of magnitude – from 100s of Watts to 10s. This significantly reduces the amount of solar power generation and mass required on the rover; ultimately extending life and functionality.
TRIPS (Thermal, Radiation Radiation and Impact Protective Shields) Technology Development Approach
The proposed concept is to undertake a steady, evolutionary technology development approach, with the long-term goal of developing nano-based materials for use in aerocapture (uses a planet’s or moon’s atmosphere to accomplish a quick, near-propellantless orbit insertion maneuver to place a spacecraft in its science orbit) and entry vehicles employed in robotic and human exploration missions. The materials would constitute a single-shield system, capable of simultaneous protection against aerodynamic atmospheric heating (occurs when aerobraking is used to modify the orbit of a spacecraft arriving at a planet with an atmosphere), solar and cosmic radiation, and micrometeorite/orbital debris strikes.
Concept: Nanotechnology-Based Shield for Solar Particles and Cosmic Rays
Lightweight materials such as hydrogen (the lightest element in the periodic table; the most abundant element in the universe), lithium (a soft, silvery-white alkali metal), and boron (in its crystalline form it is a brittle, dark, lustrous metalloid; in its amorphous form it is a brown powder) make better radiation shields than those made of high atomic weight systems since less secondary radiation is produced during the collision process with high-speed cosmic rays and solar particles. This is in contrast to X-rays (a penetrating form of high-energy electromagnetic radiation) and gamma rays (a penetrating form of electromagnetic radiation arising from the radioactive decay of atomic nuclei), which are better shielded by heavy materials.
While not as effective as hydrogen, carbon is also an effective radiation shield. Carbon chain polymers (the backbones are composed of linkages between carbon atoms; in heterochain polymers, a number of other elements are linked together in the backbones, including oxygen, nitrogen, sulfur, and silicon) such as polyethylene (the most common plastic in use today) or polystyrene (a synthetic aromatic hydrocarbon polymer made from the monomer known as styrene) contain a significant fraction of hydrogen and are often used in radiation shielding. For baseline NASA radiation shielding comparisons, polyethylene is the standard material.
Polymer-carbon nanotube (CNT) composites have the potential to improve radiation shielding performance if the nanotubes can be functionalized, or filled with significant amounts of hydrogen, lithium, or boron. Given the high strength of carbon nanotubes, these properties may enable a multifunctional material with high strength and high radiation shielding capability to be fabricated.
Approaches for Attaching Lightweight Atomic Species in Carbon Nanotubes and Making Fibers
Many approaches have been developed recently to functionalize and or fill carbon nanotubes with a variety of materials. Perhaps the largest interest comes from the fuel cell industry, where there is potential for a huge market for reversible hydrogen storage. These techniques are based either on the use of high pressure, electrochemical methods, or filling by capillary action (it causes the water in the thinnest tube to rise to a higher level than in the other tubes) as the nanotube is formed. The results have been somewhat disappointing, especially in terms of hydrogen storage, where early claims of large storage capability were later refuted.
Another method of functionalization (the process of adding new functions, features, capabilities, or properties to a material by changing the surface chemistry of the material) is ion implantation (allows precise control over the depth of penetration of dopant atoms into the silicon). This is a technique common in the electronics industry but has not received as much attention as the other methods for filling nanotubes, since the technique is not reversible. For NASA applications in radiation shielding, reversibility is not an issue, since the desire is to have the hydrogen a permanent part of the material. Furthermore, the method is compatible with either pre- or post-processing of carbon nanotube-polymer composites. This may be advantageous because many groups (such as the U.S. Army Research Labs) are working on the development of high-strength carbon nanotube-polymer composites for other applications such as bulletproof vests. Post-processing the best composites developed within or outside of NASA using hydrogen ion implantation may be an efficient use of resources.
In the ion implantation technique, ions are implanted directly into the composite with the energy selected for penetration through the film thickness. For a given energy, the distribution of the ions in the material is roughly Gaussian. Varying the energy can provide a more uniform distribution. Minimal damage to the composite during implantation will result if the film or fiber is relatively thin, enabling the use of low ion energy beams which will not break the carbon-carbon bonds and still penetrate through the proper depth. Once implanted, the hydrogen may functionalize or form a covalent bond (consists of the mutual sharing of one or more pairs of electrons between two atoms; these electrons are simultaneously attracted by the two atomic nuclei; a covalent bond forms when the difference between the electronegativities of two atoms is too small for an electron transfer to occur to form ions) with the interior or exterior of the nanotube, as well as form molecular hydrogen (H2; an abundant energy carrier, mostly produced from natural gas by steam reforming, and it is also the fuel of choice for a sustainable future, particularly if produced from water and biomass), inside the tubes or in the interstices (i.e. small or narrow spaces or intervals between things or parts).
Scientists have identified two methods of ion implantation (i.e. a low-temperature process by which ions of one element are accelerated into a solid target, thereby changing the physical, chemical, or electrical properties of the target; used in semiconductor device fabrication and in metal finishing, as well as in materials science research) that may be useful in this application. The first uses a commercially available ion gun (an instrument that generates a beam of heavy ions with a well-defined energy distribution), which simply accelerates and implants directly into the sample. The second method is plasma immersion (PIII; or pulsed-plasma doping, pulsed PIII, is a surface modification technique for extracting the accelerated ions from the plasma) or plasma source ion implantation (PSII; an ion‐implantation technique that has been optimized for surface modification of materials such as metals, plastics, and ceramics). In this technique, the sample is immersed in a low-temperature plasma chamber filled with hydrogen or other species. When the sample is biased to a negative voltage, electrons are driven away and ions are accelerated towards the sample and become implanted. It can be shown that if the energy of the ions is kept below 70 eV (an electronvolt is the measure of an amount of kinetic energy gained by a single electron accelerating from rest through an electric potential difference of one volt in vacuum), the ions will penetrate and dope thin samples but not dislocate carbon atoms from the nanotube lattice, thus retaining the high strength characteristics of the fibers. The method has many advantages over beam-line ion implantation, including a high dose rate and uniform coverage.
Thermal Properties of Lunar Dust
Solar absorptance (𝛼, alpha) and thermal emittance (𝜀n and 𝜀) of spacecraft materials are critical parameters in determining spacecraft temperature control. Because thickness, surface preparation, coatings formulation, manufacturing techniques, etc. affect these parameters, it is usually necessary to measure the absorptance and emittance of materials before they are used. Also, because most materials exhibit some amount of dye degradation to outgassing (i.e. the release of trapped gas or vapor that was previously dissolved, trapped, frozen, or absorbed in the solid), ultraviolet, and or particle damage, it is necessary to conduct laboratory testing on these materials before certifying them for use in space.
In order to survive the hot and cold extremes of the lunar environment, a vehicle would ideally use an outer insulation layer with a low solar absorptivity (a low solar absorption variation means that the solar transmittance is regulated partly via reflection), to reduce the heat absorbed during periods of illumination (to reflect the solar radiation), and a low IR emissivity (i.e. low infrared emissivity; a surface condition that emits low levels of radiant thermal (heat) energy), to reduce the heat lost during periods of darkness. However, previous lunar experience has shown that lunar dust is extremely effective in coating external surfaces. As well as being potentially kicked up by the wheels of the rover, charged solar particles can levitate dust up to 10 m from the surface as they pass through the solar terminus. An accumulation of dust on radiators over time is possible and it is necessary to cover the radiators and solar arrays before passing from lunar day into lunar night. There also exists the possibility that in passing in and out of shadows (such as boulder shadows) lunar dust may levitate the performance of radiators and solar arrays will decrease as dust thickness on them increases. The location of radiator surfaces along with mitigating actions to reduce dust must be considered in the system design. Locating radiators on the top surface will avoid dust contamination during rover operation.
Being so small, under 20 microns in diameter, the dust can easily wear away at any
exposed equipment. The dust can also become electrostatically charged due to radiation,
increasing its ability to cling to objects. Dust coating on the outer MLI will increase the absorptivity and emissivity. For the most robust design, a black insulation outer layer may be selected, as the optical properties will be relatively unaffected by the addition of a layer of lunar dust.
By using a pigment-treated paint, the electrical charge of lunar dust in contact with the spacecraft’s components can be reduced. The paint relies on atomic layer deposition (ALD) to dissipate electricity, thereby preventing dust from clinging to the spacecraft. Additionally, all motors and electronic components should be sealed to protect them from dust.
Solar Illumination – Radiator Location and Sizing
The thermal performance of the rover during periods of solar illumination is determined by the size and placement of radiator surfaces, along with the rover body insulation. Because of the impact of lunar dust on the optical properties of the insulation outer surface, a black outer layer has been taken as the baseline to provide the most robust system design.
The most effective radiator location provides the radiator with a view to cold space whilst avoiding viewing the lunar surface and avoiding direct solar illumination. Radiators located on the rover side walls would have a minimum 50% view of the lunar surface, with a greater view in areas with boulders or sloping terrain. In addition, because of the low angle of the sunlight with the ground plane, the rover side walls will experience high solar flux (i.e. concentrated sunlight; a measure of how much light energy is being radiated in a given area). For these reasons, the rover’s top surface is a much better candidate for radiator placement.
The upper surface of a lunar rover will be exposed to direct solar flux when operating on sloped terrain. Radiators exposed to direct flux absorb a large amount of heat, leading to a feedback runaway in radiator sizing. This effect can be seen in the preliminary radiator sizing. As the ground angle increases, the incident solar flux on a radiator located on the rover’s top surface increases, leading to a runaway in the required radiator area.
To avoid this runaway, it is necessary to shield the radiator surfaces from the solar flux. For radiators located on the rover side walls, it would be necessary to position side radiator surfaces out of direct sunlight, which would seriously inhibit the rover’s movement and operation. Therefore it is preferable to limit the required radiator area to the top rover surface only, and protect this surface from direct solar illumination, either using the solar array or some additional structure. Limiting the radiator area to the rover’s top surface will also limit the impact of lunar dust on radiator performance, and allows for the option of covering the radiator surface during periods of darkness to reduce heat loss to the environment.
Alternatively, the radiator surface could be located on the back surface of the solar array. This surface, by definition, will never be illuminated; however, there are additional complexities with this design option. The radiator surface must be thermally decoupled from solar array heating and must be well coupled to the rover body to reject heat. A radiator located on the back surface of the array would also have a view of the lunar surface, which would limit its overall efficiency.
The least complex, lowest mass option is to limit the radiator area to the rover’s top surface. In order to achieve this, it is required to limit the overall rover dissipation in addition to protecting the top surface from direct solar illumination. The baseline configuration uses deployed solar array panels to shadow the top surface. The panels have been sized to ensure that the top surface is completely shadowed when operating on a ground slope of 25° and the back surface has a mirror finish to reflect any incident solar flux away from the radiator surface.
Warm Electronics Box Panels
The equipment at the heart of a rover is protected by a low conductivity, lightweight enclosure typically referred to as a warm electronics box (WEB). The outer layer of the vehicle provides structural integrity and protection against environmental loads and a warm electronics box shields vital components from the extremes of space.
The WEB is envisioned to combine structural and thermal insulation properties in order to withstand the mechanical loads while ensuring that the temperatures of the equipment inside are maintained within the operating limits. The resulting solution produces three different designs for the top, bottom, and lateral panels of the box.
The bottom panel facing the ground is designed to reduce the heat flux radiated outside so as to prevent the regolith from heating up. It is therefore constituted by a honeycomb structure made of an aluminum core and insulating carbon nanofibers (nanofibers fibers with diameters in the nanometer range, typically, between 1 nm and 1 μm; carbon nanofibres, CNFs are cylindrical nanostructures with graphene layers arranged as stacked cones, cups or plates) composite face sheets. On the external side, the surface is coated with the MLI. On the internal side, the structure is cladded (refers to components that are attached to the primary structure to form non-structural, external surfaces) with a layer of silica aerogel (a solid material of extremely low density, produced by removing the liquid component from a conventional gel), which is in turn covered with a low conductivity E-glass fibers/epoxy (also known as electrical glass is the standard glass composition used for most glass fibers; extremely strong materials used in roofing, automobiles, pipes, textile industry, composite materials, and are present in 90% of reinforcements; was originally developed for stand off insulators for electrical wiring) composite sheet. Silica aerogel has a proven history of exceptional thermal insulation performance in space.
Due to the vacuum conditions, the thermal conductivity of 12 × 10−3 𝑊/𝑚 (Weight/mass) is used in sizing the insulation layer to accommodate the performance of the WEB. The aerogel is opacified (i.e. made opaque, not transparent) to reduce the transmission of IR radiation across the wall. The glass-epoxy composite layer protects the underlying aerogel and offers a substrate to support the electronics.
Finally, this surface is covered by a goldized Kapton film that ensures additional retainment of the heat radiated by the equipment inside the box. A similar configuration characterizes the lateral panels, with few changes. Here, the innermost goldized Kapton coating is replaced by a black-painted Kapton layer. The high emissivity value of this film (𝜀 = 0.92) allows for achieving more uniform temperatures. MLI is replaced by a surface coating with an emissivity 𝜀 = 0.4 and an absorptivity 𝛼 = 0.3. These optical properties can be matched either by vacuum-deposited silver and Inconel-backed Teflon or by aluminized Kapton.
However, Teflon (Polytetrafluoroethylene, PTFE; a synthetic fluoropolymer of tetrafluoroethylene that has numerous applications) is recommended due to its thermal resistance ranging from -184 to 150°C. Finally, the top panel is designed to achieve higher radiative heat exchanges with the external environment in order to avoid internal equipment overheating. This translates into a replacement of the carbon composite facesheets covering the inner aluminum honeycomb core with more conductive aluminum face sheets. Moreover, both the silica aerogel and the glass-epoxy insulating layers are eliminated. Therefore, the honeycomb structure is internally coated with black paint and externally covered with backed Teflon.
The Moon is looked at as the first possible mission site for the exploration robots, as funding appears to be most easily acquirable due to a need for reconnaissance and mapping in view of future planned human deployment.
The discussed mission concept would allow several advantages compared to a traditional rover mission, mainly due to less restricted mobility restraints, especially in caves and heavily cratered terrain. Such a mission would set the stage for later life detection missions, for example on Mars or Jupiter’s Galilean Moons. The decision on a thermal control mechanism installation must be carefully evaluated.
As the majority of the proposed concepts and components must still be developed to meet specifications, in essence, a curved Variable Emittance Coating surface and a small lightweight heat switch, a thorough analysis must be conducted. Because the use of a thermal control provision has a direct influence on system cost, weight, complexity, energy requirement, and mission flexibility, all these effects must be included in the study to see if such a system is worthwhile installing and pursuing.