Non-Integrated Warm-Reservoir Variable Conductance Heat Pipes (VCHPs) with Hybrid Wicks for Lunar Landers and Rovers
Jeff Diebold, R&D Engineer
An increasing number of applications such as satellites and planetary rovers have unprecedented thermal requirements due to the need to dissipate high power while operating in extreme thermal environments. Thermal management systems need to reject large heat loads into hot environments while also having high heat rejection turn-down ratios in order to minimize power needs during periods of extreme colds, such as the 14-day-long lunar night. Variable Conductance Heat Pipes (VCHPs) are capable of passively transporting large quantities of heat and provide high thermal turn-down ratios ideal for surviving extreme cold environments. For lunar and planetary surface applications, warm-reservoir VCHPs offer superior passive thermal control.
ACT recently developed a non-integrated warm-reservoir VCHP with a hybrid wick for lunar surface applications.
The non-integrated reservoir and the hybrid wick each solve a problem common to this application
- The non-integrated reservoir improves control over the distribution of working fluid and non-condensable gas (NCG) during periods of non-operation.
- The warm reservoir provides excellent thermal control.
- The hybrid wick allows the VCHP to operate effectively in both microgravity (transit) and on the surface in a gravity-aided orientation.
Warm-reservoir Variable Conductance Heat Pipes (VCHPs) offer superior passive thermal control compared to cold-reservoir VCHPs, but over time their performance can be degraded due to the migration of working fluid into the reservoir. This process displaces non-condensable gas (NCG) from the reservoir resulting in a higher nominal operation temperature for the heat pipe. If the reservoir is not integrated with the evaporator (separated from the evaporator), then during periods of non-operation an independent heater can be applied to the reservoir in order to purge the working fluid and restore normal operation.
HEAT PIPES MAXIMIZE POWER CAPABILITY
When a heat pipe must operate in microgravity (during transit) a grooved wick structure is generally required due to the high permeability of the grooves. On the lunar or planetary surface, it is generally desirable for the heat pipe to operate in a gravity-aided orientation to maximize its power carrying capability. In a gravity-aided orientation, grooved wicks have a tendency to exhibit large temperature spikes during startup due to the working fluid pooling at the bottom of the evaporator. The hybrid heat pipe, shown in Figure 1, utilizes a hybrid wick that contains screen mesh, metal foam or sintered evaporator wicks for the evaporator region, which can operate on planetary surfaces and sustain high heat fluxes and axial grooves in the adiabatic and condenser sections that can transfer large amounts of power over long distances due to their high wick permeability and associated low liquid pressure drop.
DESIGN AND TESTING FLIGHT HARDWARE
ACT designed and fabricated the non-integrated warm-reservoir VCHP with hybrid wick shown in Figure 2, as flight hardware a future Lunar Landers. The envelope material was aluminum and the working fluid was ammonia. The envelope and porous wick in the evaporator was 3D printed and were designed to interface with the grooved extrusion of the adiabatic and condenser sections. The hybrid wick will allow the VCHP to operate in microgravity and on the lunar surface in a gravity-aided orientation. The non-integrated reservoir of NCG contained an independent heater for purging the reservoir of working fluid.
“The hybrid wick will allow the VCHP to operate in microgravity and on the lunar surface in a gravity aided orientation.”
The VCHP for Astrobotic was designed to transfer 40W at a temperature of 55°C into a sink temperature of 40°C. The maximum operating power of 40W was well below the theoretical maximum for this pipe based on flooding and capillary limits. Figure 3 shows temperature measurements on the VCHP as a function of time. Temperatures corresponding to the evaporator, adiabatic, condenser, and reservoir are indicated. At approximately 4,000 seconds a steady-state operating temperature of 58.9°C was reached at the evaporator. This high temperature was due to ammonia migrating into the reservoir and displacing NCG increasing the pipe’s thermal resistance. Two purge tests were then carried out. During the purging process, the heat source at the evaporator was shut down, the sink temperature was decreased to 0°C and a small amount of power was applied to the reservoir heater. After the first purge, the steady-state operating temperature decreased to 56.5°C and after a subsequent purge operation, the design operating temperature of 55°C was achieved.
THERMAL CONTROL TEST RESULTS
This test successfully demonstrated the ability of the non-integrated warm-reservoir VCHP to purge the reservoir of working fluid in order to maintain the nominal operating conditions. Working fluid migrates into the reservoir during long periods during which the pipe is not operating, for example prior to launch. A purging process can be applied shortly before the pipe is required to operate in order to ensure the proper operation.
Figure 4 shows the results of a thermal control test of the non-integrated warm-reservoir VCHP. The system began at the nominal operating point of 40W, an evaporator temperature of approximately 55°C, and a chiller block set point of 40°C. While maintaining the power constant at 40W, the chiller block set point (the heat sink) was decreased to -100°C (a decrease of 140°C). During this time the evaporator temperature decreased only 16°C due to the increased thermal resistance provided by the expanding NCG. This highlights the excellent thermal control capability of the warm-reservoir VCHP during normal operation. The power supplied to the evaporator was then decreased to 1W and the evaporator achieved a new steady state of approximately -40°C, a reasonable survival temperature for many electronics. This represents a turndown ratio of approximately 185:1.
3D Printed Evaporators for Loop Heat Pipes
The traditional process for Loop Heat Pipe (LHP) evaporator fabrication is labor-intensive due to complex manual steps such as wick fabrication, vapor groove cutting, and wick-envelope integration. These steps contribute to high costs and extended lead times, particularly in the case of small evaporators for mini loop heat pipes, due to the added challenges posed by small-scale manual fabrication. An innovative solution was therefore needed to significantly reduce the manufacturing costs of evaporators in order to make the system financially viable for use on SmallSats and in particular, CubeSats.
ACT has been conducting an R&D campaign aimed at creating 3D printed evaporators for Loop Heat Pipes (LHPs) funded through the NASA SBIR program. With 3D printing, the entire evaporator is fabricated in a single operation, complete with such features as the wick, solid envelope, and vapor grooves, thereby eliminating the need for complex manual steps in evaporator fabrication. This innovation amounts to over an order-of-magnitude manufacturing cost reduction and significant time savings as compared to the traditional fabrication process.
As one of the primary components of this R&D effort, the 3D printing build parameters are optimized to improve wick pore size and the associated capillary limit. A high capillary limit is, in turn, crucial to achieving the desired power limit of the LHP. The optimization process follows an iterative routine which has been successful so far in reducing the maximum pore radius from over 8.5 μm (Sample 1) to under 5 μm (Sample 4) in just four iterations, as shown in Figure 1.
With the current best build parameters, several evaporators of varying sizes were fabricated, shown in Figure 2, and tested across a range of input powers, condenser temperatures, and working fluids. A sample test data for a 4-inch-long evaporator is presented in Figure 3. In this test, the Loop Heat Pipe (LHP) was subjected to a maximum power of 350 W, which, as seen from Figure 3, is well within the power limit of the system. Additional testing is planned at higher input powers and even lower condenser temperatures. As a particularly noteworthy development, a full Loop Heat Pipe (LHP) architecture with a 3D printed evaporator was designed and fabricated to mission-specific demands for thermal management on NASA’s Volatiles Investigating Polar Exploration Rover (VIPER). The goal of this system was to effectively transfer heat from one or more of the scientific instruments to the radiators onboard the lunar rover. The system thus developed was tested successfully in reflux orientation in the laboratory to emulate the near-vertical operation in lunar gravity.
Current efforts are directed to the development of advanced wick structures that can go beyond the performance of traditional wicks and improvement of the overall thermal conductance of the system. Further details on the work completed so far can be found in the following reference:
Gupta, R., Chen, C-H., and Anderson, W.G., “Progress on 3D Printed Loop Heat Pipes.” Proceedings of the 50th International Conference on Environmental Systems, ICES-2021-154, 2021. (https://www.1-act.com/resources/tech-papers/progress-on-3d-printed-loop-heat-pipes/)
Evaporator fabrication using 3D printing leads to lower manufacturing costs, with current estimates indicating an order-of-magnitude cost reduction as compared to the traditional fabrication process. 3D printed evaporators are also associated with a significantly smaller lead time as compared to a traditional evaporator.
What is the impact on the spacecraft industry?
3D printed evaporators enable the fabrication of low cost and low lead time mini loop heat pipes that are well-positioned to meet the thermal management demands of the ever-growing SmallSat industry, particularly CubeSats, that are used for both research and commercial purposes.
NASA partners with ACT for challenging aerospace welding requirements
The goal of NASA’s Europa Clipper mission is to assess its habitability by orbiting Jupiter to investigate Europa’s geology, composition, and the water that is suspected to exist under an icy crust on Jupiter’s moon, Europa. Europa possesses the smoothest surface known of any solid object in the solar system. This visible smoothness and apparent youth of Europa’s surface led researchers to the hypothesis that a water ocean exists beneath the icy crust of the surface and could conceivably lead to extraterrestrial life upon the moon. This ocean is expected to cover the moon and would contain more water than all of Earth’s oceans combined.
The Europa mission will use the clipper to do flybys near the moon and collect data from the surface at certain points within the orbit, without actually landing any crafts upon the surface. It is critical that the clipper is equipped with the best tools possible to collect data during these flybys; NASA has announced 9 named specialty instruments.
NASA has collaborated with ACT engineers on dozens of successful projects over the last 15 years. So, when NASA wanted to work through a welding issue that arose during the design of the Europa Clipper, ACT’s aerospace team was happy to take on the challenge.
The flange tubing for a pumped single-phase fluid cooling loop needed to be welded together and both withstand the journey to, and operation in, deep space for an extended period of time. While aerospace welding standards are extreme by nature, this weld joint is critical to systems operation, and thus requires standards that are not commonly seen by NASA’s aerospace partners.
Previous bi-metallic welding of heat exchanger components for aerospace have been successfully produced by ACT personnel in the past, however, the latest work required much more detailed inspection and qualification of the weld quality– examination that is beyond the ability of the human eye, and therefore, must be completed at the microscopic level.
“This project required aluminum to bimetallic welding, which is a unique and uncommon welding process for most of the world’s welding needs.” Said Brent Bennyhoff, Aerospace Certified Welding Technician at ACT. He further explained that “it’s challenging because when welding a temperature-sensitive component that has materials of different CTE’s (coefficient of thermal expansion), extra measures are needed in the process to ensure no damage occurs to the bimetallic fitting’s functionality.”
ACT’s in-house radiography capabilities allow a project such as this to be quickly x-rayed and evaluated for requirements. “We’re proud to support NASA flagship missions by providing flight-critical hardware. In a mission such as the Europa Clipper, all aspects of a complex spacecraft need to operate as expected; we’re proud our hardware is being relied upon for a mission undertaking such a significant scientific milestone.” Said ACT’s Ryan Spangler, Lead Engineer of Defense and Aerospace Products about this challenging welding project.
Ryan’s team works closely with the NASA teams across many locations as well as with ACT’s Research & Development teams, who are collaborating on a NASA Phase II contract on another portion of the Europa mission. What aerospace challenges can our thermal engineers and technicians help you tackle?
Read more about NASA’s mission to Europa!
Super Cool Small Satellite Thermal Control System
As the nanosatellite and microsatellite industry continues to expand, there is a continuing need to deploy components that output higher power from smaller packages. As power increases, devices typically require more surface area to dissipate the additional waste heat. However, on size and weight sensitive small satellites, this is not a practical solution. Recently a worldwide leader in small satellite technologies faced this challenge of insufficient radiator volume for required component heat dissipation. They sought Advanced Cooling Technologies (ACT) to help them resolve their thermal management challenge.
The customer was seeking to use components that put out a tremendous amount of heat relative to systems designed and flown in in the past. Preliminary thermal calculations showed that more radiator surface area than was available was necessary for heat dissipation. Fortunately the components operated on a duty cycle, only operating at full power 10-20% of the time, while dormant the remainder flight orbit. The customer sought the assistance of ACT to utilize the 80-90% off time during orbit and change from an intermittent full power heat dissipation solution to a reduced power, time averaged dissipation solution– thus minimizing the needed radiator surface area.
To achieve this, ACT utilized a military and terrestrial flight-proven technology: phase change material (PCM). PCM is a material that, during solid to liquid phase transition, absorbs heat and utilizes the latent heat of the material to store energy during phase change. ACT designed a solution that, while the heat source was operating at full power, the PCM inside the heatsink was transformed from solid to liquid, without increasing the device temperature. The solution was also designed to have thermal resistance that would provide uniform dissipation of the stored heat through the radiator panel over the entire orbit cycle. This made the radiator panel size an order of magnitude smaller, and an acceptable size for their spacecraft.
An additional challenge was in the radiator panel itself. Traditional honeycomb-structure radiator panels can have a long lead time to manufacture. ACT designed a simple aluminum sheet metal radiator panel capable of surviving the anticipated launch shock & random vibration profile. However, thin aluminum is not a good mechanism for efficiency in radiation as it doesn’t spread heat in plane very well over long distances; however, increasing the thickness to a needed value would greatly increase mass of the system. ACT utilized a proven spaceflight technology – Aluminum/Ammonia Constant Conductance Heat Pipes (CCHPs) – to both 1) take the heat from the PCM to the radiator panel and 2) to spread the heat across the radiator, providing a high enough efficiency for dissipation of the received heat load.
The resulting thermal control system reduced the needed radiator panel surface area by an order of magnitude, reducing mass and providing a feasible solution to implement, reduced radiator panel complexity and cost, provided fully passive thermal control, and required minimal survival power to maintain minimum allowable temperature of the electronics.
Accumulators for Pumped Fluid Systems
The Japanese Experiment Module – Exposed Facility (JEM-EF) is a platform on the International Space Station (ISS) that is continuously exposed to the space environment. Up to 8 instruments can be installed on the facility at any time, typically focusing on Earth observation, as well as communication, scientific, and engineering experiments.
JEM-EF provides utilities to each payload location, including power, control, and active cooling using a single-phase pumped loop. All instruments plugged into the JEM-EF system require accumulators to accommodate fluid volume changes during launch, and while plugged into the cooling system. ACT developed, fabricated, and tested accumulators to accommodate the JEM-EF requirements. All stainless-steel accumulators provide volumetric compensation for fluid property changes across the broad temperature range experienced during launch and operation, from -40°C to 55°C. This product was designed, manufactured and qualified by ACT to operate for up to 3 years in this challenging environment. ACT introduced the accumulators in response to multiple requests for a rapid design and development supplier for custom, flight-qualified, fluid compensation systems.
Four of ACT’s four fluid compensation accumulators installed in NASA’s Cloud-Aerosol Transport Systems (CATS) instrument were successfully delivered to the International Space Station (ISS) aboard SpaceX’s CRS 5 Dragon capsule. After installation on the Japanese Experiment Module – Exposed Facility (JEM-EF), the CATS system then successfully operated for over two years. Additional accumulators in another instrument are scheduled for installation on JEM-EF.
Titanium Water Heat Pipes for Space Fission Power Cooling
Continuing earlier work on heat pipes for fission power (further reading also available: Low-Cost Radiator for Fission Surface Power II) Advanced Cooling Technologies, Inc. (ACT) developed a series of titanium-water heat pipe radiators to remove the waste heat from Kilopower system convertors. The titanium-water heat pipes have an advanced fluid management design, enabling the heat pipes and the Kilopower system to:
- Thermal Management: Transport waste heat in space and surface operations
- Survive and startup smoothly after being exposed to a frozen condition during launching or afterward.
The Kilopower system is an affordable, small-scale nuclear fission power plant that is designed to produce 1 to 10 kW of electricity to support NASA future space transportation and planetary exploration missions (figure 1). The Kilopower system uses Stirling conversion to generate power. The thermal energy generated from a nuclear fission reactor is transferred to the Stirling convertor hot-end via a series of high temperature (>800°C) alkali metal heat pipes. Parts of the thermal energy will be converted into usable electricity. The remaining waste heat needs to be removed from the Stirling convertor cold-end and ultimately rejected to the space environment through radiators.
ACT designed and fabricated multiple titanium-water heat pipes with radiator panels attached. Their thermal performance was validated through experimental measurement conducted both in ambient and a space-relevant environment (i.e. thermal vacuum chamber). The titanium-water heat pipes, operating at 400K, must be able to function and survive under the following four conditions:
- Operating in space without gravity forces.
- Operating on a planetary surface with a reduced gravity force for working fluid return.
- Testing on the ground to estimate space operation performance. To do so, the heat pipe evaporator should be slightly higher ( ~ 0.1 inch) than the condenser.
- Survival and recovery from a frozen condition. During a launch period, the heat pipes will be orientated in extreme against gravity and the sink temperature could be lower than the freezing point of the working fluid. It is necessary to incorporate a special wick design to manage the working fluid within the heat pipe, which can (a) avoid liquid staying inside the condenser and bursting the pipe while freezing and (b) supply enough amount of working fluid to startup the heat pipe after being frozen.
ACT developed a series of titanium water heat pipes for the Kilopower system cooling, based on our earlier titanium/water heat pipe work. As Figure 2 shows, the titanium water heat pipe has a C-shape evaporator to interface with the Stirling Convertor. Inside the evaporator, ACT inserted two types of screen mesh with different pore sizes which will enable the heat pipe to survive and recover from freezing. The rest of the pipe has an axial groove structure. Test results show that each titanium-water heat pipe is capable of transferring more than 400W of heat in the slightly adverse gravity inclination with a low thermal resistance at 0.01°C/W. The freeze/thaw test result (see figure 3) further demonstrates that the heat pipe can successfully recover from a frozen condition at -50°C to a normal space operation mode.
ACT also integrated aluminum flat sheets with the titanium water heat pipes through S-bonding, a cost-effective approach to join dissimilar metals. The titanium heat pipes with aluminum radiator panels are shown in Figure 4. Their thermal performance were tested in a space-simulated environment. As figure 5 shows, temperature distribution along the heat pipe with a radiator attached is very uniform, validating that the heat pipe radiators can effectively carry and reject the required waste heat in a space-relevant environment. ACT has delivered 7 titanium heat pipes to NASA Glenn Research Center for further performance validation.
ACT’s Ti-water heat pipes have the following key advantages:
- Highly reliable – no pumps, no fans, and no compressors involved
- Low mass – the weight of each Ti-water heat pipe with Al radiator is less than 0.75 kg.
- Highly conductive – the overall thermal resistance of the heat pipe radiator is 0.02°C/W.
- Operable in microgravity– it has been validated through ground testing that the heat pipe can transfer more than 350W of heat slightly against gravity.
- Freeze/thaw tolerance – it has been demonstrated that the heat pipe can smoothly recover from a frozen state to normal operation.
The advanced titanium-water heat pipe technology can be further applied in high heat flux electronics cooling and various spacecraft thermal management system requiring low-mass, high-performance solutions. Learn more about ACT’s work in the development of water heat pipe technology, at the following links:
High-Heat-Flux (>50 W/cm2) Hybrid Constant Conductance Heat Pipes
Constant Conductance Heat Pipes (CCHPs) have been a proven technology for spacecraft thermal control for more than 40 years. A CCHP transports heat over long distances (up to 3 m (10 feet) or more) from a heat source to a heat sink with a very small temperature difference. ACT currently has over 50 million hours in orbit for our CCHP product line (as of March 5th, 2018).
Future spacecraft and instruments developed for space science missions will involve highly integrated electronics, such as for CubeSat/SmallSat and high power laser diode arrays (LDAs). This high-density electronics packaging leads to substantial improvement in performance per unit mass, volume and power. However, it also requires sophisticated thermal control technology to dissipate the high heat flux generated by these electronics systems. For example, the current incident heat flux for laser diode applications is on the order of 5-10 W/cm2, although this is expected to increase towards 50 W/cm2. This is a severe limitation for the commonly employed axial groove aluminum/ammonia CCHPs. Hence, high flux heat acquisition and transport devices are required.
Typically, aluminum/ammonia CCHPs are used for transferring the thermal loads on-orbit due to their high wick permeability and associated low liquid pressure drop, resulting in the ability to transfer large amounts of power over long distances in micro-gravity. The maximum heat flux in a CCHP is set by the Boiling limit, where the working fluid within the evaporator wick structure starts to boil. If the heat flux is high enough, vapor bubbles will form and partially block the liquid return from the condenser to the evaporator, resulting in wick dryout. As the boiling limit is approached, the thermal resistance will continue to increase beyond the design parameters. In order to increase the heat flux limit to more than 50 W/cm2, Advanced Cooling Technologies, Inc. (ACT) developed a high heat flux heat pipes with a hybrid wick that contains screen mesh, metal foam, or sintered evaporator wicks, which can sustain high heat fluxes, for the evaporator region. The axial grooves in the adiabatic and condenser sections can transfer large amounts of power over long distances due to their high wick permeability and associated low liquid pressure drop as shown in Figure 1.
Heat spreaders are currently used in high heat flux applications to reduce the heat flux from the heat source to a level that can be accepted by the conventional axial grooves CCHP. This heat spreader adds weight, volume, thermal resistance, and cost to the system. ACT’s high heat flux CCHPs as shown in Figure 2 will eliminate the need of using these heat spreaders and will address a need for more demanding spacecraft thermal performance.
The performance of the new high-heat-flux heat pipe was validated after testing at ACT and Lockheed Martin (LM) Coherent Technologies, Inc. Figure 3 shows the comparison of the thermal resistance at ACT an LM as a function of power for the high heat flux heat pipe at ~ 5º adverse tilt. The high-heat-flux aluminum/ammonia CCHP transported a heat load of > 300 Watts with heat flux input of > 50 W/cm2 and thermal resistance < 0.012 ºC/W. Note the differences in the thermal resistance for the two test are caused by minor differences in the test setup. This demonstrates an improvement in heat flux capability of more than 3 times over the standard axial groove aluminum-ammonia CCHP design, also shown in Figure 3.
The testing results demonstrate that these new heat pipes perform efficiently, consistently and reliably and can adapt to many high heat flux applications. Accelerated life test show that the expected life for the hybrid high-heat-flux heat pipe exceeds 20 years.
 Ababneh, Mohammed T., Calin Tarau, William G. Anderson, Jeffery T. Farmer, and Angel R. Alvarez-Hernandez. “Hybrid Heat Pipes for Lunar and Martian Surface and High Heat Flux Space Applications.” 46th International Conference on Environmental Systems 10-14 July 2016, Vienna, Austria (2016).
Copper/Water Heat Pipes and HiK™ Plates for Spacecraft Applications
The waste heat from electronics must be removed to keep them from over-heating. On Earth, the ultimate heat sink is typically either a liquid coolant or the atmosphere. In space, the waste heat is typically transported by grooved-aluminum, Constant Conductance Heat Pipes (CCHPs) or Loop Heat Pipes (LHPs) to a radiator, and radiated into the environment. Constant Conductance Heat Pipes (CCHPs) represent an effective way to transport the heat over several meters, however, they have two limitations:
1. Maximum operating temperature
The maximum operating temperature is roughly 60°C for ammonia Constant Conductance Heat Pipes (CCHPs) and Loop Heat Pipes (LHPs), before the power carrying capability drops off. It is desirable to operate at as high a temperature as possible, since the thermal radiation scales with T4. Operating at a higher temperature allows a smaller and therefore lighter radiator.
2. Ground Testability
Grooved Constant Conductance Heat Pipes (CCHPs) have a very high permeability for flow, but a very low pumping capability. CCHPs are normally tested on Earth with an adverse elevation (evaporator above condenser) of 0.1in (2.5 mm). The power carrying capability drops to zero with an adverse elevation of about 0.4 in (1 cm). During ground testing of the spacecraft, the CCHP must be gravity aided or level. LHPs do not have this constraint but are considerably more complex and expensive.
Heat is generally transferred by conduction from the electronics to the aluminum/ammonia CCHPs. Two additional devices have been used to improve the heat transfer to the two-phase device: flexible thermal straps, and encapsulated conduction plates. Flexible thermal straps are usually used to transfer small amounts of power when the electronics move relative to the heat sink. Encapsulated conduction plates have an effective thermal conductivity of around 550 W/m-K, with two-dimensional spreading. In some encapsulated conduction plates, the effective thermal conductivity can decrease with thermal cycling. They are extremely expensive since they are formed by hot isostatic pressing. They also require thermal vias located below the electronics, so the electronics locations are fixed.
ACT has recently worked with NASA Johnson Space Center and NASA Marshall Space Flight Center to demonstrate flight heritage for two additional spacecraft thermal control devices: copper/water heat pipes and High Conductivity (HiK™) plates. Copper/water heat pipes are commonly used in many military and consumer electronics, including almost all laptops, typically in heat pipe assemblies. The benefits of copper/water heat pipes include their ability to operate at temperatures up to about 150°C, operate against adverse elevations of up to 25cm, and tight bend radius. Their major limitation is that the heat pipes carry only low powers at temperatures below ~20°C, and only transfer heat by conduction when the water is frozen. However, by controlling the water inventory so that no free liquid is available, copper/water heat pipes have been shown to withstand thousands of freeze/thaw cycles during terrestrial testing.
HiK™ plates use copper heat pipes that are flattened and embedded in an aluminum plate to increase the effective thermal conductivity. The heat pipe layout is tailored to most efficiently conduct heat from the electronics to the area where the plate is cooled. Water is the most common working fluid, but methanol can be used when the heat pipe needs to operate at lower temperatures. The benefits of HiK™ plates over encapsulated conduction cards include higher effective thermal conductivity (500 to 1200 W/m K), lower cost, no degradation with thermal cycling, and the ability to conduct heat around corners.
ACT’s copper-water heat pipes and HiK™ plates are commonly used for thermal management of electronics equipment on earth and aircraft but have not been used in spacecraft thermal control applications to date (2017), due to the satellite industry’s requirement that any device or system be successfully tested in a microgravity environment prior to adoption. ACT, NASA Marshall Space Flight Center, and the International Space Station office at NASA’s Johnson Space Center, demonstrated flight heritage in Low-Earth Orbit. The testing was conducted aboard the International Space Station (ISS) under the Advanced Passive Thermal experiment (APTx) project. In the ISS test as shown in Figure 2, the heat pipes were embedded in a HiK™ aluminum base plate, and subject to a variety of thermal tests over a temperature range of -10 to 38 ºC for a ten-day period. Results showed excellent agreement with both predictions and ground tests. The HiK™ plate underwent 15 freeze-thaw cycles between -30 and 70 ºC during ground testing as shown in Figure 3, and an additional 14 freeze-thaw cycles during the ISS testing. The following was demonstrated during 10 days of testing on the ISS:
- Successful operation of the copper/water heat pipes and HiK™ plate
- Ability of the copper/water HPs and HiK™ plate to survive multiple freeze/thaw cycles
- Copper/water heat pipes can carry the required power
- Copper/water heat pipes and HiK™ plate can start up from a frozen state
In addition to ACT’s current aluminum/ammonia constant conductance heat pipes (CCHPs), ACT can now offer a broader toolbox for the spacecraft thermal control engineer: spot cooling of electronic devices with copper/water heat pipes, effective heat spreading of electronic boards and enclosures with our HiK™ plates, and efficient heat transport outside the electronics control box to dissipate the heat with our CCHPs. ACT is the only company in the world that can currently offer these capabilities.
Long Term Partner – Satellite Thermal Management
November 2016 Launch of GOES-R
The Aerospace industry demands reliable and high quality products from their supplier. ACT has provided critical support to many high profile programs. ACT’s work with ITT/Harris Corporation on the recently launched Advanced Baseline Imager (ABI) aboard NOAA’s GOES-R satellite is one example of those successful collaborations.
The partnership between ACT and ITT/Harris dates back more than a decade to 2006, when ACT first began manufacturing aerospace grade Constant Conductance Heat Pipes (CCHPs). At that time ITT was developing an advanced meteorological sensor system called the ABI for the GOES-R satellite. Each ABI unit utilizes fourteen unique heat pipe geometries to isothermalize mounting structures and transport excess heat from the electronics to the thermal dissipation radiators. ACT worked closely with the ITT team to successfully manufacture and deliver the complex heat pipes for integration at ITT. At the end of that program, ACT received an outstanding supplier award from ITT for our effective support. According to ITT, “ACT was a critical supplier for what will be an important national asset”, referring to the GOES-R satellite. Since then, ACT has delivered similar CCHPs for the GOES-S, T and U satellites.
To date, ACT has produced space qualified CCHPs for over thirty satellites, and has accumulated over 15 million operating hours on orbit.
Test System for Simulation of ISS JEMS Module Fluid Loop
ACT designed, fabricated and tested a Fluid Control Unit (FCU) for NASA to simulate a thermal fluid loop using FC-72 for heat rejection onboard the International Space Station (ISS). The FCU is designed to provide FC-72 volume flow, pressure and heat rejection for ground testing of instruments destined for the JEMS module on the ISS. To date, ACT has delivered two FCU’s to NASA in support of the CATS and CREAM instruments. The user interface enables control of fluid flow, differential pressure and temperature while logging these values. The design includes 40 micron filters, automated valves, vacuum pump and an accumulator. The vacuum pump and accumulator allow the user to evacuate and charge the instrument. The Fluid Control Units are fabricated using high quality materials and meet and exceed the cleanliness requirements of ISO 14952-2:2003.
Custom Fluid Control Unit Capabilities
- Pump fluid (Fluorinert™ FC-72) at varying flow rates up to 3 GPM at 225 psi (11 liter/min at 1.5 MPa).
- Maintain fluid inlet temperature at a user-specified value within 1°C by rejecting heat to an air-cooled vapor compression system or providing heat using an electric heater
- Maintain system pressure at a user-controlled setting
- Record and display several system parameters, such as pressures, temperatures, and flow rates. This could be done using the onboard data logger or by connecting a laptop
- Pull a rough vacuum and charge the instrumentation system
- Purge the system with Nitrogen and collect purged fluid in the accumulator
- Provide a port and functionality necessary for fluid sampling
- Be the brightest object in the laboratory (FCU color scheme selected by customer)
ACT communicated regularly with our customer throughout the design and fabrication of the FCU to ensure the unit met their needs. Prior to shipping, testing of the FCU demonstrated that the unit met all of the customer’s specifications. Test results were included with the FCU manual which was delivered with the unit. Two FCU’s have been operating at our customer’s facility since 2013.
Fluid Control Unit for NASA JPL
ACT designed, fabricated, and tested a second Fluid Control Unit for the NASA Jet Propulsion Laboratory that was used for ground testing of the Orbiting Carbon Observatory 3 (OCO-3). OCO-3 was built using the spare OCO-2 instrument and is installed at the International Space Station (ISS). More info on the mission: https://science.nasa.gov/missions/oco-3.
The OCO-3 FCU was constructed to provide precisely metered flow at a stable temperature while measuring and recording flow rate, fluid pressures, and fluid temperatures. The system has an integrated membrane contactor which allows for dissolved gases to be removed from the fluid without draining. This is useful since oxygen readily dissolves in the fluorinated working fluid. Also included in the system is a bellows type accumulator which allows for pressurization of the working fluid without direct contact between any gases and the working fluid. Figure 2 shows ACT’s FCU unit during thermal ground testing at JPL.
If you are interested in learning more, contact ACT today.