National Aeronautics and Space Administration

DRAFT Thermal Management Systems Roadmap Technology Area 14

Scott A. Hill, Chair Christopher Kostyk Brian Motil William Notardonato Steven Rickman Theodore Swanson

November • 2010 DRAFT This page is intentionally left blank

DRAFT Table of Contents

Foreword Executive Summary TA14-1 1. General Overview TA14-5 1.1. Technical Approach TA14-5 1.2. Benefits TA14-6 1.3. Applicability/Traceability to NASA Strategic Goals, AMPM, DRMs, DRAs TA14-6 1.4. Top Technical Challenges TA14-6 2. Detailed Portfolio Discussion TA14-7 2.1. Cryogenic Systems TA14-8 2.1.1. Passive Thermal Control TA14-8 2.1.2. Active Thermal Control TA14-10 2.1.3. System Integration TA14-12 2.2. Thermal Control Systems (Near Room Temperature) TA14-13 2.2.1. Heat Acquisition TA14-13 2.2.2. TA14-14 2.2.3. Heat Rejection and Energy Storage TA14-16 2.3. Thermal Protection Systems (TPS) TA14-19 2.3.1. Ascent/Entry TPS TA14-19 2.3.2. Plume Shielding (Convective and Radiative) TA14-22 2.3.3. Sensor Systems and Measurement Technologies TA14-23 3. Interdependency with Other Technology Areas TA14-23 4. Possible Benefits to Other National Needs TA14-23 Acronyms TA14-25 Acknowledgements TA14-25

DRAFT Foreword NASA’s integrated technology roadmap, including both technology pull and technology push strategies, considers a wide range of pathways to advance the nation’s current capabilities. The present state of this effort is documented in NASA’s DRAFT Space Technology Roadmap, an integrated set of fourteen technology area roadmaps, recommending the overall technology investment strategy and prioritization of NASA’s space technology activities. This document presents the DRAFT Technology Area 14 input: Thermal Management Systems. NASA developed this DRAFT Space Technology Roadmap for use by the National Research Council (NRC) as an initial point of departure. Through an open process of community engagement, the NRC will gather input, integrate it within the Space Technology Roadmap and provide NASA with recommendations on potential future technology investments. Because it is difficult to predict the wide range of future advances possible in these areas, NASA plans updates to its integrated technology roadmap on a regular basis.

DRAFT Executive Summary perature radiators for pre- gas, two phase The Thermal Management Systems Technology flow radiators that serve as passive liquefiers) op- Area (TA) cross-cuts and is an enabler for most timized for the given environment is important. other system-level TAs. Technology development Thermal control systems maintain all vehicle in the Thermal Management Systems TA is cen- surfaces and components within an appropriate tered on the development of systems with reduced temperature range throughout the many mission mass that are capable of handling high heat loads phases despite changing heat loads and thermal with fine temperature control. Technologies with- environments. Effective thermal control systems in the Thermal Management Systems TA are or- provide three basic functions to the vehicle/sys- ganized within the three sub-areas of Cryogenic tem design: heat acquisition, heat transport, and Systems, Thermal Control Systems, and Thermal heat rejection while being mindful of the opera- Protection Systems. tional environment and spacecraft system. Tech- Cryogenic systems require special care for nu- nology advances for heat acquisition devices are merous reasons. The primary reason is the large centered on high thermal conductivity materials range of temperatures to which the cryogenic sys- with a high strength-to-mass ratio and increasing tem is subjected. Secondarily, the maintenance the specific energy of the systems (i.e. high and production of cryogenic propellants requires thermal performance and low mass). Once waste large amounts of power which can be a driver for heat has been acquired, it must be transported to some systems of the spacecraft. Due to the Carnot a or for reuse or ultimate penalty, 1 watt of heat at 20K most likely requires rejection to space. The specific technology em- 150-200 W at 300K to maintain it. This dictates ployed for transport is dependent on the temper- the need for very efficient systems so power re- ature and/or heat flux and thus a wide variety of quirements are not increased. Without effective equipment and techniques can be used. The de- insulation, large flow rates of gases will be vented velopment of single loop architectures could save from the tank. Fortunately, the high vacuum and significant weight, reduce system complexity, and low temperatures of the space environment sim- increase reliability of the thermal design of crewed plifies the thermal control of cryogens in some as- systems. An additional heat transport technology pects. requiring development is in the area of heat pipes. The performance and efficiency of cryogenic Loop Heat Pipes (LHP) and Capillary Pumped systems will have to significantly increase in order Loops (CPL) provide significant heat transport to enable the missions being considered over the over long distances with low temperature drop. next twenty-years. New materials capable of as- Thermal energy can also be stored for later use or cent venting without performance loss or physical rejection into a more favorable environment, thus damage and self-healing Multi Layer Insulation significantly reducing the thermal control system (MLI) or other insulation concepts must be devel- mass by smoothing out the effects of peak and oped and demonstrated. Insulation systems that minimum thermal loads as well as the extreme en- are built into cryogenic tank structure and the use vironments. A method of coping with the peri- of low-conductive composite materials will offer odic long-duration extremely-cold environments reductions in the combined structure and insu- that will occur on planets that do not have an at- lation mass fraction while significantly reducing mosphere is to devise a method of ameliorating cryogen boil-off losses. In addition, techniques for the thermal environment which can significant- tailoring regolith properties to increase the ther- ly reduce the required mass of the thermal system mal performance as an insulation system will have design. to be developed as a mission enabler for space- Thermal protection consists of materials and craft operating on other planetary or near-Earth systems designed to protect spacecraft from ex- objects. The development of cryocoolers and oth- treme high temperatures and heating during all er active cryogenic fluid management systems for mission phases. Reusable thermal protection sys- thermal control of cryogenic propellants in space tems (TPS) are also key technologies for hyper- is a high priority and mission enabler for cryogen- sonic cruise vehicles. Despite the current trend to ic fuel depots and long duration missions outside move away from systems requiring this kind of of low Earth orbit (LEO). Overall system goals TPS there is a national need to not only maintain for these systems are for reduced vibration, lower this technology and its manufacturing, but also to mass, and lower specific power. Also, development advance the state of the art (SOA) in several ar- of large capacity liquefaction cycles (e.g., low tem- eas, particularly maintainability, system size, mass, DRAFT TA14-1 and system robustness. Additional technology de- ments would impact almost every figure of merit velopment is needed to increase the robustness (e.g., mass, reliability, performance, etc.). Some and reduce the maintenance required for reus- advancements in TPS technology fall under the able TPS. In the area of hot structures, high tem- category of “game changing,” while others would perature heat pipes hold the promise of providing represent significant advancements in technolo- high heat flux capability far in excess (5-10x) of gy currently available. Implementation of a sin- high temperature materials. Large inflatable/flex- gle-loop thermal control system is a significant ible/deployable heat shields enable the consider- system simplification thereby increasing the sys- ation of an entirely new class of missions – flexi- tem reliability while decreasing integration efforts ble TPS is enabling for deployable entry systems. for the system. Finally, 20 K cryocoolers capable For many exploration missions rigid ablative ma- of 20 W of refrigeration would offer a significant terials are an enabling technology and are need- mass savings in cryogen storage through a signifi- ed for dual- heat pulse reentries and for very high cant reduction of cryogen boil off and would be a velocity entries. Advances are required to signifi- mission enabler for long term cryogen storage for cantly lower the areal mass of TPS concepts, dem- long duration missions. onstrate extreme environment capability, high re- In summary, the Thermal Management Sys- liability, improved manufacturing consistency and tems TA cross-cuts and is an enabler for most oth- lower cost, and dual-heat pulse capability. From er system-level TA’s with specific interdependen- an analytical perspective, recent efforts have re- cies identified with ten of the remaining fourteen vived ablation analysis capabilities but these need TA’s. The primary benefits from investment in the to be further developed to include development technologies outlined for cryogenic systems, ther- of material response/flow field coupling codes, in- mal control systems, and thermal protection sys- tegration of ablation models into standard 3-di- tems are enabling missions, reducing system mass, mensional thermal modeling codes, and ground & increasing system reliability. Finally, the strate- testing to generate data for code correlation and gic roadmap for the Thermal Management Sys- validation. tems TA is balanced between Technology Push & Future missions show the need for higher heat Mission Pull. rejection, cryogenic propulsion stages, and high energy atmospheric reentry trajectories. Based on these criteria, the Thermal Management Systems TA has prioritized the following technical chal- lenges for thermal management systems: 1. Low density ablator materials and systems for exo-Low Earth Orbit (LEO) missions (>11 km/s entry velocity) 2. Innovative thermal components and loop architecture 3. 20K Cryocoolers and Propellant Tank Integration 4. Low Conductivity Structures/Supports 5. Inflatable/Flexible/Deployable heat shields 6. Two-phase Heat Transfer Loops 7. Obsolescence-driven TPS materials and processes 8. Supplemental Heat Rejection Devices (SHReDs) 9. Hot structures 10. Low temperature/power cryocoolers for science applications Successful development of the various tech- nologies captured under the Cryogenic, Ther- mal Control, and Thermal Protection System ele-

TA14-2 DRAFT Figure 1: Thermal Management Systems Technology Area Strategic Roadmap (TASR)

DRAFT TA14–3/4 This page is intentionally left blank 1. General Overview pressors. Many materials used in heat exchangers have large decreases in conductivity and specific 1.1. Technical Approach heat as they approach 30K. Then there is brittle- The Thermal Management Systems Technology ness issues limiting the classes of metals designers Area (TA) cross-cuts and is an enabler for most can work with. In addition to the material com- other TAs. Thermal management runs the gamut plexity are thermodynamic considerations. Due to from milliwatt cryogenic fluid management sys- the Carnot penalty, 1 watt of heat at 20K most tems to megawatt thermal control systems for nu- likely requires 150-200 W at 300K to maintain clear propulsion architectures; from achieving zero it. This dictates the need for very efficient systems boil off (ZBO) for large scale in-space cryogenic so power requirements are not increased. Finally, fluid storage systems to protection of vehicles to thermal control is important since a large quanti- aerothermodynamic heating during reentry at ve- ty of super-cold fluid depends on it. Without ef- locities of 11 km/s and higher; and from integra- fective insulation, large flow rates of gases will be tion of vehicle structure and the thermal manage- vented from the tank. This has a major impact on ment system to insulation systems that also serve mission architecture. Fortunately, the vacuum and as micrometeoroid and orbital debris (MMOD) ultra-cold sink temperature of the space environ- protection. Technology development in the ther- ment help to simplify thermal control of cryogens mal control area is centered on the development of as compared to on Earth. systems with lower mass that are capable of han- In its most basic form, thermal control is the dling high heat loads with fine temperature con- maintenance of all vehicle surfaces and compo- trol. Technologies within the Thermal Manage- nents within an appropriate temperature range ment Systems TA are organized within the three throughout the many mission phases despite sub-areas of Cryogenic Systems, Thermal Control changing heat loads and thermal environments. Systems, and Thermal Protection Systems. For satellites this requires that the thermal control As long as chemical propellants are the most ef- system must maintain all of the equipment with- ficient primary propulsion systems used in space, in its operating and/or storage temperature range. there will be the need for cryogenic propellants. Similar to the system for satellites, the thermal The performance of LOX/LH2 engines surpass- control system for human-rated vehicles must also es other competing technologies. However, cryo- maintain all of the equipment within the appro- genic systems require special care for numerous priate temperature ranges. In addition to compo- reasons. First is the large temperature range the nent-level temperature maintenance, the crewed system must endure. This has a wide effect on ma- spacecraft’s thermal control system must also safe- terials properties. Control or heat rejection from ly maintain the internal cabin temperature within this temperature range require large amounts of the proper temperature range to ensure both crew power to produce the propellants, which can be survivability and comfort. This section focuses a driver for some systems of the spacecraft. How- on the technologies required to maintain thermal ever, with proper design thermal management of control of the vehicle within the "mid" level tem- cryogenic propellants may be easier in space than perature range. on Earth. Many of the cryogenic technologies de- An effective thermal control system must pro- tailed in this TA are driven by the mission pull vide three basic functions to the vehicle/system of in-space cryogenic servicing needs of chemical design: heat acquisition, heat transport, and heat propulsion stages in the current Human Explo- rejection while being mindful of the operational ration Framework Team (HEFT) Design Refer- environment and spacecraft system. The following ence Mission (DRM) and the potential Flagship sections discuss the critical technologies required cryogenic storage and transfer mission. However, to advance these three functions with the under- significant technology push opportunities exist as standing that some of the proposed technologies materials advances allow for development of effi- overlap two or more functions and each function cient low TRL heat and energy transport process- is dependent to some degree upon the other two. es at very low temperatures. Thermal protection consists of materials and Material properties tend to change as they are systems designed to protect spacecraft from ex- operating in the cryogenic regime. One of the treme high temperatures and heating during all most obvious is variations in coefficient of ther- mission phases. Reusable thermal protection sys- mal expansion between materials, which can af- tems (TPS) are also key technologies for hyper- fect rotating equipment such as and com- sonic cruise vehicles. Extreme high temperatures

DRAFT TA14-5 and heating may be due to not only aerothermo- 1.3. Applicability/Traceability to NASA dynamic heating effects but engine plume and ex- Strategic Goals, AMPM, DRMs, DRAs haust heating effects as well. Zero boil off cryogenic storage in space has been Development of new thermal protection mate- a feature of many past and current NASA archi- rials, systems, and technologies requires extensive tectures, including Mars, Constellation Lunar, testing using unique facilities such as arc jets and and current HEFT DRMs. In addition, the Space radiant heat chambers. Development of high fi- Operations Mission Directorate (SOMD) contin- delity analytical models and the associated tech- ues to help pull the state of the art in sensor cryo- niques, anchored in test and flight data, with an genic technology. understanding of the physics of heat transfer, The Fundamental Aero goals as listed in the stress, surface chemistry, interaction with the aero- Agency Mission Planning Manifest (AMPM) in- thermodynamic convective and radiative heating clude hypersonics elements that are directly sup- environments and relevant pressures and enthal- ported by the TPS technologies presented here, pies, decomposition chemistry, and overall sys- including enabling heat pipe technology. Ablative tem performance is also required. Hence, testing TPS technology advancement is explicitly identi- and analysis, as well as a thorough characterization fied as a must for DRM 2B, and inflatable TPS of material properties are assumed to be integral development is critical for all DRMs whose ulti- parts of technology development and maturation. mate end is to land a large payload on Mars. TPS Finally, it should be noted that TPS technology Health Monitoring Systems (HMS) and integrat- is integral to Entry, Descent, and Landing (EDL). ed thermo-electric generators (TEGs) have been The entry TPS technologies identified under this identified as technologies that may potentially en- TA are consistent with key technologies identified hance any exploration mission. during the EDL technology road mapping process 1.4. Top Technical Challenges and have been fully coordinated with the EDL TA team. Future missions show the need for higher heat rejection, cryogenic propulsion stages, and high 1.2. Benefits energy atmospheric reentry trajectories. Based on Successful development of the various tech- these criteria, the Thermal Management Systems nologies captured under the Cryogenic, Thermal TA has prioritized the following technical chal- Control, and TPS elements would impact almost lenges for thermal management systems: every figure of merit (e.g., mass, reliability, perfor- 1. Low Density Ablator Materials and Systems mance, etc). Some advancements in TPS technol- for Exo-Low Earth Orbit (LEO) Missions ogy fall under the category of “game changing” (>11 km/s Entry Velocity) – For many (e.g., inflatable TPS for large mass payload deliv- exploration missions, such as near-Earth ery to Mars, heat pipes for hypersonic cruise vehi- asteroid and Mars missions, ablative materials cle), while others would represent significant ad- are an enabling technology and are needed vancements in technology currently available (e.g. for dual heat pulse reentries and for very high lighter, cheaper, smaller, more robust, environ- velocity entries (i.e., >11 km/s). mentally-friendly insulation materials with few- 2. Innovative Thermal Components and Loop er maintenance requirements and built-in energy Architecture – An enabling thermal technology harvesting). A TPS fitting the previous description offering significant mass and power savings would save precious spacecraft weight, thereby in- and increased reliability will result from creasing performance and payload capacity. Im- more efficient systems capable of operating plementation of a single-loop thermal control sys- over a wide range of heat loads in varying tem is a significant system simplification thereby environments (for example, a 10:1 heat load increasing the system reliability while decreasing range in environments ranging from 0 to 275 integration efforts for the system. Finally, a 20 K K). A system level approach should be taken cryocooler capable of 20 W of refrigeration would which includes advanced fluids, advanced offer a significant mass savings in cryogen storage radiator design, and other components. through a significant reduction of cryogen boil 3. 20K Cryocoolers and Propellant Tank off and would be a mission enabler for long term Integration – Active thermal control of cryogen storage for long duration missions. cryogens in space can eliminate boil off and dramatically decrease required propellant mass for long duration space missions.

TA14-6 DRAFT Development of low temperature cryocoolers coatings, material characterization, structural and cryocooler to tank integration techniques design and manufacturing processes, and life are needed. and damage assessment methods will enable 4. Low Conductivity Structures/Supports the design optimization of advanced re-entry – Current propulsion stages use high and hypersonic flight vehicles. conductivity aluminum as supports, leading 10. Low Temperature/Power Cryocoolers for to high heat leak. Low thermal conductance Science Applications – Advanced low or reconfigurable supports will reduce this temperature cryocooler technology enables heat leak and minimize power requirements operation of detectors for scientific observation for active cooling systems. of the universe. Advances in size, efficiency and 5. Inflatable/Flexible/Deployable Heat Shields -- reduced vibration/interference are needed. Analytical studies have shown that large heat shields provide a means to increase the down- 2. Detailed Portfolio Discussion mass to the Martian surface. Large inflatable/ The Thermal Management Systems TA has flexible/deployable heat shields enable the identified and detailed numerous technologies in consideration of an entirely new class of the following subsections that are a mix of both missions. Technology Push and Mission Pull. The missions 6. Two-phase Heat Transfer Loops – This that have been identified as Mission Pull candi- technology allows the transfer of small or dates for technologies from this TA are identified large amounts of (typically a across the top row of Figure 1 which is identified 1:100 ratio) over long distances, with very as Major Milestones. little temperature drop. Advanced two-phase For Cryogenic Systems, the Mission Pull oppor- loops allow heat load sharing thus conserving tunities are the Cryostat Demonstration and the energy. Cryogenic Propulsion Stage which will require 7. Obsolescence-Driven TPS Materials and advanced multi-layer insulation and high-capacity Processes – This effort continues development 20K cryocoolers. Thermal Control Systems have of replacement cryoinsulation, primer, identified Mission Pull opportunities for advanced adhesive, and ablator TPS materials that are phase change materials, advanced thermal control currently facing obsolescence. These four system fluids, and variable heat rejection radiators classes of materials are each subject to unique which are demanded by NEO pre-cursor robotic obsolescence issues that will limit their and Crew-to-LEO missions. Technology push op- availability for future programs. portunities for Thermal Control Systems include 8. Supplemental Heat Rejection Devices high temperature materials and components for (SHReDs) – Future technology development megawatt systems; high flux cooling with precise efforts should focus on heat rejection hardware temperature control; and advanced heat exchang- required for transient, cyclical applications. ers and lightweight radiators. Finally, for Thermal Depending on the duration of the mission Protection Systems, the Mission Pull suite of tech- phase, this function can be accomplished nologies that are identified are rigid ablative TPS using either Phase Change Material (PCM) which will be pulled by Crew-to-LEO, Mars Sam- heat exchangers or evaporative heat sinks. An ple Return, Mars pre-cursor, Hypervelocity Earth evaporative utilizes a consumable Return Demo, and Crewed NEO missions; ob- fluid and future development efforts should solescence-driven TPS will be pulled by Crew-to- focus on the efficient use of this consumable LEO; and structurally integrated TPS and multi- when an evaporator is used as a SHReD. functional TPS will both be pulled by MMSEV PCM heat exchanger development, on the and Deep Space Habitat. other hand, should focus on improving the The following subsections are devoted to the de- energy storage capacity of these devices while scription of the current state of the practice, limita- minimizing the hardware mass. Particular tions of the practice, identification of the technol- attention should be focused on combining the ogies to exceed these limitations, and an estimate function of PCM with radiation shielding for of the current technology readiness level (TRL) crew members. and timeframe required to advance the technol- 9. Hot Structures – Advancements in high- ogy to TRL 6 for the Thermal Management TA. temperature materials, environmental The first-level Technical Area Breakdown Struc-

DRAFT TA14-7 Figure 2. First-Level Thermal Management TA TABS ture hierarchy is classified via the temperature re- Thermal Control, and System Integration. The gimes that thermal management systems are re- lower-tiered TABS for Cryogenic Systems is pro- quired to operate: Cryogenic Systems, Thermal vided in Figure 3. Control Systems, and Thermal Protection Sys- 2.1.1. Passive Thermal Control tems. This first-level hierarchy of the TABS is pro- vided in Figure 2. 2.1.1.1. Large-Scale Multi-Layer Insulation 2.1. Cryogenic Systems (MLI) The first major technology area within the Ther- MLI systems have been in use for in space cryo- mal Management TABS is Cryogenic Systems genic propulsion applications for many years but which refers to those systems that are operating use has been limited in size and performance. Ex- below -150 °C. Cryogenic Systems is further dis- amples include the Atlas and Delta upper stages cretized into Passive Thermal Control, Active which typically have three layers of MLI and are

Figure 3. Cryogenic Systems TABS TA14-8 DRAFT intended for a few hours use. Evaporation rates can repair damage from handling or micromete- are on the order of 2 percent per day. Future ap- oroids while maintaining thermal performance plications such as Earth departure stages require should be investigated. These self healing systems larger volumes and much longer orbital storage are perhaps 15-20 years in the future. A more near timelines. This will require an order of magni- term goal is insulation built into tank structures, tude increase in thermal performance with passive such as evacuated honeycomb tank walls or aero- evaporation rates on the order of 0.2 percent per gel filled annular tanks. Multifunctional systems day. Application of large number of layers and in- that serve as cryotank insulation as well as high tegration with vapor cooled shields on high sur- temperature thermal protection is an ideal long face area tanks must be demonstrated. Of partic- term goal. Demonstrations of these cross cutting ular interest are methods of minimizing losses at capabilities will take approximately 8-10 years. seams and penetrations and other areas of chang- 2.1.1.4. Ground to Flight Insulation ing geometry by overlapping of layers, heat sta- A high percentage of the overall heat transferred tioning blankets, or using hybrid aerogel/MLI sys- to flight tanks occurs during the ascent phase of tems. Materials capable of ascent venting without the mission. There is as much thermal energy performance loss or physical damage must be de- transfer during the ascent phase as there is dur- veloped and demonstrated. Currently, high per- ing 6 days of steady state orbital operations us- formance large scale MLI systems for space appli- ing conventional MLI. MLI is very effective while cations is at Technology Readiness Level (TRL) in vacuum but not as good in soft vacuum or at- 3 and successful development to TRL 6 will take mospheric conditions as other insulation meth- 3-5 years. ods. Cryopumping of atmospheric moisture can 2.1.1.2. Advanced MLI Systems also damage MLI and hurt on orbit performance. Current MLI concepts utilize a number of layers Hybrid insulation schemes that are effective dur- of radiation shielding separated by layers of low ing ground and ascent phases while still offering thermal conductance spacers to minimize con- optimal performance for long duration on orbit duction losses across the radiation layers. New storage are needed. Foam/MLI and aerogel/MLI MLI concepts have been proposed that eliminate hybrid schemes are potential options for develop- the need for low conductivity layers of paper be- ment. Hydrophobic materials or coatings can be tween the radiation shields. These new insulations considered. Deployable ground insulation pan- use more rigid metallic layers separated by a sys- els which work during launch countdown but tem of discrete molded polymer spacers to pre- are then detached to minimize mass to orbit are cisely control spacing and layer density. Lower another potential solution. While hybrid foam/ mass and higher performance are predicted bene- MLI systems have been tested in ground cham- fits over current systems. The possibility also exists bers, aerogel/MLI and deployable insulations are to expand these concepts so the outer layer is ca- at TRL 3 and will require 3-5 years of develop- pable of supporting a soft vacuum while on Earth, ment to achieve TRL 6. compressing slightly while being supported by the 2.1.1.5. Low Conductivity Supports spacer system. This offers large performance ben- Conduction heat leak across mechanical sup- efits during the ground and launch phases of the ports such as struts, skirts, and feedlines can be mission where typical MLI systems are not very greater than the convection/radiation heat leak effective. Materials development and process- across the tank surface. Innovative methods of ing and manufacturing improvements needed to minimizing/eliminating that loss are needed. One bring this to TRL 6 will take approximately 3-5 mitigation is to use materials with lower thermal years. conductivity. Low thermal conductivity compos- 2.1.1.3. Multifunctional Insulation/MMOD ite struts are to be used on the James Webb Space Protection Telescope (JWST). Insulation can also offer struc- Integration of multi-functional insulating ma- tural support while still providing thermal perfor- terials into other spacecraft systems can reduce mance, such as load bearing aerogels. Further en- spacecraft mass and increase simplicity. For in- hancements can be made to intercept conduction stance, MLI has shown some ability to serve as an heat leak at a higher temperature by actively cool- effective MMOD protection. Analysis and opti- ing or vapor cooling these solid structures. Meth- mization of MLI systems to increase this protec- ods of integrating heat intercept stations with tion effect is needed. Self-healing materials that supports and feedlines must be proven. The op-

DRAFT TA14-9 timal long term solution is structure that is part needed, so more durable and less fragile insulation of the load path during ground and launch phas- blankets can be developed. In addition, insulation es but disconnects on orbit. Passive orbital discon- optimized for 7 Torr CO2 atmospheres also need nect struts, magnetic levitation, and shape mem- to be developed. There has been some small scale ory alloy materials are all proposed solutions for thermal testing of simulated lunar and Martian this issue. For very sensitive systems, methods of regolith, but research on tailoring regolith prop- electrical power and data transmission without erties for thermal performance is needed. Current conductive wiring could be used. Low conductiv- TRL is on the order of 2 and will require up to 10 ity materials can be developed and proven within years of development to achieve TRL 6. the next 2-3 years but more exotic solutions such 2.1.1.8. Low Temperature Radiators as heat intercept or supports that are capable of disconnect may require 6-8 years of development The effective heat sink temperature of deep space to reach TRL 6. is 3K. This offers potential for a simple source of cryogenic temperature. However, due to the T4 2.1.1.6. Low Conductivity Tanks nature of radiation, low temperature radiators re- Use of composite materials can minimize ther- quire very large areas for any large loads. Pres- mal conduction across tank walls and across tank ently, the low temperature limit is approximate- surfaces. If conduction can be decreased so much ly 50K for LEO applications, and this is for very of the hydrogen tank outer wall temperature is small loads (milliwatt scale). The JWST, which is above 77K, then helium purging can be replaced located at L2, far away from any planetary ther- with nitrogen. Helium is an expensive non-renew- mal load, requires about 16 m2 to radiate approx- able resource and its use should be limited to ap- imately 0.5 watts to deep space. Advances are plications where it is absolutely necessary. This needed including deployable systems, so they are also minimizes cryopumping inside MLI lay- packaged for launch and deployed on orbit with a ers which could affect thermal performance. Sur- much larger radiative area. Ideal materials are flex- face treatments or nanoscale gas barriers on the in- ible for launch but are capable of being rigidized side of tanks to minimize convective heat transfer while maintaining very high emissivity. Concepts from the tank wall to the are possible en- that offer large deployable radiators with integrat- hancements. Non-isotropic heat conduction ma- ed shielding for lunar surface systems or on orbit terials are another solution, where heat is easily spacecraft are needed. Shields that are reconfigu- transported down the length of the tank but has rable to minimize spacecraft orientation require- much higher thermal resistance across the tank ments have benefits. Elimination of parasitic heat thickness. These are definite areas where materi- leads to allow for lower temperature passive cool- als research can push radical changes in propellant ing such as those described in the low conductiv- storage concepts. Composite tanks are currently ity supports section also are applicable here. 50K being tested on the ground, but these have been radiators in the milliwatt range are currently TRL designed for lower mass than corresponding alu- 9, but increasing the capacity to the watts range minum tanks, not necessarily for thermal perfor- while decreasing the operating temperature to mance. Surface treatments or anisotropic materi- 20K will require approximately 8-10 years prior al developments are probably 10 years away from to reaching TRL 6. TRL 6. 2.1.2. Active Thermal Control 2.1.1.7. In Situ Insulation 2.1.2.1. 20 K Cryocoolers for Propellant This section considers the availability of local re- Management sources and environment as a part of the passive thermal control system. Technologies in this area Development of cryocoolers for thermal control are based on ideas such as using lunar or Martian of cryogenic propellants in space is a high priori- regolith as bulk insulation materials in a manner ty. This technology will enable long duration stor- similar to perlite on Earth. Separation of regolith age of cryogenic propellants in orbit. Our current components to remove high thermal conductivi- propulsion stage experience threshold is about 9 ty metals is needed. Use of geographical features hours time on orbit prior to reentry, but will need such as craters on the lunar surface or lava tubes to be extended to 2-3 years for some future mis- sions. Applications include on orbit cryogenic de- to minimize environmental transients must be in situ considered. Insulation schemes that are reusable pots, long duration Mars stages, and re- and reconfigurable for secondary applications are source utilization (ISRU) production on lunar or

TA14-10 DRAFT Mars surface. There is an immediate need for the genic refrigeration for space missions. However, planned Cryostat demonstration mission in 2016. pulse tube coolers have a disadvantage when dis- For liquid hydrogen systems, higher power 20K tributed cooling is required. Methods of convert- coolers are an immediate need. The current state- ing pulsing AC cryocoolers to enable circulation of-the-art (SOA) is pulse tube refrigerators with (DC flow) are sought. Of particular interest are a capacity of 1 W at 20K, with specific power in low- temperature rectifiers so there is no need for the range of 180 W/W. There is a general range a recuperative heat exchanger in the flow stream. of needed capacities listed below, but specific mis- Further optimization of the rectifier into the pulse sions will have exact requirements. Initial require- tube cooler is also needed and characterizing the ments are ~5W at 20K for a Cryostat-scale ZBO performance of the distribution system must be system. Later stages will require 20W at 20K for done. Current TRL is 4 and development to TRL larger storage volumes and possible propellant 6 will take 3-5 years. conditioning in space. Both of these applications Effective recuperative heat exchange is a criti- should target a specific power of 100 W/W. Re- cal function for DC cycle coolers such as Joule- duction in power input and heat rejection require- Thompson (JT) and Brayton cycles, as well as ments is critical, as leads to mass savings in other open cycle liquefaction systems. Current recuper- spacecraft systems. Eventually even higher power ators tend to be the heaviest components in space systems will be required for liquefaction of hydro- cooling systems. The current SOA is perforated gen gas produced by ISRU. Such systems could plate designs. Major losses come from longitu- include open cycle liquefiers, closed cycle refriger- dinal conduction from warm to cold ends. Ways ators, or a hybrid option of the two. Overall sys- of maximizing conduction across the flow paths tem goals are for reduced vibration, lower mass, while minimizing conduction on the flow direc- and lower specific power. Staging for minimal tion are needed. Anisotropic materials would be power requirements is necessary. TRL for 5W and an ideal solution. These materials include heat 20W at 20 K systems is currently at 4 and devel- transfer surfaces such as carbon or metallic foams opment up to TRL 6 will take 2-3 years. High- that can act as fins to increase radial heat transfer er power stages for ISRU liquefaction will take an and provide support for thinner walls. Microplate additional 3-5 years. heat exchangers that use very thin walls and long In addition to cryocooler systems, specific com- flow paths that can be easily produced with maxi- ponent development is needed. These compo- mum flow uniformity within the channels are also nents include regenerative and recuperative heat an option. Several companies are working on suit- exchangers, rectifiers, and turboalternators. able solutions and it is estimated that higher per- Materials that have high thermal conductivity formance options will take 3-5 years to fully de- and high heat capacity at low temperatures (be- velop to TRL 6. low 30K) are needed for low temperature Alter- Turboalternators provide a method of recover- nating Current (AC) cycle regenerative heat ex- ing the work produced by the turboexpander for changers. Current higher temperature coolers use increased cryocooler efficiency. The use of turbine stainless steel or lead materials that have a dramat- shaft work to drive a compressor can be investi- ic decrease in heat capacity as temperatures ap- gated. Microturbines with low leakage and bypass proach 20K. Rare earth elements such as Er-Pr rates when using hydrogen and helium need to have better thermal properties at those tempera- be developed. It is also important to increase the tures but have mechanical limitations. Processing thermodynamic efficiency of the expansion pro- methods for rare earth elements to make regener- cess. Alternatives to gas bearings such as magnet- ative screens, beads, fibers or foils in reproducible ic bearings should be examined. Miniature foil and cost effective manner need to be developed. It bearings would allow numerous start/stop cycles. is important to characterize the mechanical prop- Turboexpanders that can handle multiphase flow erties and address potential lifetime issues and fail- while remaining balanced and operating correct- ure modes. Alternatives to the rare earth materi- ly on the bearings will be important. Some work als currently being used in this application should is being done on a prototype 20K turboalternator, also be investigated. Currently, the TRL of low and this could achieve TRL 6 within 1 year. More temperature regenerators is 4 and will take 3-5 advanced options such as magnetic bearings and years to raise this to TRL 6. integrated turbine/compressor devices will require Pulse tube cryocoolers have proven to be a re- 5-8 years for development. liable and effective method of producing cryo-

DRAFT TA14-11 2.1.2.2. High Capacity Cryocoolers current cryocooler systems at 20-70K. Higher Current pulse tube cryocoolers have scaling is- density magnetic materials are needed. New para- sues when increasing the capacity beyond the cur- magnetic materials are needed for ADR operation rent SOA. It will be important to scale up pulse at 20mK. For all applications, cryocooler devel- tube cryocoolers so they can produce hundreds of opment goals included minimization of mass and watts of cooling at lower temperatures. Acoustic specific power, and reduced vibration levels. streaming phenomenon such as DC flow in dual 2.1.2.5. Cryopumps orifice pulse tubes should be eliminated. Radial Development of cold gas compressors and cold flow and turbulence issues have to be understood liquid circulating pumps with long life, variable and controlled. Characterizing proper phase an- speed operation, and very low leakage are need- gles for larger geometries will potentially be differ- ed for intermediate heat transport loops. Cold ent than for smaller coolers. Development of these control valves with low heat leak are also need- systems is estimated to require 8-10 years to reach ed. Piezoelectric materials have potential to offer a sufficient TRL. lower heat leaks from actuators. Advances in com- 2.1.2.3. In Space Liquefaction Cycles pressor design using non-contact bearings such as Eventual human bases on the Moon or Mars magnetic or foil configurations will increase sys- will rely on in situ resource utilization to produce tem reliability of these components. These tech- necessary propellants and life support consum- nologies are important for cryogen processing and ables. For propellant use, liquefaction will be re- sub-cooling. Development of these systems will quired. The current SOA is large scale liquefaction require 3-5 years to reach TRL 6. and production plants on Earth. Development 2.1.2.6. Thermal Energy Storage of large capacity liquefaction cycles that are op- Advanced fluids systems that are sub-cooled timized for the given environment will be impor- or even solid that can store refrigeration energy tant. This includes low temperature radiators for at some times while providing thermal margin at pre-cooling gas as well as potential two phase flow other times so cryocooler loads can be balanced radiators that serve as passive liquefiers. Integra- and minimized. Sufficient sub-cooling on the tion of the cooling cycle with the ISRU plant form ground also increases the vent-free storage time in a thermal perspective, including effective recuper- early mission phases dramatically. Advances in this ator heat exchangers and high pressure electroly- area can help push rocket motor development to- sis systems that serve as compressors for the lique- wards densified propellants. To maximize this ad- faction cycle. Methods of expanding isentropically vantage, utilizing the heat of vaporization to ab- with multiphase working fluids need to be inves- sorb heat leak can be accomplished. An additional tigated. Development of these liquefaction cycles benefit of solid to liquid phase change storage sys- will require extensive integration with ISRU sys- tems is minimization of convective heat leak into tems and will require 10+ years of development. the tank. Basic liquid sub-cooling systems can 2.1.2.4. Low T, Low Q Coolers for Science reach TRL 6 within 3-5 years but phase change Instruments systems will require 8-10 years. Applications include a variety of space based sci- 2.1.3. System Integration ence platforms including space telescopes, Earth observing systems, and a variety of instruments 2.1.3.1. Shields ranging from X-ray to infrared (IR). The use of deployable shields to eliminate radia- Development of low power 35K, 10-6K and 2K tion to storage tanks and cold instruments should cryocoolers is required to cool the next genera- be investigated. Radiation shields offer critical tion of science instruments. 35K coolers are need- thermal protection for low temperature systems ed for mercury cadmium telluride long wave IR and are currently being baselined for the JWST. detectors. 10K - 6K coolers are needed for arse- Future development will be closely related to ad- nic-doped silicon detectors which operate in the vances in flexible materials that have low emissivi- IR. 2K coolers are needed as upper stage for lower ty and non diffuse surfaces. Adjustable shields and temperature Adiabatic Demagnetization Refriger- louvers to control the thermal energy flow to the ation (ADR) systems for X-ray spectrometers. Ad- system are another example that is at a relatively vances in magnetic materials for ADR’s to increase high TRL. Actively cooled shields that intercept the temperature regime to include high temper- heat at higher temperatures will also be needed. ature superconductors could offer alternatives to The overall goal is to increase the effectiveness of

TA14-12 DRAFT active cooled shields and distributed cooling sys- cooler system which will increase efficiency. Meth- tems. Use of shields for in-space propellant stor- ods of removing waste heat from coolers such as age will require 3-5 years development to reach integrated heat exchangers and liquid cooled com- TRL 6. pressors should be studied. Heat exchangers with 2.1.3.2. Heat Transport more than two working fluids should also be de- Integration of cryocoolers with science instru- veloped for integrated heat transport. These sys- ments is a critical function. Ways to transport the tems are used extensively in ground applications cooling capacity efficiently across small distanc- and transition to space systems should be relative- es on spacecraft have to be advanced. Distributed ly straightforward and take 2-3 years to reach TRL cooling systems including Brayton and JT coolers 6. have to be characterized. Advances in heat pipes 2.1.3.5. Superconducting Systems and loop heat pipes with higher conductance in Eventual power transmission applications will the 10K temperature range need to be achieved. benefit from using superconducting tapes to elim- Advances in heat switches and heat pipe technolo- inate electrical resistance heating, both reducing gy are required to enable high conductance in the waste heat as well as reducing overall power re- “on” position and low conductance when “off”, es- quired. Such systems could be also incorporated pecially for temperatures below 10K. The ultimate into motor windings for a superconducting mo- goal is thermal superconductivity, and successful tor and possible for low voltage instrumentation development can push the incorporation of cryo- to eliminate conduction losses on extra sensitive genic systems into the mainstream energy trans- science packages. Higher temperature supercon- port sector. Surface treatments for materials may ducting magnets have projected applications in allow for lower thermal contact resistance. Evo- science missions. Development of active and pas- lutionary advances in heat transport systems are sive cryogenic systems for both low temperature possible over the next 3-5 years but revolution- and high temperature superconducting applica- ary advances such as surface treatments and inte- tions is needed. There are opportunities for tech- grated switch/pipe systems may require 6-8 years nology push in this area. to reach TRL6. 2.2. Thermal Control Systems (Near Room 2.1.3.3. Staging Temperature) Integration of 20K coolers will use cryocool- The next major technology area within the Ther- er staging at higher temperatures. To optimize mal Management TABS is Thermal Control Sys- the system for minimal power and mass, staging tems which refers to those systems that are op- of the coolers and heat intercept at higher tem- erating between -150 °C and 500 °C. Thermal peratures is needed. Methods of higher tempera- Control Systems is further discretized into Heat ture shielding and integration into passive thermal Acquisition, Heat Transfer, and Heat Rejection control scheme are needed. Heat intercept across and Energy Storage. The lower-tiered TABS for solid interfaces such as struts, skirts, and feedlines Thermal Control Systems is provided in Figure 4. should be developed. Heat transport systems such 2.2.1. Heat Acquisition as circulation pumps, AC/DC flow rectifiers, and Heat acquisition is the process of acquiring ex- DC cryocooler flow control are needed to effec- cess thermal energy from various components in- tively transport cooling loads across distances of cluding power, , avionics, computers, meters and areas of tens of square meters. Multi- and metabolic loads from crewmembers. Heat ac- stage 20K coolers with high temperature heat in- quisition is typically accomplished using a myriad tercept will require 2-4 years of development to of hardware components: mainly, but not limited reach TRL 6. to, coldplates, air/liquid heat exchangers, and liq- 2.1.3.4. Integration with High Temperature uid/liquid heat exchangers. Heaters are also con- Systems sidered a component of heat acquisition and are Cryocooler systems that reject waste heat to generally used to address temperature differences warmer temperature environmental control and created by the environment or differences in heat life support system (ECLSS) cooling loops as op- generation between components or areas of the posed to dedicated radiators at higher tempera- spacecraft. tures should be investigated. This allows for a lower 2.2.1.1. Coldplates and Heat Exchangers ΔT between the cold and warm sides of the cryo- Heat exchangers and coldplates are presently

DRAFT TA14-13 Figure 4. Thermal Control System TABS made of metals such as aluminum and stainless of investment is very low in developing highly re- steel. Both heat exchangers and coldplates include liable heat generation devices. The SOA has not heat transfer fins which are required to enhance changed much over the past 50 years. The major the unit’s heat exchanger efficiency. Higher effi- challenges include increasing reliability and mini- ciency/lower mass designs can be realized through mizing overall waste heat generation. New materi- the use of micro-channel fabrication techniques or als and heater concepts to increase iso-thermality, the use of composite materials. Composite mate- increase control reliability, and minimize burnout rials are desirable for the fabrication of heat ex- or thermal fatigue are critical. Improved sensors changers due to their potentially high thermal and are desired to allow more precise conductivity and high strength-to-mass ratio (es- temperature control and better redundancy. These pecially composites enhanced by nanotechnol- devices are at TRL 1-3 and could be advanced to ogy). Micro-channel fabrication techniques can TRL 6 in 7-10 years. also be used to reduce the fin spacing by an or- 2.2.2. Heat Transfer der of magnitude increasing the thermal perfor- mance while also dramatically reducing the hard- Once waste heat has been acquired, it must be ware mass and volume. Both concepts are at a transported to a heat exchanger or radiator for re- TRL of 2-3, but can be advanced to TRL 6 in 3-5 use or ultimate rejection to space. The specific years. The materials development work should be technology employed for transport is dependent coordinated with the Materials, Structural & Me- on the temperature and/or heat flux and thus a chanical Systems, and Manufacturing (TA12) TA, wide variety of equipment and techniques can be but requirements and testing should be led by the used. In some cases, heat transfer is not desired Thermal Management TA. Improvements in cool- or must be tightly controlled and technologies to ing and dehumidification of airflow in condens- limit or prevent heat transfer are also critical. ing heat exchangers are the responsibility of the 2.2.2.1. Insulation Human Health, Life Support and Habitation Sys- Thermal transport of waste or to-be-reused heat tems (TA06) TA. must first be controlled within a defined path. In- 2.2.1.2. Heaters sulation is critical to this function. This is typi- An efficiently designed spacecraft would use cally accomplished via MLI, specialized thermal available waste energy to provide heat where re- coatings such as gold or various paints on a space- quired, using as few heaters as possible. While ad- craft body, or aerogels or other solids in situations vanced two-phase loops can provide this function, where MLI or coatings are ineffective such as on engineering compromises require that heaters be balloons or planetary surfaces. Room temperature used to address the cold component tempera- MLI technology is well established. One area still tures caused by the environment and to provide at a lower TRL (4-5) is with insulation intend- condensation control for internal cabin surfaces. ed for very high temperature environments (e.g., Heaters consume power and frequently add to the aerogels or special MLI), such as for missions go- overall heat load for the rejection system. The re- ing close to the Sun or landing on Venus. Plane- liability of heaters and their associated control cir- tary environments such as Venus present a special cuits are low enough that generally at least one re- case since there is both a dense atmosphere and dundant circuit is required. NASA’s planned level very high temperature. In this situation MLI is TA14-14 DRAFT useless and advanced high temperature/high per- proved thermophysical properties. Advanced ther- formance insulations will be needed. Additional- mal fluids are currently at TRL 2-3 and could be ly, high performance insulation systems, which are advanced to 6-7 in 3-5 years. more easily fabricated than traditional MLI sys- 2.2.2.3. Advanced Pumps tems, are desired for both hot and cold environ- Mechanical pumps are often used to move flu- ments. Both of these technologies could be ad- id through a thermal control system. Advanced vanced to TRL 6-8 in 3-5 years. technologies such as piezoelectric pumps and/ 2.2.2.2. Transport Fluids or electrohydrodynamic pumps offer very high For robotic spacecraft where human life support pumping efficiencies (>90%), longevity and neg- is not an issue, the most common heat transport ligible vibration. Reliable check valves would al- fluids are ammonia, fluorocarbons, and propylene low the inclusion of a redundant , if needed. for room temperature applications with meth- Current TRL for these technologies is around 3-4 ane, ethane, nitrogen, or oxygen employed for and could be advanced in 5 years to TRL 6. cryogenic applications (dependent upon specific 2.2.2.4. Heat Straps temperature). Such fluids are used for both sin- gle-phase applications (typically via the mechan- Occasionally it is necessary to passively trans- ical pumping of a liquid) or for two-phase ap- port heat from one specific location to another plications (transported either mechanically or by within the spacecraft itself or to spread the heat capillary forces). load across a wider area. In these situations it is For human tended spacecraft, it is desirable to often convenient to use a heat strap, which is a develop technologies that enable single-loop ther- mature technology that typically uses higher mass mal control system architectures rather than the materials such as and aluminum. Emerg- current state-of-art, which is an internal/exter- ing technologies using ultra high conductivity car- nal two-loop architecture. The single-loop archi- bon fibers, artificial diamond films, carbon nano- tecture would have benefits of improved reliabili- tubes, boron nitride nanotubes, (now at TRL 2-4) ty, system simplicity, and significant mass savings. or other such materials promise lower mass/ther- Typically, thermal control systems, including the mal conductivity and could be developed to TRL vehicle’s radiator, are sized for the maximum con- 6 in 3-5 years. tinuous heat load in the longest sustainable ther- 2.2.2.5. Heat Switches mal environment. This design approach results in Heat switches are often employed with heat a fairly large radiator surface area. Unless some straps, or at connection points between equip- sort of protection is provided, the heat transfer ment. These devices allow, or prevent, a thermal- fluid will need to have a relatively low freeze tem- ly conductive link. Heat switches may be designed perature must be incorporated into the design to to be purely passive, by opening or closing at spe- avoid freezing the fluid during cold and low-load cific pre-selected temperatures, or they may be un- mission phases. Unfortunately, most of the fluids der active command. Technology improvements that have low freeze temperatures (fluorocarbons, are possible and the goals are typically higher ther- ammonia, etc.) are also toxic to crew members and mal conductivity with lower mass for the device, cannot be used inside the pressurized spacecraft and/or higher ratios of conductance in the “on” volume for fear of an inadvertent system leak. As a and “off” conditions. Desired performance of heat result, two-loop thermal control system architec- switches is currently at a TRL of 2-3 and could be tures are typically designed to use low freeze point advanced to TRL 6 in 3-5 years. temperatures in the external loop and a more be- 2.2.2.6. Heat Pipes nign fluid in an internal loop which is connected The most efficient heat transport phenomenon via an inter-loop heat exchanger which adds sub- involves convective flow associated with a change stantial mass to the system. in phase of the heat transport fluid. The fluid is Single loop architectures could save significant generally contained within a tube, or separate weight. They could be enabled either by devel- tubes for the liquid and vapor sides, and is contin- oping advanced fluids and/or developing variable uously recycled. The fluid can be circulated either heat rejection radiator technologies. The devel- by a mechanical pump or by the capillary forces opment of advanced fluids should be focused on generated by a wick. Two-phase loops driven by low toxicity while depressing the freeze tempera- capillary forces have the advantage of not requir- ture and ensuring that the advanced fluid has im- ing an external force, such as a mechanical pump,

DRAFT TA14-15 to circulate the fluid. ment, contaminant control, phase separation, Heat pipes are the most common capillary-driv- etc.) are in use, intermittent operation in micro- en device used in spacecraft today. A traditional gravity and in severe environments, such as hard heat pipe includes a hollow tube, sealed on both vacuum, radiation, and extreme temperatures, are ends, with an interior wick and a circulating fluid. real concerns. Exceptionally long life, low mass, Heat pipes can transport up to hundreds of watts high efficiency, and operating with high temper- for several meters, at negligible temperature drop, ature lifts (50 °C or more) are key improvements along the length of the pipe. Heat pipes are a well- for space-based heat pump technology. developed technology, TRL 8-9. A variety of heat Depending upon the application, it may be pos- pipes for cryogenic and high temperature applica- sible to make productive use of the local environ- tions have also been flown (TRL 6-9). These tend ment to improve the operation of a heat pump to be specialty applications and require unique de- based system. For example, in a lunar or other signs, but the basic technology is established. planetary application it may be most efficient to Another type of capillary based, two-phase heat drive the heat pump compressor with a photovol- transport technology involves the use of a loop taic system sized such that it provides 100% of the where there are separate liquid and vapor lines compressor power without the use of heavy and and the wick is located only at the evaporator. expensive batteries. System level TRL is 4-5 and Such loops are termed Loop Heat Pipes (LHP) could be developed to 6-7 in 5-10 years. and Capillary Pumped Loops (CPL). Loops have 2.2.2.8. Thermal Electric Coolers (TECs) been built which can transport 20+ kW over a ten meter length. Conversely, smaller loops have There are special situations where precise, but been built that transport only a few watts to sever- very localized thermal control is needed, and cold al hundred watts. These technologies offer signif- biasing with bump-up heaters is impractical. Ex- icant heat transport over long distances with low amples include spot cooling of electronics or sen- temperature drop, are inherently self regulating, sors and tight control of laser diodes. In such sit- need no mechanical pump, and can last indefi- uations it is possible to use thermoelectric devices nitely. They are capable of very tight thermal con- based on the Peltier effect. Typical applications trol (i.e. +/- 0.1 °C). All of these advantages are are at a higher TRL; however it is currently a ma- very important for scientific spacecraft that have jor challenge to operate TECs below 150 K. The instruments in need of precise thermal control. TRL for the low temperature applications is cur- Existing LHPs and CPLs have only one evapora- rently at 2-3 and could be developed to TRL 6 in tor and one condenser/radiator. Hence, an impor- 5 years. tant advancement would be the development of 2.2.2.9. Architecture and Flow Control LHPs with multiple evaporators and condensers, Current system architectures and available flu- a technology that is currently TRL 5. A flight ex- ids can result in severe limitations on system op- periment would raise the TRL to 6-7 within 2-3 erations. The current SOA two-loop system with years. a regenerator bypass temperature control can op- Better analytical models of two-phase loops are erate at low heat loads in cold environments, but also needed. Current models are functional, pre- its operational heat load/environment envelope is dict transients reasonably well, and are useful for still limited. Systems with wide operating enve- design purposes, but more complete models with lopes that have few, if any, operating restrictions zero-G validated correlations are desired. This is must be developed for future missions. This will particularly true for systems with multiple evap- require innovative system architectures and/or so- orators and condensers. The current TRL of 5/6 phisticated control schemes. Analytical models could be raised to 7/9 within 4 years. to support such system level design and trade-off 2.2.2.7. Heat Pumps studies are also needed. The TRL for advanced ar- Heat pumps have been employed for ground chitectures is very low, but could be developed to applications for many decades and are a relatively TRL 6 in 5 years. mature technology. Earth-based heat pump tech- 2.2.3. Heat Rejection and Energy Storage nology is not directly applicable to the rigors of Heat rejection is accomplished using radiators, the space environment. Although ground-based evaporators, and/or sublimators. Thermal energy designs that are not reliant on gravity for elements can also be stored (either as latent heat in a phase of heat pump operation (e.g., lubricant manage- change material or through sensible heating of a

TA14-16 DRAFT large mass) for later use or rejection into a more allow a substantial reduction in mass by reducing favorable environment, thus significantly reduc- radiator area. A successful coating will be insen- ing the thermal control system mass by smooth- sitive to the effects of environmental degradation ing out the effects of peak and minimum thermal or build up of contamination. Exactly what those loads as well as the extreme environments. degradation mechanisms are depends on the en- 2.2.3.1. Radiators vironment, but the degradation will probably be centered on those components directly exposed, Radiator advancement is perhaps the most crit- such as radiators and thermal protection layers. ical thermal technology development for future For example, the Hubble Space Telescope multi- spacecraft and space-based systems. Since radia- layer insulation underwent dramatic degradation tors contribute a substantial portion of the ther- due to the combined effects of radiation and ther- mal control system mass. For example, the Altair mal cycling as it traveled in and out of Earth’s (Lunar Lander) vehicle radiator design represents shadow. As another example, when dust got on 40% of the thermal system mass. Radiators can be the Apollo lunar roving vehicle (LRV), the result subdivided into two categories; the first is for re- was that the batteries did not cool down between jection at temperatures below 350 K and the sec- EVAs. In all three LRVs (Apollo 15, 16, and 17) ond is for nuclear or high power systems at tem- the batteries were run at higher temperatures then peratures around 500 K. Technology development specified for most of the third EVA. In addition efforts in the low temperature category should fo- to dust effects, there may be atmospheric chemis- cus on heat rejection variability, advanced coat- try interactions with exposed surfaces. Perhaps the ings, and mass reduction. Technology develop- extreme example is Venus with its high tempera- ment for high temperature radiators should focus ture, high pressure, and acidic atmosphere. Pas- on advanced coatings and compatibility with liq- sive (e.g., “Lotus Coating”) or active techniques uid metals or other exotic heat transfer fluids. for cleaning radiator coatings are also being pur- Radiators are typically sized to reject heat during sued, currently at TRL 3-5, and could be ma- the worst case combination of peak heat load and tured to TRL 6-8 within a few years. Prototype least favorable environment. As mentioned earli- versions of variable emittance coatings have been er, for crewed spacecraft, this design often results flown on ST-5 spacecraft, but additional develop- in a two-loop thermal control architecture. The ment is necessary to achieve a dynamic change of two-loop architecture can be replaced by a low- emittance of at least 4:1 and survive in a space en- er mass single-loop architecture by either devel- vironment for several years. There are continuing oping advanced fluids and/or developing variable efforts to gradually improve the coating proper- heat rejection radiator technologies. A variable α/ε heat rejection radiator must be capable of varying ties, but substantial decreases in (throughout its effective heat transfer coefficient with the envi- the entire mission) are critical to next generation ronment depending on the given mission phase. spacecraft. This variability could be accomplished by active- Specialized thermal control systems will be re- ly changing the radiator’s infrared emissivity (e.g., quired for nuclear or other high power spacecraft. a variable emissivity coating), draining the radia- The roadmap for nuclear power for spacecraft is tor fluid, changing the radiator surface area with addressed by the Space Power and Energy Stor- covers, or stagnating/freezing the fluid inside the age Team (TA3), but the key thermal components radiator in a predictable fashion. These technol- required include high temperature liquid met- ogies are at a TRL 2-4 and could be developed al with high thermal conductivity and to TRL 6 in 5-7 years. Hence, there are multiple large, multi-megawatt deployable radiators. Radi- paths towards a technical solution to a significant ator surfaces will be required to reject heat at high challenge, and these involve interlinking push and temperatures (~500K). Current SOA does not ex- pull technologies. ist for high temperature radiator coatings with Thermal coatings are critical for moderate tem- acceptable longevity and emissivity. This will be perature radiator surfaces, and a wide variety of critical to a Rankine-cycle thermal nuclear pow- specialty coatings have been developed over the er system. Close coordination with TA3 will be past 50 years. These coatings typically have an ab- required. The current TRL is at 1-2 and could be α/ε developed to TRL 6 in 7-10 years. In addition to sorptivity to emissivity ratio ( ) of approximate- the hardware, analytical models for high tempera- ly 0.1. Reducing the ratio to near zero would al- ture systems are needed. low effective heat rejection in full sun, which may Finally, lightweight radiators or thermal stor-

DRAFT TA14-17 age devices for extra vehicular activity (EVA) tasks ature and/or increasing the minimum required must be developed for extended missions on ex- heat load to avoid radiator freezing. The specific traterrestrial surfaces. Current SOA involves the process that must be developed is the technique to use of sublimators which consumes significant store energy into modified lunar regolith and then mass. Compact specialized radiators or thermal retrieve the energy. Initial modeling efforts require storage devices could significantly reduce the con- testing in appropriate environments (such as a full sumables required for EVA. The goal is to devel- vacuum). The TRL for this technology is currently op a non-venting, closed-loop heat rejection sys- at a 2-3 and could be advanced to a TRL 6 with- tem with no consumables for EVA missions. The in 2-3 years. current TRL is 2-3 and could be advanced to TRL 2.2.3.4. Heat Sinks and Storage 6 in 3-5 years. Significant technology development is required 2.2.3.2. Two-Phase Pumped Loop Systems for transient, cyclical applications. These scenarios The benefits of two-phase systems are principal- can occur when the thermal environment is vary- ly realized in phase change heat transfer where a ing in a cyclical fashion such as those occurring significant amount of heat is transferred to a rel- during a planetary orbit. Another cyclical applica- atively small amount of fluid through the latent tion is when the spacecraft’s heat rejection require- heat of vaporization. This is approximately a two ment is varying throughout a particular mission order of magnitude improvement in heat trans- phase. The hardware used to accomplish this spe- ferred per unit of system mass. Two-phase systems cific heat rejection function is commonly referred give the additional flexibility of maintaining near- to as SHReDs. This name is used because these isothermal conditions in critical locations over a devices are used to supplement another device wide range of flow rates. such as a radiator. Two-phase heat transport systems, such as Loop There are two primary types of hardware that Heat Pipes, have been used successfully in the can be used as a SHReD. The first, which requires space environment but have been limited to small a consumable fluid such as water, is an evaporative heat load applications (less than 1 kW) that re- heat sink. This device rejects energy by evaporat- quire the isothermal advantages. The advantages ing a fluid and venting the consumable to the am- for large heat load systems, such as those requiring bient environment. The second type is a supple- the use of Rankine cycle power system, have yet to mental, regenerable heat rejection device, which be designed and tested. Furthermore, we still lack has the benefit of not requiring a consumable and a fundamental understanding of the mechanisms is therefore more suitable for long mission dura- involved in pool and flow boiling and condensa- tions. One such device is a PCM heat sink. tion in partial or microgravity environments. Re- The mass of PCM required for a given applica- sults from upcoming experiments on the Interna- tion is inversely related to the material’s heat of fu- tional Space Station (ISS) will provide a TRL of sion and the time of operation. Traditional PCM 2-3 and then should be applied to a larger scale heat exchangers use paraffin as the PCM. How- thermal control system (TCS) to develop TRL 6 ever, these materials tend to have lower heats of in 10-15 years. Development of microgravity sep- fusion leading to relatively heavy PCM heat ex- arators is also required. Other techniques such as changers. Water has a heat of fusion approximate- the imposition of electric and acoustic fields and ly 70 percent higher than a typical PCM with the forced flow on the bubble nucleation and vapor appropriate control (melt) temperature but it has bubble removal should be examined to control unique challenges associated with its use. Un- film dry out increasing heat transfer. like most fluids, water expands significantly when 2.2.3.3. Environmental Amelioration it freezes, which results in unique structural de- A method of coping with the periodic long dura- sign challenges. A concentrated technology devel- tion extremely cold environments that will occur opment effort should be performed to make the on planets that do not have an atmosphere is to use of water as the PCM a viable option for fu- devise a method of ameliorating the thermal envi- ture heat exchangers. Water has another advan- ronment. By trapping heat during the warm peri- tage because it is a very good material for radiation ods and giving it up slowly during the cold periods shielding. Developing a thin water shield around (much as a desert solar house rock-wall does), en- crew living space for exo-LEO missions and us- vironmental amelioration can raise the minimum ing it as a combined SHReD and radiation shield sink temperature, increasing the radiator temper- should be a high priority. TRL for water based sys-

TA14-18 DRAFT tems is currently at 3 and could be developed to ger being manufactured, however future nano- TRL 6 in 3-5 years. structured films hold promise (see TA10 Section Sublimators have traditionally been used for 3.2.1.3, 7 years to technology maturity). There- steady state heat rejection applications. However, fore, there is a need to maintain the present re- previous test programs have shown that the overall usable TPS design and manufacturing capability sublimator feedwater efficiency is quite low when for use with future spacecraft. Additional technol- used as a supplemental heat rejection device. This ogy development is needed to increase the robust- is because of start-up and shutdown inefficiencies. ness and reduce the maintenance required for re- In addition, many sublimators have a minimum usable TPS. For example, the RTV breaks down heat load requirement due to concerns about hard- at 340 °C, but higher temperature adhesives (new ware failure caused by freezing within the feedwa- compounds, or those benefiting from nano-fillers ter reservoir. This minimum applied heat load can [see TA10 Section 3.2.1.4]) would greatly bene- lead to additional inefficiencies because it would fit TPS designers in reducing the insulation need- result in consumable use even when evaporative ed to reduce bondline temperatures. Further- cooling is not truly required. In addition to devel- more, nanostructured materials (see TA10 Section oping sublimators capable of performing supple- 3.2.1.2) promise to greatly increase the damage mental heat rejection, a concerted effort should be tolerance of composites (5-10 years), and poten- performed to develop evaporative heat sinks capa- tially large increases in ceramics (as much as a fac- ble of operating in a wide range of external pres- tor of 1000 in 10-20 years). This would directly sures (post-landing terrestrial cooling, Venus or benefit commercial crew transportation vehicles, Martian environments, etc.). Indeed, along with some of which are envisioned to rely on reusable advanced high temperature insulation, sublima- TPS. This effort should also include development tors will be needed to extend operational times for of thermal barriers, seals, and gap fillers that are a Venus Lander. Ensuring zero carryover would usually required between segments (tiles or pan- allow the use of back pressure control valves for els) and the development of waterproofing agents. temperature control. Sublimators and evaporators Reusable TPS provides benefit not only during as- for this use are currently at TRL 3 and could be cent, but also contributes to the spacecraft’s over- advanced to TRL 6 in 3-4 years. all on-orbit heat balance and serves as entry TPS 2.3. Thermal Protection Systems (TPS) as well. Advanced Reusable TPS may take many forms (all of which require material development The final technology area within the Thermal and characterization) including tiles, blankets, re- Management TABS is Thermal Protection Sys- fractory composite/refractory metallic structures, tems which refers to those systems that are oper- UHTCs, embedded PCMs, coatings, nanostruc- ating between above 500 °C. Thermal Protection tured materials, and heat pipes. Of these TPS Systems is further discretized into Entry/Ascent technologies CMC fabrication technology devel- TPS, Plume Shielding, and Sensor Systems and opment stands out as a priority as the US has fall- Measurement Technologies. The lower-tiered en far behind several other countries in this both TABS for Thermal Protection Systems is provid- promising and critical area. Nanostructured ma- ed in Figure 5. terials could incorporate radiation and MMOD 2.3.1. Ascent/Entry TPS protection as well as tailoring thermophysical 2.3.1.1. Reusable TPS properties with values in excess of current SOA. High temperature heat pipes hold the promise of With the retirement of the Shuttle Orbiter, the providing high heat flux capability far in excess primary user for reusable TPS, which includes (5-10x) over high temperature materials with the Carbon/Carbon hot structures (up to 1650 °C), benefit of being light weight and a passive design. tiles (up to 1260 °C), and fibrous insulation blan- Also included in this category are hot structures, kets (820 °C), will be gone. Despite the current which may include refractory composites (Car- trend to move away from systems requiring this bon/Carbon (C/C), Carbon/Silicon Carbide (C/ kind of TPS there is a national need to not only SiC)), refractory metallics (Inconel, NiTiAl and maintain this technology and its manufacturing, Gamma TiAl systems), and the associated attach- but also to take this opportunity to advance the ment hardware and insulation systems. High TRL SOA in several areas, particularly maintainability, (4-5) items may be developed in 2-3 years of ef- system size, mass, and system robustness. Water- fort. Lower TRL (2-3) items may require 4-5 years proofing agents for tiles, for example, are no lon- of development time.

DRAFT TA14-19 Figure 5. Thermal Protection System TABS 2.3.1.2. Flexible TPS (cross cutting with TA09- quirements for “soft goods.” Concept maturation EDL) to TRL 5 will require extensive ground testing, System studies have shown that large heat while maturation to TRL 6 may require a small- shields provide a potentially enabling means to in- scale component level flight test. Development crease landed mass on the Martian surface. Large time estimated to be 5-6 years. Primary areas of inflatable/flexible/deployable heat shields enable recommended NASA investment include: the consideration of a whole new class of missions. • Non-ablative (insulative or transpiration Flexible TPS is enabling for such deployable entry cooled) material concepts with high flexibility systems, and will provide strongly enhancing ben- and stowability; efits for rigid systems as well (in terms of reduced • Ablative material concepts, including systems life cycle cost and ease of manufacturing, as evi- that rigidize in-space or during entry. Includes denced by Orbiter LCC data). The current oper- high-heat flux semi-flexibles (q>150 W/cm2) ational SOA is the AFRSI blankets employed on for low-cost application to rigid aeroshells the leeward side of the Orbiter, which are reus- 2 (push technology); able systems designed for ~5 W/cm of maximum • Multifunctional materials that provide heating. Future deployable entry systems will re- MMOD/radiation protection and/or carry quire TPS concepts that can be stowed for months structural loads (push technology); and in space and then deployed into an entry configu- ration that can withstand 20-150 W/cm2 of heat- • Improved thermal response and reliability ing on Mars or Earth. These are envisioned as sin- quantification models. gle or dual use systems, therefore multi-mission 2.3.1.3. Rigid Ablative TPS (cross cutting with reusability will not be required. Both non-ablat- TA09-EDL) ing and ablating concepts may be suitable, with All NASA entry vehicles to date have employed the key trade being TPS development complexity rigid TPS, ranging from the reusable tiles on the versus system scalability and controllability. The Orbiter to ablative systems employed for plane- current NASA portfolio includes small invest- tary entry and Earth return from beyond LEO. ments in q<50-75 W/cm2 non-ablative materials For many exploration missions, such as near- (ARMD Hypersonics demonstration 2012), and Earth asteroid and Mars missions, ablative mate- q<150 W/cm2 ablative materials (ESMD ETDD). rials are an enabling technology and are needed All of these materials are low TRL at this time, al- for dual heat pulse reentries and for very high ve- though the first, an insulative 20 W/cm2 multilay- locity entries (i.e., >11km/s). However, the cur- er system, is planned to be flight tested in 2012. rent selection of high TRL rigid TPS materials Major technical challenges include maintaining is inadequate for future mission objectives. TRL thermal and structural properties after long dura- 6+ heatshield materials include PICA (Stardust, tion storage in space, performance under aeroelas- MSL), SLA-561V (MER, Phoenix), ACC (Gen- tic and shear loading, and planetary protection re- esis), Shuttle tiles (Orbiter), and RCC (Orbiter

TA14-20 DRAFT WLE). Advances are required to significantly low- materials with at least 40 percent lower areal er the areal mass of TPS concepts, demonstrate ex- mass than the current SOA (overlap with treme environment capability, demonstrate high TA12). reliability, demonstrate improved manufacturing »»Extreme environment (q > 2 kW/cm2, p > 1 consistency and lower cost, and demonstrate du- atm) materials, including redevelopment of al-heat pulse (aerocapture plus entry) capability. extremely high-reliability Carbon-Phenolic Current agency investments include ablative ma- for Mars Sample Return, and analog materials terials development within ARMD (Hyperson- for future missions. ics) and ESMD (Orion and ETDD), primarily • Improved thermal response models, including in support of crewed return from the Moon and high fidelity ab initio in-depth ablation/ Exploration missions to Mars. It should be noted thermally-coupled response, gas-surface that Avcoat, while flown on Apollo and baselined interactions, and computational materials for Orion, has not yet reached TRL 5 as a re-en- design capability. (push technology, overlap gineered concept. There is currently minimal in- with TA12) vestment in materials concepts for other mission classes and highly innovative multifunctional ma- • Improved processes for quantification of terials. Notably, missions that involve Earth re- TPS margin and system reliability, including turn of crew from beyond the Moon, or robotic statistical analysis, testing techniques, and entry to Venus or the Giant Planets, have no avail- archival storage of agency thermal test data, as able TPS solution at this time. Many other mis- required for crewed vehicles and Mars/Europa sion classes must resort to capable, but extreme- sample return missions. ly heavy, TPS solutions. The addition of carbon 2.3.1.4. In-Space TPS Repair nanotubes and nanofibers to strengthen the char A significant risk with spaceflight is damage layer promises the ability to reduce required abla- to a vehicle due to MMOD impact. The Space tor mass by up to 50 percent (See TA10 Section Shuttle Orbiter on-orbit repair techniques activ- 3.2.1.5, and 3.2.1.6 for development timeframe). ities should be continued to provide a repair ca- Recent efforts under Constellation have revived pability for future spacecraft, both commercial ablation analysis capabilities. These efforts should and NASA-run. Assuming the development ef- be expanded to include development of material forts leveraged off of methods developed for Space response/flow field coupling codes (for both, en- Shuttle Orbiter TPS repair such as the Tile Re- gineering calculations and high-fidelity compu- pair Ablator Dispenser (TRAD) and Reinforced tational solutions), integration of ablation mod- Carbon-Carbon (RCC) crack repair, significant els into standard 3-dimensional thermal modeling improvement in these techniques could be made codes, and ground testing to generate data for code with 2-3 years of effort. Repair techniques for sig- correlation and validation. There is some TPS in- nificantly different TPS architectures, such as in- vestment by other government agencies, but their flatable/deployable heat shields, will require lon- primary focus is on reusable systems to support ger development times. hypersonic cruise. Major technical challenges in- 2.3.1.5. Self-Diagnosing/Self Repairing TPS clude the fidelity of current response models, availability of suitable ground test facilities, high One way to advance the TPS SOA is to study uncertainties in input aerothermal environments more advanced capabilities already displayed (covered in Section 2.1.6) and the inherent con- in nature – in this case, the ability to diagnose flict between low mass and robust performance. and heal itself. The self-diagnosing aspect of this Concept maturation to TRL 5 will require exten- would come from a health monitoring system sive ground testing, while maturation to TRL 6 (HMS) – an area that holds promise with recent may require a small-scale component level flight advances in flight-worthy, quantitative fiber op- test. Development time estimated to be 4-5 years tic sensing, acoustic emission technology, wireless for lower TRL concepts and 2-3 years for high- sensing, other full-field optical techniques, and er TRL concepts. Primary areas of recommended other forms of non-destructive evaluation (NDE). NASA investment include: The self-repairing aspect of this technology is not as mature but has seen some success in low TRL • Advanced Ablator materials efforts. Advances in nanostructured materials »»Low- and Mid-density ablators for Earth (based on nanotubes and self-assembled materi- return from beyond LEO and Mars entry als) would enable both self-diagnoses as well as re- missions, including dual heat pulse capable DRAFT TA14-21 pair. Development of some of the self-diagnosing Ablative TPS, used to protect vehicles from high- technologies is likely in the 5 year range, but ad- er heating environments, requires additional de- vancement of the self-repairing technology would velopment to increase low-temperature substrate probably require 10 - 20 years after advancements adhesion, reduce weight, and to identify replace- in design and manufacturing methods (see TA10 ment weatherproof coatings that have recently be- Section 3.2.1.2). come commercially unavailable. TPS primer sys- 2.3.1.6. Multi-Functional TPS tems, used for both cryoinsulations and ablators, Multidisciplinary approaches to traditional sys- are designed to increase adhesion performance tems like TPS hold potential for enabling mis- and maintain corrosion protection for spacecraft sions currently out-of-reach. The driving motiva- structures. Current primers are based on hexava- tion behind this is, ultimately, the development lent chromium materials which continue to have of efficient spacecraft structures and systems. For reduced availability due to migration to more en- example, spacecraft that are used for prolonged vironmentally-friendly alternatives. Unfortunate- on-orbit periods (such as Orion) require robust ly, the materials developed by other government MMOD protection. Structurally integrated TPS agencies are often incapable of meeting the more improves not only MMOD damage tolerance but stringent spacecraft requirements. Adhesive ma- also could provide significant weight savings as a terials, currently used for cryoinsulation applica- load-bearing structure. This technology area in- tions, are mercury-based and in need of replace- cludes TPS that is designed to be robust in the ment to limit hazardous waste and ease processing MMOD environment. Other multi-function- restrictions. Based on lessons learned from the al TPS, such as materials that also improve radia- Ares I Upper Stage development work, addition- tion protection, or combine cryogenic insulation al activities have been linked to this task to in- with ascent/entry heating protection may also be crease effectiveness and reduce overall risk. These included. The integration of TEGs into a TPS activities include development of a small-scale holds the potential to harvest free energy and save compounding/blending facility to directly sup- weight by reducing onboard power systems. Sim- port new material research and characterization, ilarly, the development of TPS that could harvest construction of a TPS tooling and cold storage fa- ascent or plume heating for storage and later use cility, and development of advanced TPS testing (in space or on a planetary surface) would prove techniques and capabilities to support the overall beneficial to several potential missions.. Develop- development effort. The time horizon for matur- ment of these technologies is estimated to be a 5 - ing these technologies is estimated to be 1-6 years 10 year effort due to the low TRL level (2-3) with depending on which material focus and support- the TEG integrated TPS having a higher TRL lev- ing activities are chosen. el (4-5) and probably requiring less time to ma- 2.3.2. Plume Shielding (Convective and ture after recent successes. Radiative) 2.3.1.7. Obsolescence-Driven TPS Materials 2.3.2.1. Plume Shielding and Process Development Protection of the spacecraft from the convec- This technology effort would ultimately provide tive and radiative heating components of rocket TPS materials and processes that are directly ap- engine and thruster plumes is required. It is be- plicable to heavy lift launch vehicles, commercial lieved that the Comet Nucleus Tour (Contour) vehicles, space-based cryogenic propellant depots, spacecraft was lost due to the impingement of hot and specialized ground-based test equipment re- gases from an engine burn. Clearly, the physics of quiring thermal insulation. This effort contin- plume heating and impingement need to be bet- ues development of replacement cryoinsulation, ter understood and techniques to analyze and de- primer, adhesive, and ablator TPS materials that sign protection for the vehicle and its components are currently facing obsolescence. These four class- from this heating are required. . This technology es of materials are each subject to unique obso- area includes high-temperature insulation blan- lescence issues that will limit their availability for kets and coatings. Development time estimated to future programs. The current generation of cryo- be 3 years. insulation materials are due to be phased out of production by 2015 and require replacement with more environmentally-compliant systems that provide no reduction in performance capabilities.

TA14-22 DRAFT 2.3.3. Sensor Systems and Measurement ulation, and Information Technology Systems) Technologies and TA12 (Materials, Structural & Mechanical 2.3.3.1. Sensor Systems and Measurement Systems, and Manufacturing). Our technologies Technologies are then required as technology push or through collaborative development for the remaining elev- Flight safety would be greatly improved with en systems-based TA’s. Specifically, for TA1 and advances in ultra light-weight TPS sensor systems TA13, advances in cryogenic tank and feedline measuring temperature, strain, recession, flux, joints insulation, helium conservation and/or and other quantities of interest. This would be elimination through the use of aerogels instead of achieved by providing data needed for on-orbit/in active purge systems provide a significant impact situ, or self-repairing mechanisms, or to adaptive to launch capability. Technologies developed in control algorithms that can compensate for dam- Cryogenic Systems will have direct impact on re- age without repair. Two sensor system technolo- fueling operations for TA4, superconducting tech- gies currently in development show great promise: nologies for TA5, consumables, production and FO, and wireless. Distributed FO sensor systems environmental protection for TA7, and cryocool- providing full-field data (strain, temperature) have ers required by TA8. In addition, technologies de- been extensively ground tested, and have recent- veloped in Thermal Control Systems will have a ly been successfully flight tested on aircraft, TRL direct impact on autonomous control and moni- 6-8. Current development efforts include redun- toring or robotic spacecraft developed by TA4, on dant pathway FO networks that allow fiber failure the habitation and life support systems developed without losing all downstream data (SOA TRL 3), within TA6. Finally, technologies developed un- ~5 years to TRL 6. Current SoA for high temper- der Thermal Protection Systems have direct link- ature RFID thermal sensors is TRL 2 for passive age to propulsion systems and instrumentation capability up to 500°C with TRL 6 expected with isolation and protection for TA9. 5 years development. Higher temperature capabil- Interdependencies between the Thermal Man- ity may be leveraged from current Air Force fund- agement Systems TA and other TA’s are highlight- ed research. Other wireless sensors may be pos- ed in Figure 6. sible and may benefit from energy scavenging or increased energy density technology development 4. Possible Benefits to efforts (see TA03). Other National Needs 3. Interdependency with The technologies identified in Section 2 will Other Technology Areas have spin-off potential to both the public and pri- vate sectors in applications or systems where con- The Thermal Management Systems TA does rely trol of the flow of heat is desired. A matrix rep- on collaboration and technology push from the resenting what sectors the Thermal Management TA10 (Nanotechnologies), TA11 (Modeling, Sim- Systems technologies would be applicable and the

Figure 6. Thermal Management Systems TA Interdependencies DRAFT TA14-23 estimated impact to that sector is provided in Fig- dients. Obvious benefits would be in the develop- ure 7. ment of commercial spacecraft as well as military Advances in space cryogenic thermal control sys- spacecraft and aircraft. Other applications may tems have direct relevance to industry on Earth. benefit reactors or other energy generation tech- Cryogenic processes are very energy intensive and nologies where high heat dissipation is present. greater system efficiencies will reduce power con- Additionally, any advancement in the multi-func- sumption and minimize product losses which will tional TPS could have strong spin-off potential. have a direct impact on the transportation, super- TEG technology could find broad application in conducting power transmission, biomedical, and furnace technology, automotive applications, re- remote sensing sectors. Technological advances actors, and may have clean energy applications, as in thermal control systems and in reducing ther- well. HMS applications are very broad and could mal conductivity will have a significant impact on be employed in commercial aircraft, the automo- ground-based heating and cooling industries that tive industry, and a variety of structures. Lastly, would result in significant reductions in power the continued development of aerogels and other consumption and operating costs. nano-structured materials would benefit, as they The development of new thermal protection already have, technologies centered on environ- and hot structure systems may find application mental protection or filtration systems. anywhere there are high heat loads or thermal gra-

Figure 7. Potential benefits of Thermal Management Systems technologies to other national needs

TA14-24 DRAFT Acronyms Acknowledgements AC Alternating Current The draft NASA technology area roadmaps were ADR Adiabatic Demagnetization Refrigeration developed with the support and guidance from AMPM Agency Mission Planning Manifest the Office of the Chief Technologist. In addition C/C Carbon/Carbon to the primary author’s, major contributors for the C/SiC Carbon/Silicon Carbide TA14 roadmap included: the OCT TA14 Road- CPL Capillary Pumped Loop mapping POCs, Tibor Balint and Tammy Gafka; DC Direct Current the NASA Center Chief Technologist and NASA DRA Design Reference Architecture Mission Directorate reviewers, and the follow- DRM Design Reference Mission ing individuals Ryan Stephan, Stan Bouslog, Eric ECLSS Environmental Control and Life Hurlbert, and the Multi-Center Cryogenic Fluid Support System Management Team. EDL Entry, Descent, and Landing EVA Extra Vehicular Activity FO Fiber Optic HEFT Human Exploration Framework Team HMS Health Monitoring System IRVE Inflatable Reentry Vehicle Experiment ISRU In Situ Resource Utilization ISS International Space Station JT Joule-Thompson JWST James Webb Space Telescope LEO Low Earth Orbit LHP Loop Heat Pipe MLI Multi Layer Insulation MMOD Micro Meteoroid and Orbital Debris NDE Nondestructive Evaluation PCM Phase Change Material RCC Reinforced Carbon-Carbon RFID Radio Frequency Identification RTV Room Temperature Vulcanizing elastomer SHReD Supplemental Heat Rejection Device SOA State of the Art SOMD Space Operations Mission Directorate TA Technology Area TABS Technology Area Breakdown Structure TASR Technology Area Strategic Roadmap TCS Thermal Control System TEC Thermal Electric Cooler TEG Thermo-Electric Generator TPS Thermal Protection Systems TRAD Tile Repair Ablator Dispenser TRL Technology Readiness Level UHTC Ultra High Temperature Ceramic ZBO Zero Boil Off

DRAFT TA14-25 November 2010

National Aeronautics and Space Administration

NASA Headquarters Washington, DC 20546 www.nasa.gov TA14-26 DRAFT