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Components for Extreme Environments

Description:

ScopeTitle:

Mechanisms for ExtremeEnvironments

ScopeDescription:

Proposals are sought for mechanisms and mechanicalsystems that can operate on the dusty surface of the Moon and Mars formonths to years. These systems will be exposed to the harsh extremeenvironments and will have little to no maintenance. These mechanisms inextreme environments must function in the presence of lunar regolith andcharged dust, micrometeoroids, plume ejecta, extreme temperaturevariations, high vacuum, cosmic rays and other high-energy ionizedparticles, plasma, solar ultraviolet (UV) and other electromagnetic (EM)ionizing radiation, static electricity charging, changing gravitationalconditions, and other electrically induced effects.

Proposals shouldbe focused on the following mechanisms and technologies that canfunction in these environments:

  • Sealing materials, fabrics, and flexible covers and technologiesthat can seal/protect mechanisms by preventing regolith intrusion andremain compliant and functional in the extreme Moon/Marsenvironments.
  • Dust-tolerant electrical connectors that can function with (ormitigate) light dust coating in the relevant Moon/Marsenvironments.
  • Moving components for dust protection (iris, hatch, covers,louvers, airlocks, closures, hinges, joints, trusses,etc.).

Successfulsolutions will have the following performance characteristics:

  • Operational for extended service of 10 to 100 months with limitedor no maintenance.
  • Linear and static joints will function and perform thedesigned actuation/motion/mate-demate cycles of 1,000 orhigher.
  • Mechanisms will function with minimal solid film or withoutlubrication.
  • Operational lifetimes on the order of hundreds of thousands ofcycles.
  • All mechanisms will function throughout lunar temperaturecycles between 127 °C (260 °F)and -173 °C (-280°F).
  • All mechanisms will function in the extreme cold of permanentlyshadowed regions (‑238 °C) (-396°F).
  • All mechanisms will function reliably with lunar regolith(simulant) coating on the exposed mechanismsurfaces.
  • All mechanisms will function in the high-vacuum lunar environmentof 10-9 Torr.
  • All mechanisms and materials will function in the lunarelectrostatic and radiation environment.

Expected TRL or TRL Range at completion of theProject: 2 to 6

Primary TechnologyTaxonomy:

  • Level 1 07Exploration DestinationSystems
  • Level 2 07.2 MissionInfrastructure, Sustainability, andSupportability

DesiredDeliverables of Phase I and PhaseII:

  • Research
  • Analysis
  • Prototype
  • Hardware
  • Software

DesiredDeliverables Description:

Research should be focused on solving one of the NASA technologyneeds listed above. Applications with direct infusion path to currentand future NASA projects/programs are sought.

Phase I Deliverables: Aproof-of-concept or breadboard demonstrating technical feasibility andoperability in a laboratory environment, and a report that includesanalytical and model simulations in a relevant environment to answercritical questions focused on functional performance of the mechanisms.In addition, the report shall include recommendations for brassboard orprototype development during Phase II that is directly applicable to acurrent or future NASA project/program.

Phase II Deliverables: Delivery of abrassboard or prototype with a goal of achieving TRL 5 or 6, andlaboratory testing demonstrating operability over the range of expectedenvironmental conditions. The prototype shall be designed to conform toa NASA project/program need and include a well-developed flightdemonstration and infusion plan. A report shall be written that includesfunctional, performance, analytical, and test results; and an evaluationof the technology’s maturity level (i.e., TRL) including therisk of proceeding with the development.

State of the Art and CriticalGaps:

Previoussolutions used in the Apollo program did not address the current need oflong-term usage. Terrestrial solutions often employ materials or methodsthat are incompatible with the Moon/Mars environment.

Critical Gaps:

Seals at rotary and linear joints arevery common for actuation in dusty environments. Most of these seals,however, use elastomers that would off-gas and become brittle in a lunarradiation environment and at lunar temperatures. Solutions are neededthat employ advanced materials, metallic seals, or nontraditionaltechniques that can operate in the lunar environment for an extendedperiod of time (months to years).

Operations on the lunar surface willinvolve the mating/demating of electrical, fluid, and cryogenicconnections. Dust on the surface of these connectors will impede theirproper function and lead to failures. Solutions are needed to developconnectors that can function in dusty Moon/Mars extremeenvironments.

Dust-protective enclosures, flexiblecovers, boots, hatches, and moving covers are needed to protect delicatemechanism components.

Relevance / ScienceTraceability:

Developing mechanisms for extreme environments willbe one of the biggest challenges for operation on the lunar surface forthe Artemis program.

References:

DustMitigation Gap Assessment Report, International Space ExplorationCoordination Group (ISECG): https://www.globalspaceexploration.org/wordpress/docs/Dust%20Mitigation%20Gap%20Assessment%20Report.pdf

ScopeTitle:

Freeze-Tolerant Radiators, HeatExchangers, and Water Containers

ScopeDescription:

Proposals are sought to develop freeze-tolerantradiators, heat exchangers, and water containers. The goal isto develop these components that can freeze and thaw without sufferingdamage or performance degradation on human-rated spacecraft onthe lunar surface. Current ground rules and assumptions (GRAs) for lunarpressurized habitats include:

  1. Single-phase nontoxic external andinternal active thermal control system (ATCS) coolant loops.
  2. Heat exchangers and deployableradiators operating at turbulent flow to remove and reject heat.
  3. Operate near the lunar south poleand survive the lunar nights (lasting up to 14 days), whereenvironmental temperatures can drop below the freezing point of heritageand candidate ATCS coolants (e.g., ammonia, water, Freon, HFE 7200) andas low as -213 °C (-351 °F).
  4. Total heat loads varying between 2and 15 kW, or 6,824 to 51,182 BTU/hr.

Based on these GRAs, the risk of lossof mission (LOM) due to rupturing radiator and heat exchanger coolanttubes because of freeze-thaw cycles is high, and the development offreeze-tolerant radiators and heat exchangers is necessary to reducethis risk and reduce heater power during Artemis missions.

Specifically, developments in radiatorsand heat exchangers are sought in these areas:

  • Lightweight, corrosion-resistant,freeze-tolerant metallic coolant tubes ranging from 0.127 to 3.81 cm(0.05 to 1.5 in.) inner diameter, 51 to 304 cm (20 to 120 in.) long, andoperating under turbulent flow conditions.
  • Lightweight, high-strength,corrosion-resistant, freeze-tolerant nonmetallic flexible coolant tubesranging from 0.127 to 3.81 cm (0.05 to 1.5 in.) inner diameter, 51 to304 cm (20 to 120 in.) long, and operating under turbulent flowconditions.
  • Radiators and exchangers withvariable thermal resistance that can temporarily eliminate or reduceheat rejection. Examples include, but are not limited to, low-power(less than 1 kW) devices that are capable of suctioning, temporarilystoring, then refilling the coolant to and from a radiator or heatexchanger and variable emissivity devices or materials (e.g., louvers,thermochromic and electrochromic coatings).

Developments in freeze-tolerant watercontainers are sought in these areas:

  • Develop flexible, freeze-tolerantwater containers that can survive the extremely cold environmentaltemperatures at unpressurized and pressurized conditions on the lunarsurface. Water recovered from in situ devices may be contained in bagsthat are subjected to an unpressurized environment on the lunar surfaceand will be exposed to temperatures from -213 to 127 °C (-351 to260 °F). The water containers may be brought inside apressurized habitat or rover at atmospheric conditions, then processedand treated to produce potable water for contingency use. Therefore, thecontainers need to withstand pressure and thermal cycles, prevent thewater from freezing while on the lunar surface, and be flexible so theycan shrink when empty to reduce volume and expand when full; full toempty container ratio >100:1 and maximum water mass of 250 kg(555 lbm).

Expected TRL or TRL Range at completion of theProject: 3 to 6

Primary TechnologyTaxonomy:

  • Level 1 14 ThermalManagement Systems
  • Level 2 14.2Thermal Control Components andSystems

DesiredDeliverables of Phase I and PhaseII:

  • Analysis
  • Prototype
  • Hardware

DesiredDeliverables Description:

Phase I Deliverables: A proof-of-concept or breadboarddemonstrating technical feasibility and operability in a laboratoryenvironment, and a report that includes analytical and model simulationsin a relevant environment and heat loads to answer critical questionsfocused on reducing the risk of freezing radiators or heat exchangers.In addition, the report shall include recommendations for brassboard orprototype development during Phase II.

Phase II Deliverables: Delivery of abrassboard or prototype with a goal of achieving TRL 5 or 6, andlaboratory testing demonstrating operability over the range of expectedenvironmental conditions. The prototype shall be designed to conform toa NASA project/program need and include a well-developed flightdemonstration and infusion plan. A report shall be written that includesfunctional, performance, analytical, and test results; and an evaluationof the technology’s maturity level (i.e., TRL) including therisk of proceeding with the development.

State of the Art and CriticalGaps:

State of the art (SOA) ATCSs on human-rated spacecraft like theApollo Service Module (SM) and International Space Station (ISS) usemechanically pumped, single-phase coolant to collect, transport, andreject heat, and the components that are most vulnerable to rupturingdue to freeze-thaw cycles are the radiators and heat exchangers becausethey are exposed to the environment.

The Apollo SMradiators were designed to partially stagnate, and only the coolanttubes, not the manifolds, in the ISS radiators were designed towithstand the high-pressure transients induced by freeze-thaw cycles.This required small-inner-diameter (0.18-cm, or 0.07-in.) metallic(Inconel or stainless steel) coolant tubes with thick walls (outerdiameter of 0.32 cm, or 0.125 in.), optimal spacing between tubes, andturbulent flow. Bigger inner diameters may be required for futureradiators to enhance hydraulic and thermal performance, but increasingthe outer diameter to enable freeze tolerance will increase mass andcounter thermal performance.

Similarly, theApollo SM and ISS heat exchangers used metallic coolant tubes with largeinner diameters (2.5 cm, or 1 in.) and thin walls to achieve high heattransfer coefficients, but increasing the outer diameter for freezetolerance will impact thermal performance. Inconel and stainless-steelcoolant tubes were used in these systems for their higher thermalconductivity, corrosion resistance, and strength for micrometeoroid andorbital debris (MMOD) protection but consequently limit freezeprotection.

Therefore,nonmetallic flexible coolant tubes that are corrosion resistant withhigh strength are also desired to enable freeze tolerance while meetingthermal and hydraulic requirements. There are no SOA ATCSs that can varythe thermal resistance of a radiator or heat exchanger to temporarilyeliminate or reduce heat rejection, but this capability is desired toenable freeze tolerance.

SOA contingencywater containers (CWCs) used on the space shuttle and the ISS weredesigned to be stored in an atmospheric environment and were not ratedfor the vacuum conditions, pressure cycles, and extreme environmentaltemperatures expected at the lunar south pole. Current containers have areasonable full water mass to empty volume ratio of 25:1, and theinternal space on the ISS and space shuttle constrained the maximumwater mass to 45 kg (99 lbm). Critical gaps are the flexible,freeze-tolerant water containers for unpressurized and pressurizedconditions at temperatures ranging from -213 to 127 °C (-351 to260 °F); full to empty container ratio >100:1; andmaximum water mass of 250 kg (555 lbm).

Relevance / ScienceTraceability:

Pressurized habitats or rovers stationed near the lunar south polefor future Artemis missions will be exposed to extremely coldenvironmental temperatures as low as -213 °C (-351 °F)during lunar nights (up to 14 days). These temperatures are below thefreezing point of heritage or candidate ATCS coolants (e.g., ammonia,water, Freon, HFE 7200). Preliminary analysis results of the conceptuallunar surface habitat ATCS architecture showed that significant heaterpower (up to 4 kW, or 13,648 BTU/hr) is required to prevent the coolantfrom freezing and maintain operations. Thus, freeze-tolerant radiatorsand heat exchangers are needed to reduce heater power, avoid rupturingthe coolant tubes, and reduce the risk of loss of mission (LOM).

NASA is developing in situ water retrievaltechnologies to excavate or drill into regolith-based water depositsfrom various regions on the lunar surface, then transport, store, andprocess into potable water, propellant, fuel cell reactants, and lifesupport consumables for Artemis missions.

References:

  • Babiak, S., Evans, B., Naville, D.,and Schunk, G., "Conceptual Thermal Control System Design for aLunar Surface Habitat," Thermal Fluids & AnalysisWorkshop (TFAWS), August 24-26, 2021.
  • Binns, D., and Hager, P.,"Thermal Design Challenges for Lunar ISRU Payloads,"50th International Conference on Environmental Systems (ICES), July12-15, 2021. 
  • Samonski, F.H., Jr., and Tucker,E.M., “Apollo Experience Report: Command and Service ModuleEnvironmental Control System,” NASA Technical Note (TN)D-6718, March 1, 1972.
  • “International SpaceStation (ISS) Active Thermal Control System (ATCS) Overview,”https://www.nasa.gov/pdf/473486main_iss_atcs_overview.pdf
  • Carter, L., et al.,“Status of ISS Water Management and Recovery,” 49thInternational Conference on Environmental Systems (ICES), July 7-11,2019.
  • Tobias, B., et al.,“International Space Station Water BalanceOperations,” 41st International Conference on EnvironmentalSystems (ICES), 2011.
  • Li, S., et al., “DirectEvidence of Surface Exposed Water Ice in the Lunar PolarRegions,” PNAS, 115, 2018, pp. 8907-8912, https://www.pnas.org/content/pnas/115/36/8907.full.pdf
  • Colaprete, A., Schultz, P.,Heldmann, J., Wooden, D., Shirley, M., Ennico, K., and Goldstein, D.,“Detection of Water in the LCROSS Ejecta Plume,”Science, 330 2010, pp. 463-468.
  • Schultz, P.H., Hermalyn, B.,Colaprete, A., Ennico, K., Shirley, M., and Marshall, W.S.,“The LCROSS Cratering Experiment,” Science, 330,2010, pp. 468-472.

Scope Title:

ActivelyControlled Louvers

ScopeDescription:

NASAplans to develop infrastructure to enable a sustaining human presence onthe Moon as part of the Artemis program. Current lunar orbit and surfacehabitat concepts incorporate conventional single-phase radiators toreject heat, and these habitats will be exposed to ionizingultraviolet (UV) radiation and lunar dust. The UV and lunar dustenvironments can significantly degrade the radiator’s Z-93absorptivity properties and reduce heat rejection capability. Inaddition, the radiator coolant tubes may rupture when exposed tosubfreezing environmental temperatures during transit to lunar orbit andat nighttime in lunar south pole regions. Radiator coating degradationand coolant freezing jeopardize the success of Artemis missions. Louvertechnology is a promising solution to maintain radiator performance andintegrity, but heritage louvers are passively controlled. Active-controllouvers are sought to improve thermal response times and allow groundcontrol. The louver design must be compliant with the current groundrules and assumptions (GRAs) as follows:

  • Maintain radiator heat rejectioncapability between 2 and 15 kW.
  • Minimum 15-year life.
  • Louver shall vary the effectiveradiator emissivity from 0.14 (blades closed) to 0.74 (bladesopen).
  • Louver blade thickness between 0.5and 2 cm.
  • Electromagnetic charging shall bemitigated.
  • Dust-tolerant design that mitigatesthe effects of a dusty environment.

Specifically, developments in louvers aresought in these areas:

  • Lightweight and corrosion-resistantmaterial; ideally less than 1 kg (2.2 lbs.) and compliant with NASASTD-6016A.
  • Thermal response time to setpoint changes toless than 15 min.
  • Electrically powered and dust-protectedactuation.
  • Actuation power less than 500 W.

Expected TRL or TRL Range at completion of theProject: 2 to 4

Primary TechnologyTaxonomy:

  • Level 1 14 ThermalManagement Systems
  • Level 2 14.2Thermal Control Components andSystems

DesiredDeliverables of Phase I and PhaseII:

  • Analysis
  • Prototype
  • Hardware

DesiredDeliverables Description:
Phase I Deliverables: A proof-of-conceptsmall-scale demonstration (no less than 1,400 cm2)showing technical feasibility and operability (i.e., thermalresponse time and emissivity range) in a laboratory environment, and areport that includes analytical results in a relevant environment andrecommendations for brassboard or prototype development during Phase II.In addition, the report should include a material trade study assessingthe louver weight against the emissivity range.

Phase II Deliverables: A brassboard orprototype representing a no-less-than ~7-m2 radiatorpanel with louvers in a vacuum environment. The goal is to achieve a TRLof 4 or 5. The testing should demonstrate operability over the range ofexpected environmental conditions and heat loads. A report shall bewritten that includes functional, performance, analytical, and testresults; an evaluation of the technology’s maturity level(i.e., TRL), including the risk of proceeding with the development; anda well-developed flight demonstration and infusion plan.

State of the Art and CriticalGaps:

State-of-the-art (SOA) louverblades are made from aluminum and are passively actuated using abimetallic spring. The louver blade transition from open to closed orvice versa and resulting thermal response time can take 1 to 2 hr. Thelouver thermal response time needs to be less than 15 min for humanlunar habitats. Studies have shown that 1 to 4 kW of heat power isneeded to keep coolant in a 48-m2 deployable radiatorfrom freezing. Passive louvers are not electrically powered, and activelouver power should be less than 500 W. A conventional 14-blade aluminumpassive louver weighs ~1 kg (2.2 lb). The active louver mass, includingthe control mechanism, needs to be less than 0.5 kg (1.1 lb).

Relevance / ScienceTraceability:

A lunarhabitat will be exposed to high-energy, or ionized, UV radiation whiletraveling through the Van Allen belts and can last from hours to days.Experiments have shown exposure to more than 500 equivalent sun hours(ESH) in the Van Allen belts can degrade the radiator’s Z-93absorptivity from 0.16 to 0.24, or 50%. An absorptivity reduction of 50%results in approximately 9 to 3 kW, or two-thirds reduction in heatrejection capability based on conservation of energy. Conventionalaluminum louver blades are approximately 1.3 cm thick, and theUV intensity through the Van Allen belts can be eliminated withthis thickness based on the Beer-Lambert law. Lunar dust is copious andhighly adhesive. Tests have shown Z-93 absorptivity linearly degradeswith the amount of dust coverage on the coating. As little as 20% dustcoverage can increase the absorptivity by 75% and decrease the heatrejection capability by 30%. Lunar habitats stationed near the lunarsouth pole will be exposed to extremely cold environmental temperatures(as low as -213 °C or -351°F) during lunar nights (up to14 days). The cold environmental temperatures are below the freezingpoint of heritage or candidate active thermal control system (ATCS)coolants (e.g., ammonia, water, Freon, HFE 7200). Conservation of energyanalysis results showed significant heater power (up to 4 kW, or 13,648BTU/hr) is required to prevent heritage coolants from freezing andmaintain operations. Louvers can reduce the radiator’seffective emissivity to 0.14 while in the closed position and keep theradiator outlet temperature above the HFE 7200 working and freezingpoints.

References:
Cowan, D., “Actively ControlledLouver for Human Spacecraft Radiator Ultraviolet (UV), Dust, and FreezeProtection,” International Conference of Environmental Systems(ICES), July 16-20, 2023.

Sawyer, D.M., and Vette, J.I.,“AP-8 Trapped Proton Environment for Solar Maximum and SolarMinimum,” NASA Technical Memorandum TM-X-72605, pg. 87,December 1976.

Edwards, D.L., and Zwiener, J.M.,“Radiation Induced Degradation of White Thermal Control PaintsZ-93 and Z-93P,” NASA Technical Memorandum 108518, October1996.

Gilmore, D., Spacecraft Thermal ControlHandbook, 2nd ed., The Aerospace Press, California, 2002, Chap.6.

Gaier, J.R., Siamidis, J., et al.,“The Effect of Simulated Lunar Dust on the Absorptivity,Emissivity, and Operating Temperature on AZ–93 and Ag/FEPThermal Control Surfaces,” NASA Technical Memorandum 215492,December 2008.

Sierra Nevada Corporation (SNC) ProductCatalog, 2015.

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