NASA SBIR 2022-I Program Solicitations

The Life Support and Habitation Systems Focus Area seeks key capabilities and technology needs encompassing a diverse set of engineering and scientific disciplines, all of which provide technology solutions that enable extended human presence in deep space and on planetary surfaces such as Moon and Mars, including Orion, ISS, Gateway, Artemis and Human Landing Systems. The focus is on systems and elements that directly support human missions and astronaut crews, such as Environmental Control and Life Support Systems (ECLSS), Extravehicular Activity (EVA) systems, Human Accommodations, including crew and cabin provisioning, hygiene and clothing systems, and Bioregenerative Life Support, including plant growth for food production.

For future crewed missions beyond low-Earth orbit (LEO) and into the solar system, regular resupply of consumables and emergency or quick-return options will not be feasible. New technologies must be compatible with attributes of the environments expected, including microgravity or partial gravity, varying atmospheric pressure and composition (both internal to the cabin and external to the vehicle), space radiation, and the presence of planetary dust. Technologies of interest are those that enable long-duration, safe, economical, and sustainable deep-space human exploration. Special emphasis is placed on developing technologies that will fill existing gaps as described in this solicitation, that reduce requirements for consumables and other resources, including mass, power, volume and crew time, and which will increase safety and reliability with respect to the state-of-the-art. Spacecraft may be untended by crew for long periods, therefore systems must be operable after these intervals of dormancy.

ECLSS encompass process technologies and monitoring functions necessary to provide and maintain a livable environment within the pressurized cabin of crewed spacecraft, including environmental monitoring, water recycling, waste management and atmosphere revitalization including particulate removal. There are two specific technical areas of interest for ECLSS submissions. Advancements in heaters and thermal swing components are needed for thermally desorbed carbon dioxide removal and compression beds, including considerations for structured monolithic sorbents created by additive manufacturing or slip casting of the sorbent itself. Secondly, proposals are sought to address challenges in carbon dioxide reduction systems, including separation, collection, removal and storage of carbon particulates, methods to recharge or recycle catalysts and solutions to prevent clogging of frits and filters in recycle gas streams. Also, of interest to ECLSS but included elsewhere in this solicitation, is lunar dust filtration and monitoring for spacecraft cabins.

For Human Accommodations, the focus in this solicitation includes advanced heating and refrigeration systems for stored food, personal hygiene including handwash, combination clothes washer and dryer systems and volumetrically efficient concepts for equipment, flexible work surfaces and stowage. In addition, textiles are sought for extreme surface environments and high oxygen atmospheres, applicable to crew clothing. Lastly, of interest to the focus area but included elsewhere, is the subtopic Plant Research Capabilities in Space, which is applicable to Bioregenerative Life Support.

Unique needs also exist for the Exploration Extra-vehicular Mobility Unit (xEMU), commonly called spacesuits. Textiles used for the xEMU Environmental Protection Garment (EPG), the outermost component of the xEMU, must resist extreme surface environments including planetary dust and also be suitable for oxygen-rich atmospheres. Applicable to the xEMU’s Portable Life Support System (PLSS), sorbent technologies are sought for a low volume, low power and low mass carbon dioxide and humidity control system. In addition, miniaturized gas sensor technologies are needed for measurement of oxygen, carbon dioxide and water vapor within the suit.

Please refer to the description and references of each subtopic for further detail to guide development of proposals within this technically diverse focus area.

Lead Center: MSFC

Participating Center(s): ARC, GRC, JPL, JSC, KSC

Scope Title: Advancements in Carbon Dioxide Reduction

Scope Description:

Air Revitalization Systems (ARS) are necessary for human survival during space exploration missions. Technologies to efficiently remove carbon dioxide (CO₂) from the cabin atmosphere and to reduce the captured CO₂ to recover oxygen are two systems that face technical challenges. Using adsorption beds to remove CO₂ is a proven technology, but optimization is needed. Please see the second scope "Advanced Heaters for Solid Sorption Systems" for more information. In the area of CO₂ reduction, several technologies produce solid carbon either intentionally or unintentionally. A current challenge to the development of these technologies is carbon management. Technologies and methods that will efficiently separate, remove, and store the carbon are sought. Technical solutions will allow for efficient operation of the carbon reduction process, prevent contamination of downstream hardware receiving effluent gases and avoid contamination of cabin atmosphere during carbon handling and disposal.

Oxygen recovery technology options, including carbon formation reactors and methane pyrolysis reactors, almost universally result in particulates in the form of solid carbon or solid hydrocarbons. Mitigation for these particulates will be essential to the success and maintainability of these systems during long-duration missions. Techniques and methods leading to compact, regenerable devices or components for removing, managing, and disposing of residual particulate matter within Environmental Control and Life Support Systems (ECLSS) process equipment are sought.

NASA has invested in many CO₂ reduction technologies over the years to increase the percentage of oxygen recovery from CO₂ in human spacecraft for long-duration missions. Examples of technologies include, but are not limited to, Series-Bosch, Continuous Bosch, methane pyrolysis, and microfluidic carbon dioxide electrolysis. Significant technical challenges still face these process technologies and are impeding progress in technology maturation. Critical technical elements of these technologies have a high degree of technical difficulty.

Examples where additional component technology development is needed include (this is a partial list):

Separation of particulate carbon from process gas streams.
Safe collection, removal, and disposal of solid carbon, including cases when continuously operating reactors are active.
Subsystems to recharge reactors with new catalyst and to efficiently reuse or recycle consumable catalysts.
Technology solutions to mitigate solid carbon clogging of frits and filters in recycle gas streams.

Separation performance approaching HEPA rating is desired for ultrafine particulate matter with minimal pressure drop. The separator function should be capable of operating for hours at high particle loading rates. If necessary, periodic operations and methods could be employed to restore capacity/functionality back to nearly 100% of its original clean state through in-place and autonomous regeneration or self-cleaning operations using minimal or no consumables (including media-free hydrodynamic separators). The device must minimize crew exposure to accumulated particulate matter and enable easy particulate matter disposal or chemical repurposing.

This subtopic is open to consider novel ideas that address any of the numerous technical challenges that face development of CO₂ reduction hardware with particular attention to solid carbon management.

Expected TRL or TRL Range at completion of the Project: 2 to 5

Primary Technology Taxonomy:
Level 1: TX 06 Human Health, Life Support, and Habitation Systems
Level 2: TX 06.1 Environmental Control & Life Support Systems (ECLSS) and Habitation Systems

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Hardware
Prototype

Desired Deliverables Description:

Phase I deliverables: Reports demonstrating proof of concept and test data from proof-of-concept studies, concepts, and designs for Phase II. Phase I tasks should answer critical questions focused on reducing development risk prior to entering Phase II. Conceptual solution in Phase I should look ahead to satisfying the requirement of limiting crew exposure to the raw carbon dust as well as carbon exposure to downstream hardware.

Phase II deliverables: Delivery of technologically mature hardware, including components and subsystems that demonstrate performance over the range of expected spacecraft conditions. Hardware should be evaluated through parametric testing prior to shipment. Reports should include design drawings, safety evaluation, and test data and analysis. Prototypes must be full scale unless physical verification in 1g is not possible. Robustness must be demonstrated with long-term operation and with periods of intermittent dormancy. The system should incorporate safety margins and design features to provide safe operation upon delivery to NASA.

State of the Art and Critical Gaps:

Advanced oxygen recovery systems are necessary for long-duration missions as resupply of consumables will not be available. The state-of-the-art Sabatier system, which has flown on the International Space Station (ISS) as the Carbon Dioxide Reduction Assembly (CRA), only recovers about half of the oxygen from metabolic CO₂. This is because there is insufficient hydrogen to react all available CO₂. The Sabatier reacts hydrogen with CO₂ to produce methane and water. The methane is vented overboard as a waste product causing a net loss of hydrogen. Mars missions target >75% oxygen recovery from CO₂, with a goal to approach 100% recovery. NASA is developing several alternate technologies that have the potential to increase the percentage of oxygen recovery from CO₂, toward fully closing the ARS loop. Methane pyrolysis recovers hydrogen from methane, making additional hydrogen available to react with CO₂. Other technologies under investigation process CO₂, recovering a higher percentage of oxygen than the Sabatier. All these alternative systems, however, need additional technology investment to reach a level of maturity necessary for consideration for use in a flight ECLSS.

Several of these alternative systems produce solid carbon either intentionally or unintentionally and solutions for safely filtering, removing, and storing solid carbon are critical to the maturation of these systems.

Relevance / Science Traceability:

These technologies would be essential and enabling to long-duration human exploration missions, in cases where closure of the atmosphere revitalization loop will trade over alternate ECLSS architectures. The atmosphere revitalization loop on the ISS is only about 50% closed when the Sabatier is operational. These technologies may be applicable to Gateway, lunar surface, and Mars, including surface and transit missions. This technology could be proven on the ISS as a flight demonstration.

This subtopic is directed at needs identified by the Life Support Systems Capability Leadership Team (CLT) in the area of atmosphere revitalization, and specifically, in the areas of CO₂ reduction and oxygen recovery, functional areas of ECLSS.

The Life Support Systems (LSS) Project, under the Advanced Exploration Systems (AES) Program, within the Human Exploration and Operations Mission Directorate (HEOMD), is the expected customer. The LSS Project would be in position to sponsor Phase III and technology infusion.

References:

"Hydrogen Recovery by Methane Pyrolysis to Elemental Carbon" (49th International Conference on Environmental Systems, ICES-2019-103)
"Evolving Maturation of the Series-Bosch System" (47th International Conference on Environmental Systems, ICES-2017-219)
"State of NASA Oxygen Recovery" (48th International Conference on Environmental Systems, ICES-2018-48)
"Particulate Filtration from Emissions of a Plasma Pyrolysis Assembly Reactor Using Regenerable Porous Metal Filters" (47th International Conference on Environmental Systems, ICES-2017-174)
"Methane Post-Processing and Hydrogen Separation for Spacecraft Oxygen Loop Closure" (47th International Conference on Environmental Systems, ICES-2017-182)
“Trading Advanced Oxygen Recovery Architectures and Technologies” (48th International Conference on Environmental Systems, ICES-2018-321)
NASA-STD-3001, VOLUME 2, REVISION A, Section 6.4.4.1 “For missions longer than 14 days, the system shall limit the concentration in the cabin atmosphere of particulate matter ranging from 0.5 μm to 10 μm (respirable fraction) in aerodynamic diameter to <1 mg/m3 and 10 μm to 100 μm to <3 mg/m3.” https://www.nasa.gov/sites/default/files/atoms/files/nasa-std-3001-vol-2a.pdf

Scope Title: Advanced Heaters for Sorbent Systems

Scope Description:

Spacecraft carbon dioxide (CO₂), water, and trace contaminant (organics) removal systems must be regenerable and reliable and minimize resupply and equivalent system mass (ESM). In most sorbent systems, heat is used to regenerate the beds by expelling contaminants for disposal or to downstream processes for resource recovery. In future deep space exploration missions, such as those to the Moon and to Mars, sorption systems must drastically reduce power to minimize the dependence on scarce resources. The state-of-the-art (SOA) spacecraft sorption systems utilize commercial off-the-shelf (COTS) resistive heaters coupled with conductive fins. These Joule heating methods lead to inefficiencies such as high thermal contact resistance, high temperature differential within the sorption beds, high component mass and volumes, and long ramp-up times. The SOA cooling options utilize blowers, cooling channels or cold plates in conjunction with spacecraft liquid cooling loops. Since the spacecraft cooling systems are limited in capacity, efficient cooling methods are needed. Although it is recognized that the conductivity of the sorbent material is the limiting factor to the heating of sorption beds, it is also important to design integrated thermal management systems that transfer the heat quickly, uniformly, and efficiently throughout the bed. Some suggested, but not inclusive, areas of heater improvements are listed here:

Decreasing the contact resistance between the heaters and the sorbent media.
Increasing temperature uniformity within the sorbent beds.
Improving tolerance to corrosion.
Optimizing for various configurations of sorbent media, including granules, beads, porous solids, additively manufactured, and liquid sorbents.

Some thermal management components can function both as heaters and coolers. This will lead to reduced system mass and volume of heaters, fin stock, cooling channels, various supporting hardware, and sorption materials. Proposed concepts may include different heater types as well as heater configurations. Heater configurations could include those that are bound or embedded into the sorbent media.

This subtopic solicits advanced thermal management systems that offer a significant improvement over the SOA. The heaters, coolers, configurations, and all attached hardware must meet the following operational requirements:

Continuous operation at temperatures as high as 200 ˚C or above.
Minimize both heating and cooling rates compared to the SOA heaters capability.
Heaters and cooling options must be able to operate in temperature swing sorption systems continuously 24 hours a day. Some example cycle times are those used in the current spacecraft system: the Carbon Dioxide Removal Assembly used a 144-minute cycle time; The 4BCO2 beds operate on 80-minute cycle times.
Compatible with either liquid or solid sorbent systems.
Capable of operating in microgravity and reduced gravity environment.
Must be compatible with sorbent regeneration or thermally sorbent systems.
Must be able to operate continuously for 3 years.
Offer an improvement in heat conservation, efficiency, power consumption, reliability, resupply, and ESM over the spacecraft SOA systems.
Heaters and cooling options must utilize the available power and cooling options expected in exploration spacecraft such as avionic air, the low-temperature loops, and the medium-temperature loops.
Meet the space station safety requirement.

Expected TRL or TRL Range at completion of the Project: 2 to 4

Primary Technology Taxonomy:
Level 1: TX 06 Human Health, Life Support, and Habitation Systems
Level 2: TX 06.1 Environmental Control & Life Support Systems (ECLSS) and Habitation Systems

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware

Desired Deliverables Description:

Phase I deliverables: Reports demonstrating proof of concept and test data from proof-of-concept studies, concepts, and designs for Phase II. Phase I tasks should answer critical questions focused on reducing development risk prior to entering Phase II. Phase I analysis should include a trade study between the advanced heaters and the SOA and operation in the spacecraft environments.

Phase II deliverables: Delivery of technologically mature hardware, including components and subsystems that demonstrate performance over the range of expected spacecraft conditions. Hardware should be evaluated through parametric testing prior to shipment. Reports should include design drawings, safety evaluation, test data, and analysis. Prototypes must be for sorption beds sized for 4 crew members. Robustness must be demonstrated with extended operation and with periods of intermittent dormancy. Systems should incorporate safety margins and design features to provide safe operation upon delivery to NASA.

State of the Art and Critical Gaps:

Current and future human exploration missions require regenerable systems that minimize mass, power, volume, and resupply and are highly reliable. Most SOA sorption systems in the Atmosphere Revitalization System (ARS) use COTS heaters that are inefficient, leading to high power requirements. Thermal management in systems such as the Carbon Dioxide Removal Assembly and the Sabatier could be improved by using advanced heating systems. Unfortunately, innovative heaters such as heat pipes and vapor chambers have been used elsewhere in space hardware but have yet to be developed for use in Environmental Control and Life Support Systems (ECLSS).

In addition, a significant amount of the spacecraft power is allocated to a variety of ECLSS. Alternative thermal management approaches that have multiple functions such as heating, cooling, thermal energy storage, and the thermal energy transfer over long distances will drastically reduce the loading on available resources for both in-transit and planetary base missions. These advanced heaters can be used for other NASA mission architectures as well, such as the extravehicular activity (EVA) and the Trash Compaction Processing System.

Relevance / Science Traceability:

This subtopic is relevant to Human Exploration and Operations Mission Directorate (HEOMD), especially ECLSS, by improving thermal management systems to minimize loading on facility resources such as power, heater, and cooling systems. In addition, efficient heaters minimize mass, power, and volume. The following ECLSS systems could benefit from improvements in thermal management technology: the ARSs, the Water Management Systems, and Solid Waste Management Systems including trash compaction. Other technical areas that may have interest are small satellites and EVA.

References:

Cmarik, Gregory, James Knox, and Warren Peters. "4-Bed CO2 Scrubber–From Design to Build." 2020 International Conference on Environmental Systems, 2020.
Peterson, G. P., and H. B. Ma. "Theoretical analysis of the maximum heat transport in triangular grooves: a study of idealized micro heat pipes." (1996): 731-739.
Schunk, Richard, Warren Peters, and John Thomas. "Four Bed Molecular Sieve–Exploration (4BMS-X) Virtual Heater Design and Optimization." 47th International Conference on Environmental Systems, 2017.
Wang, G., D. Mishkinis, and D. Nikanpour, "Capillary heat loop technology: space applications and recent Canadian activities." Applied thermal engineering, 2008. 28(4): p. 284-303.

Tra-My Justine Richardson and Darrell Jan. "A Trade-off Study of the Spacecraft Carbon Dioxide Management System using the Analytical Hierarchy Process", 48th International Conference on Environmental Systems, ICES-2018-332

Lead Center: JSC

Participating Center(s): JPL

Scope Title: Human Accommodations for Exploration Missions

Scope Description:

Humans have been living and working in Low Earth Orbit (LEO) for several decades; however, human accommodations such as galley and hygiene facilities are still fairly limited. As mission length and distance increase, these comforts of home will become even more important, and their resource footprint will need to be reduced. Missions to the Moon and Mars will introduce partial gravity where optimal design of human accommodations may be different than in LEO. Additionally, emerging commercial activities in LEO and the lunar vicinity will create a larger demand for human accommodations in space.

Innovative technologies that improve human accommodations over the state of the art are sought in the areas of galley, personal hygiene, laundry, and volumetrically efficient use of space for tasks.

Expected TRL or TRL Range at completion of the Project: 3 to 6

Primary Technology Taxonomy:
Level 1: TX 06 Human Health, Life Support, and Habitation Systems
Level 2: TX 06.1 Environmental Control & Life Support Systems (ECLSS) and Habitation Systems

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware

Desired Deliverables Description:

Phase I deliverables: Reports demonstrating proof of concept, test data from proof-of-concept studies, and designs leading to Phase II. Phase I tasks should answer critical questions focused on reducing development risk prior to entering Phase II. Conceptual solutions should clearly describe resource requirements such as hardware mass, volume, and power, as well as water use and crew time to operate.

Phase II deliverables: Delivery of technologically mature hardware, including components and subsystems that demonstrate performance over the range of expected spacecraft conditions. Hardware should be evaluated through parametric testing prior to shipment. Reports should include design drawings, safety evaluation, and test data and analysis. Prototypes should be full scale unless physical verification in 1g is not possible. Robustness must be demonstrated with long-term operation and with periods of intermittent dormancy. System should incorporate safety margins and design features to provide safe operation upon delivery to NASA.

State of the Art and Critical Gaps:

State of the art for most human accommodations is defined by International Space Station (ISS) hardware. For sleep and privacy, crew quarters consist of permanent rack-size compartments that accommodate sleeping bag, privacy, personal wall space, ventilation with limited temperature control, lighting, and personal entertainment. The ISS galley consists of a table, food warmer, and potable water dispenser, which can rehydrate food and drinks with hot or ambient temperature water. A small refrigerator has also been added for food and drink storage. Personal hygiene is accomplished with disposable wipes, wetted towels, no-rinse shampoo, Earth-like oral and hair care and normal clothing that is discarded when it gets too dirty. Housekeeping relies mainly on disposable disinfectant wipes and a vacuum cleaner. On ISS, there is no cooking, sink (handwash), shower, dishwasher, washing machine, or dryer.

Critical gaps include:

Rapid food heating for 4 crewmembers at the same time so that crews can dine together. Ideally, heating of 16 food packages could be accomplished in 30 to 45 minutes with less than 500 W of electricity. Food must be heated in accordance with NASA Standard 3001 and the Human Integration Design Handbook, and all equipment must meet touch temperature limits.
Food refrigeration for long-term storage on the way to and from Mars. Stored food volumes of 2 to 8 m³, with average packaged food density of 388 kg/m³, may be required at temperature ranges of -25 to 5 °C. Concepts must be volumetrically efficient, mass efficient, and highly reliable since loss of food quality can result in loss of crew performance. Secondary mass penalty for cold stowage should be below 0.2 kg per 1 kg of food. The refrigeration and insulation systems should be efficient enough to run (at steady state) on less than 0.15 W/kg of food frozen at -22 °C in a 23 °C ambient.
Personal hygiene with less consumables is needed. Currently 0.11 kg/person/day of wet wipes are supplied, and the goal is to reduce this below 0.05 kg/person/day.
Water efficient handwash for use in microgravity environment. Soap, water, and crew interface aspects must all be considered.
Clothes washer/dryer combination for use on the Moon (1/6g) or Mars (1/3g) that can clean up to 4.5 kg of cotton, polyester, and wool clothing at a time in less than 7 hours using <50 kg machine mass, <0.3 m³ external machine volume and <300 W electrical power (Note: 101.3 kPa habitat pressure may be assumed for prototype development).
Devices and systems for volumetrically efficient use of habitable volume in spacecraft. This may include random access stowage concepts where equipment and stowage could be packed together densely and slid open for random access when needed. Such a concept could optimize volumes according to real-time crew needs, while maximizing volume for stowage and equipment. Flexible work surfaces will also be considered. For example, systems that allow the crew to maximize "wall" and "ceiling" as work surfaces in a microgravity environment but allow reconfiguration if the habitat transitions into a gravity environment (i.e., walls and ceilings are less useful, but fold-out table tops or overhead features may deploy on demand). Logistics-2-Living concepts are also of interest, such as secondary stowage structure repurposing for the real-time creation of partitions, furniture, glove-boxes, etc.

Out of Scope:
Proposals are not solicited for toilets nor hardware considered life support systems, including air revitalization, water processing, or waste processing. Lunar dust mitigation technologies are covered elsewhere, but innovative interior surface cleaning including dust removal may be submitted to this subtopic. Crew quarters, exercise devices, and electronic devices for entertainment are not in scope here.

Relevance / Science Traceability:

The Logistics Reduction (LR) Project, under the Advanced Exploration Systems (AES) Program, within the Human Exploration and Operations Mission Directorate (HEOMD), is the expected initial customer. The LR Project will consider sponsoring Phase III Small Business Innovation Research (SBIR) activities and assist with technology infusion into NASA Moon-to-Mars missions.

References:

"Life Support Baseline Values and Assumptions Document", NASA/TP-2015–218570/Rev. 1
" NASA Spaceflight Human-System Standard, Volume 2: Human Factors, Habitability, and Environmental Health", NASA-STD-3001 Vol. 2, https://www.nasa.gov/hhp/standards
"Human Integration Design Handbook, Revision 1", https://www.nasa.gov/feature/human-integration-design/
"Dual Use of Packaging on the Moon: Logistics-2-Living", AIAA-2010-6049
"Lessons Learned for the International Space Station Potable Water Dispenser", ICES-2018-114
"Will Astronauts Wash Clothes on the Way to Mars?", ICES-2015-53

Lead Center: JSC

Participating Center(s): N/A

Scope Title: Spacesuit Gas Sensors

Scope Description:

As the design for the new Exploration Extravehicular Mobility Unit (xEMU) is developed, technology gaps have been identified for the gas sensors employed in the portable life support system (PLSS). These gaps need to be fulfilled to meet the new exploration requirements.

In order to ensure the safe operation of the spacesuit there is a need to measure the following major constituents in the gas stream across a total pressure range of 3.5 to 23.5 psia and temperature range of 35 to 125 °F: O₂ = 20 to 100%; CO₂ = 0 to 30 torr over 3.5 to 23.5 psia; H₂O = 5 to 90% relative humidity (RH). During ground testing these measurements can be made by ancillary equipment, however, the current design of the PLSS only includes nondispersive infrared (NDIR) sensors for CO₂. For reference, the outer mold line for these sensors is approximately 2.3 by 2.2 by 6.1 inches.

Since these sensors are continuously powered during an extravehicular activity (EVA) their power consumption is a direct driver of spacesuit battery capacity and, in consequence, spacesuit mass. It is, therefore, desirable to have a sensor power consumption below 2.5 W. The current CO₂ sensors consume 2 W during operation.

The intended use case for these sensors in the PLSS is to provide general situational awareness of the major constituents, in contrast to highly accurate measurements. The required accuracy of the sensors is therefore 1% or better for O₂ concentration and RH and 0.3 torr for CO₂ partial pressure.

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:
Level 1: TX 06 Human Health, Life Support, and Habitation Systems
Level 2: TX 06.2 Extravehicular Activity Systems

Desired Deliverables of Phase I and Phase II:

Prototype

Desired Deliverables Description:

Phase I products: By the end of Phase I, it would be beneficial to have a concept design for infusion into the xEMU. Testing of the concept is desired at this Phase.

Phase II products: By the end of Phase II, a prototype ready for system-level testing in the xEMU or in a representative loop of the PLSS is desired.

State of the Art and Critical Gaps:

As the design for the new xEMU is developed, there are obvious gaps in technologies that need to be fulfilled to meet the new exploration requirements. The currently employed gas sensors are functionally limited, draw significant power, and require new, innovative ideas. This solicitation is an attempt to seek new technologies for low-power multi-gas sensors. NASA has plans to go to the Moon and as the mission extends further out of low Earth orbit, the additional information provided by such sensors will be indispensable for the situational awareness of astronauts in space, as well as flight controllers on the ground.

Relevance / Science Traceability:

It is relevant to the new xEMU, International Space Station (ISS), as well as commercial space companies. As the xEMU is being designed, built, integrated, and tested at Johnson Space Center, solutions will have a direct infusion path as the xEMU is matured to meet the design and performance goals.

References:

https://www.nasa.gov/image-feature/exploration-extravehicular-mobility-unit-xemu

Lead Center: JSC

Participating Center(s): N/A

Scope Title: Spacesuit CO2 and Humidity Control Technology

Scope Description:

This solicitation is seeking to identify sorbent candidates that will compete with or outperform the current baseline sorbent technology used within the Exploration Extravehicular Mobility Unit (xEMU) for carbon dioxide (CO₂) and humidity control. It is desired that sorbent candidates meet or exceed the characteristics and goals listed below.

Key goals for sorbent characteristics and performance:

600- to 1,000-µm-sized beads.
Vacuum desorb technology (desorb at a pressure of 140 Pa).
CO₂ loading uptake (noncyclic) 25 °C, 8 mmHg CO₂, 10 °C dewpoint, 6.0 g CO₂/100 g sorbent.
H₂O loading uptake (noncyclic) 25 °C, 15 °C dewpoint, 7.0 g H₂O/100 g sorbent.
Uptake (cyclic) 25 °C, 8 mmHg CO₂, 10 °C dewpoint, 2.0 g CO₂/100 g sorbent at 2 to 3 minute half-cycle timing (e.g., adsorb for 2 minutes/desorb for 2 minutes).

In order to ensure the safe operation of the xEMU, CO₂ and humidity levels need to be controlled to levels in accordance with requirements established by the NASA medical community. The technology currently baselined for the xEMU is the Rapid Cycle Amine (RCA) technology and information on the RCA is also available in the reference section below. New technology alternatives to the RCA are desired in order to have a robust suit program that is able to fall back on alternate technologies if the need arises.

For the majority of an extravehicular activity (EVA), the CO₂ partial pressure required at the breathing gas inlet to the helmet of the spacesuit needs to be maintained at or below 2.2 mmHg when the astronaut is generating 2.44 g/min of CO₂. The flow rate of the oxygen ventilation loop that circulates the breathing gas from the suit, through the CO₂ and humidity removal unit and back to the helmet of the suit is maintained at 6 ft³/min.

The driving humidity requirement is to maintain the relative humidity of the breathing gas flowing into the helmet between 5 and 45% with the water vapor production level of 0.2 lb/hr and 6 ft³/min ventilation flow rate through the suit. The CO₂ and humidity control unit should also be able to handle situations where the generation rates are higher during shorter periods as described in the detailed requirements listed in the references section.

The goals for the mass of alternate technology units to be less than 14 lb, the volume to be less than 0.4 ft³, and the power consumption to be less than 1.4 W on average.

This subtopic is relevant to the xEMU, International Space Station (ISS), Gateway, and human landing system (HLS), as well as other endeavors currently in development by commercial space companies. The goal is to have proposed solutions to be designed, built, integrated and tested at Johnson Space Center and integrated into the xEMU. These solutions have the potential for a direct infusion path as the xEMU is matured to meet the design and performance goals.

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:
Level 1: TX 06 Human Health, Life Support, and Habitation Systems
Level 2: TX 06.2 Extravehicular Activity Systems

Desired Deliverables of Phase I and Phase II:

Prototype

Desired Deliverables Description:

Phase I products: By the end of Phase I, it would be beneficial to have candidate sorbent(s) identified that meet the goals listed for this solicitation. Testing of sorbent candidate is required at this Phase.

Phase II products: By the end of Phase II, testing of sorbent in the xEMU equivalent application and conditions is desired. Vendors may collaborate with research institutes if desired.

State of the Art and Critical Gaps:

The current state-of-the-art utilized on the ISS EMU is a metal oxide technology that requires astronauts to remove the unit from the PLSS, regenerate it in an oven, and reinstall it into the PLSS prior to the subsequent EVA.

The baseline xEMU technology provides regenerative CO₂ and humidity removal via a pressure swing adsorption system with a high-capacity sorbent that desorbs upon exposure to vacuum and requires little to no maintenance by the astronaut. This technology is well developed, but unparalleled. Ultimately, this solicitation is an attempt to lead to an alternate CO₂ and humidity removal system with regenerable capabilities requiring minimal astronaut maintenance to provide options for the new xEMU should unforeseen issues arise with the current technology. Additionally, xEMU has goals of reducing power draw, volume envelope, and mass while maintaining the current CO₂ and humidity removal capacity at the conditions described previously.

Relevance / Science Traceability:

It is relevant to the new xEMU, ISS, as well as commercial space companies. As the xEMU is being designed, built, integrated, and tested at Johnson Space Center, solutions will have a direct infusion path as the xEMU is matured to meet the design and performance goals.

References:

ICES-2016-073 Design and Development Comparison of RCA 1.0, 2.0, and 3.0 (Design and Development Comparison of Rapid Cycle Amine 1.0, 2.0, and 0 (tdl.org))
ICES-2019-400 RCA Testing History (Rapid Cycle Amine Testing History (tdl.org))

https://www.nasa.gov/image-feature/exploration-extravehicular-mobility-unit-xemu

Lead Center: LaRC

Participating Center(s): GRC

Scope Title: Lunar Surface 50-kW-Class Solar Array Structures

Scope Description:

NASA intends to land near the lunar South Pole (at S latitudes ranging from 85° to 90°) by 2024 in Phase 1 of the Artemis Program, and then to establish a sustainable long-term presence by 2028 in Phase 2. At exactly the lunar South Pole (90 S), the Sun elevation angle varies between -1.5° and 1.5° during the year. At 85 S latitude, the elevation angle variation increases to between -6.5° and 6.5°. These persistently shallow sun grazing angles result in the interior of many polar craters never receiving sunlight while some nearby elevated ridges and plateaus receive sunlight up to 100% of the time in the summer and up to about 70% of the time in the winter. For this reason, these elevated sites are promising locations for human exploration and settlement because they avoid the 354-hr nights found elsewhere on the Moon while providing nearly continuous sunlight for site illumination, moderate temperatures, and solar power [Refs. 1-2].

Under a recently announced “Game Changing” project in NASA’s Space Technology Mission Directorate (STMD) named Vertical Solar Array Technology (VSAT), several firms are developing relocatable 10-kW vertical solar arrays for initial modular power generation at the lunar South Pole [Refs. 3-4]. These adaptable 10-kW arrays can be retracted and moved as needed to support evolving requirements for initial South Pole human occupation. Their relatively small size (35 m² of deployed area) allows them to be used individually or in combination to power loads up to a few tens of kilowatts. However, because the Sun is always near the horizon at lunar polar sites, using numerous small interconnected arrays for electrical power loads >>10 kW can result in excessive shadowing of one array onto another as well as considerable positioning, leveling, and deployment challenges when locating them at optimally illuminated locations.

This subtopic seeks structural and mechanical innovations for relocatable 50-kW-class (40- to 60-kW) lightweight solar arrays near the lunar South Pole for powering second-generation lunar base infrastructure including habitats and laboratories, rechargeable rovers, and in situ resource utilization (ISRU) mining and processing machines, and that can deploy and retract at least 5 times. Increasing the unit solar array size from first-generation 10 kW to second-generation ~50 kW is a logical course of action as power needs increase for new higher-power capabilities such as ISRU or the Foundation Surface Habitat, which can require >>10 kW of power. This increase in size by 5 times while maximizing specific power (>75 W/kg) needs structures and mechanisms innovations and development effort to ensure compact packaging, safe transportation in space and on the lunar surface, reliable deployment, stable operation while sun tracking, and retraction and relocation as needed. Small Business Innovation Research (SBIR) contracts provide important near-term investment to flesh out specific technical requirements and new technical challenges for these larger 50-kW-class solar arrays based on VSAT results for smaller 10-kW arrays and on assumed Design, Development, Test, and Evaluation (DDT&E) schedules.

These 50-kW-class solar arrays are listed in NASA’s HEOMD-405 Integrated Exploration Capabilities Gap List as tier 1 (highest impact) development gap #03-04 for which at least 1 potential solution has been identified, but additional work is required to ensure feasibility of the new and/or novel performance or function in a specific operational application [Ref. 5]. The largest similar lightweight solar array under development is the 30-kW “ROSA” wing for NASA’s Lunar Gateway, but it is considerably smaller than desired for second-generation lunar surface arrays, and it is not designed to retract or to survive the unique lunar gravity, insolation, and dust and terrain environments. Exploration Capabilities Gap #03-04 is described as “Medium-power solar array technology for human-rated missions with specific power (>75 W/kg) and operation in mission specific environment.”

Retraction will allow valuable solar array hardware to be relocated, repurposed, or refurbished and possibly also to minimize nearby rocket plume loads and dust accumulation. Also, innovations to raise the bottom of the solar array by up to 10 m above the surface to reduce shadowing from local terrain are required [Ref. 6]. The ability to be relocated is assumed to be through use of a separate surface-mobility system (i.e., not necessarily part of the solar array system), but design of array structures and mechanisms should accommodate loads likely to be encountered during transport along the lunar surface. Suitable innovations, variations, or combinations of existing 10-kW array components to these much larger 40- to 60-kW arrays including those being developed under the VSAT project are of special interest.

Design guidelines for these deployable/retractable solar arrays are:

Deployed area: 140 m² (40 kW) initially; up to 210 m² (60 kW) eventually per unit, assuming state-of-the-art space solar cells.
Single-axis Sun tracking about the vertical axis.
Up to 10-m height extension boom to reduce shadowing from local terrain.
Deployable, stable base for supporting tall vertical array on unprepared lunar surface.
Base must accommodate a local 15° terrain slope with adjustable leveling to <0.5° of vertical.
Retractable for relocating, repurposing, or refurbishing.
Number of deploy/retract cycles in service: >5; stretch goal >10.
Lunar dust, radiation, and temperature resistant components.
Specific mass: >75 W/kg and specific packing volume: >20 kW/m³, including all mechanical and electrical components.
Factor of safety of 1.5 on all components.
Lifetime: >10 years.

Suggested areas of innovation include:

Novel packaging, deployment, retraction, and modularity concepts.
Novel lightweight, compact components including booms, ribs, solar cell blankets, and mechanisms.
Novel actuators for telescoping solar arrays such as gear/rack, piezoelectric, ratcheting, or rubber-wheel drive devices.
Mechanisms with exceptionally high resistance to lunar dust.
Load-limiting devices to avoid damage during deployment, retraction, and solar tracking.
Methodology for stabilizing large vertical arrays such as compactly packageable support bases, using regolith as ballast mass, or novel guy wire and surface anchor systems.
Optimized use of advanced lightweight materials, including composite materials with ultra-high modulus (>280 GPa) combined with low coefficient of thermal expansion (<0.1 m/m/°C).
Integration of novel structural health monitoring technologies.
Validated modeling, analysis, and simulation techniques.
Modular and adaptable solar array concepts for multiple lunar surface use cases.
Completely new concepts: e.g., thinned rigid panel or 3D-printed solar arrays, nonrotating telescoping “chimney” arrays, or lightweight reflectors to redirect sunlight onto solar arrays or into dark craters.

Proposals should emphasize structural and mechanical innovations, not photovoltaics, electrical, or energy storage innovations, although a complete solar array systems analysis is encouraged. If solar concentrators are proposed, strong arguments must be developed to justify why this approach is better from technical, cost, and risk points of view over unconcentrated planar solar arrays. Solar array concepts should be compatible with state-of-the-art solar cell technologies with documented environmental degradation properties. Design, build, and test of scaled flight hardware or functioning lab models to validate proposed innovations is of high interest.

Expected TRL or TRL Range at completion of the Project: 4 to 5

Primary Technology Taxonomy:
Level 1: TX 12 Materials, Structures, Mechanical Systems, and Manufacturing
Level 2: TX 12.2 Structures

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

In Phase I, contractors should prove the feasibility of proposed innovations using suitable analyses and tests. In Phase II, significant hardware or software capabilities that can be tested at NASA should be developed to advance their TRL. TRLs at the end of Phase II of 4 or higher are desired.

State of the Art and Critical Gaps:

This subtopic addresses capability gap #03-04 in the 2021 HEOMD-405 Integrated Exploration Capabilities Gap List titled “50 kW class solar power generation systems.” Gap #03-04 is one of just three tier 1 (highest impact) capability gaps in the 03) Aerospace Power and Energy Storage category, and is considered to be a development gap for which at least one potential solution has been identified but additional work is required to ensure feasibility of the new and/or novel performance or function in a specific operational application.

Deployable solar arrays power almost all spacecraft, but they primarily consist of hinged, rigid panels. This traditional design is too heavy and packages too inefficiently for lunar surface power. Furthermore, there is usually no reason to retract the arrays in space, so self-retractable solar array concepts are unavailable except for rare exceptions such as the special-purpose International Space Station (ISS) solar array wings. In recent years, several lightweight solar array concepts have been developed but none of them have motorized retraction capability either. The critical technology gap filled by this subtopic is a lightweight, vertically deployed, retractable 50-kW-class (40- to 60-kW) solar array for surface electrical power near the lunar South Pole for diverse needs including ISRU, lunar bases, dedicated power landers, and rovers.

Relevance / Science Traceability:

Robust, lightweight, redeployable solar arrays for lunar surface applications are a topic of great current interest to NASA on its path back to the Moon. New this year, the subtopic extends the focus area from human landers to other powered elements of the lunar surface architecture along with refined design guidelines. There are likely several infusion paths into ongoing and future lunar surface programs, both within NASA and also with commercial entities currently exploring options for a variety of lunar surface missions. Given the focus on the lunar South Pole, NASA will need vertically deployed and retractable solar arrays that generate 10 to 20 kW of power for first-generation capabilities and 40 to 60 kW for second-generation capabilities.

References:

Burke, J., “Merits of a Lunar Pole Base Location,” in Lunar Bases and Space Activities of the 21st Century, Mendell, W. (editor), 1985, https://www.lpi.usra.edu/publications/books/lunar_bases/
Fincannon, J., “Characterization of Lunar Polar Illumination From a Power System Perspective,” NASA TM-2008-215186, May 2008, https://ntrs.nasa.gov/citations/20080045536.
NASA Space Tech News, “NASA, Industry to Mature Vertical Solar Array Technologies for Lunar Surface,” March 23, 2021, https://www.nasa.gov/feature/nasa-industry-to-mature-vertical-solar-array-technologies-for-lunar-surface.
Pappa, R. S., et al., “Relocatable 10 kW Solar Array for Lunar South Pole Missions,” NASA-TM-20210011743, March 2021, https://ntrs.nasa.gov/citations/20210011743.
NASA Human Exploration & Operations, Systems Engineering and Integration, HEOMD-405 Version 1, “2021 Integrated Exploration Capabilities Gap List,” March 19, 2021.

Mazarico, E. et al., “Illumination Conditions of the Lunar Polar Regions Using LOLA Topography,” Icarus, February 2011, https://doi.org/10.1016/j.icarus.2010.10.030.

Lead Center: MSFC

Participating Center(s): JSC, LaRC

Scope Title: Inflatable Softgoods for Next Generation Habitation Systems: Testing and Structural Health Monitoring

Scope Description:

A key enabling technology for future crewed habitation systems is the development of inflatable softgoods materials and structures. In the past, habitat structures have typically consisted of metal alloys, but larger habitable volumes with lower structural mass will be required for long-duration, exploration-class missions. This subtopic seeks activities to mature inflatable softgoods through integration of sensing capabilities for structural health monitoring (SHM) and development of accelerated testing techniques.

Activities that may be undertaken under a Phase I effort include:

Development of approaches for accelerated materials creep testing, specifically for high-strength materials in inflatable softgoods such as webbings and cords. In implementing these materials in habitation structures, one long-term risk is failure of the structural material due to creep (deformation under sustained loading). Real-time creep testing at the component and subscale levels can take years and new test methods need to be developed to help certify softgoods for flight in their intended use environment. Approaches may include novel test methods to reduce the duration of the test (while still generating meaningful data relative to long term use of a material system in its environment) and/or a combination of test methods and modeling approaches to accelerate generation and capture of relevant lifetime material data. Development of new approaches will help to increase testing throughput and mitigate potentially catastrophic risks in failure of inflatable materials.

Scope of work includes:

Develop a methodology and test approach that can be validated to produce accurate predicted lifetime creep strain data and time-to-failure (TTF) for high-strength softgoods over a range of percent creep loads (50~90% nominally) within a year. (Within 1 to 3 months would be preferable).
- "High strength" refers to webbings and cordage of strength nominally in the range 5,000 to 20,000 lbs/in.
Produce master creep curves for each percent load and failure points.

The second focus area is integrated SHM. Integrated sensing capabilities in inflatable softgoods material systems are needed to monitor the structural performance of the material in situ, measure load/strain on softgoods components, detect damage, and predict further degradation/potential failures. The ability to acquire, process, and make use of this data in real time is an important risk mitigation for potential structural failure modes. The current state of the art in this field are instrumentation systems such as high-resolution strain gauges, fiber optics, accelerometers, and acoustic sensors using flexible electronics. However, there is a technology gap in developing a proven system that can integrate into a softgoods material system and continually monitor performance through its life. For this activity, integration of a system as a proof of concept and preliminary testing on an integrated inflatable softgoods structure are expected as part of the Phase I.

Scope of work includes:

Develop robust/repeatable integration approach with high-strength softgoods during manufacture or afterwards, that minimizes impact on the performance of the softgoods.
Develop sensor(s) that are robust to packaging/deployment/handling of the softgoods they are integrated into.
Repeatable performance and high-accuracy strain measurement (creep strain is typically 0.1 ~ 0.5%) once deployed and over the lifespan of the inflatable module.
Sensor(s) are inherently able to (or have a defined path to) survive the extreme environment of space.

Expected TRL or TRL Range at completion of the Project: 3 to 4

Primary Technology Taxonomy:
Level 1: TX 12 Materials, Structures, Mechanical Systems, and Manufacturing
Level 2: TX 12.2 Structures

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware

Desired Deliverables Description:

Phase I: Depending on the activity the proposer chooses to focus on, a Phase I effort would result in:

Development of accelerated creep test methods for inflatable softgoods at the component level and a laboratory proof of concept.
Approach to SHM for inflatable softgoods, a laboratory proof of concept of efficacy of approach, and/or preliminary design, which integrates the SHM approach into test articles.

Phase II: Depending on the activity the proposer chooses to focus on, a Phase II effort would result in:

Implementation of accelerated testing methods for evaluation of material creep and comparison with traditional (real-time) testing approaches at the component level.
Integration of SHM approach into inflatable softgoods components or subscale inflatable softgoods prototype and preliminary testing under load.

State of the Art and Critical Gaps:

Development of approaches for accelerated materials creep testing. Current state of the art for testing uses straps for real-time creep testing at the component level and subscale (or full-scale) inflatable softgoods test articles for (a) burst and (b) creep-to-burst testing. These tests are needed to understand the behavior of the inflatable softgoods over the mission lifetime and predict failure due to creep, which represents a catastrophic risk. Real-time testing takes months to years to collect data (depending on load level) and predictions require extrapolation from a limited number of data points. Accelerated testing techniques would enable higher fidelity characterization of the performance of the inflatable softgoods system over the entire mission scenario prior to flight and reduce risk.

Integrated SHM. Approaches for SHM in inflatable softgoods are needed to track the performance of the material system in real-time and identify when the structure has incurred damage or is at risk of failure. SHM typically uses strain gauges, digital image correlation, or accelerometers. SHM for inflatable softgoods requires novel approaches, as the material system is multilayer and fundamentally different from many other habitat structures. New techniques, such as flexile electronics, wireless systems, and fiber optics, are also generally unproven in a flight scenario for SHM.

Relevance / Science Traceability:

Technology for inflatable softgoods has historically been developed under Human Exploration Operations Mission Directorate (HEOMD) and Advanced Exploration Systems (AES). Current work on inflatable softgoods is under NASA's Next Space Technologies for Exploration Partnerships (NextSTEP) A: Habitation Systems Broad Agency Announcement opportunity, which has been ongoing since 2016 and focuses on design of next-generation habitat systems for cislunar space, the lunar surface, and Mars transit scenarios. The work under this subtopic will strongly complement ongoing work under the NextSTEP habitat project and increase the potential for the infusion of inflatable softgoods into future habitation concepts by reducing risk associated with understanding and predicting material behavior.

References:

Littiken, Doug. "Inflatable Habitat Inspection Needs." In-Space Inspection Workshop. 2017 31 January. https://www.asnt.org/~/media/Files/Events-Meetings/Conferences/ISIW/2017/3a_4_Litteken_NASA-JSC_Inflatable-Habitat-Inspection-Needs
Valle, Gerard, Doug Littiken, and Tom Jones. "Review of Habitable Softgoods Inflatable Design, Analysis, Testing and Potential Space Applications." AIAA SciTech. January 2019.
Littiken, Doug. "Inflatable technology: using flexible materials to make large structures." Proceedings of Electroactive Polymer Actuators and Devices. 13 March 2019. https://www.spiedigitallibrary.org/conference-proceedings-of-spie/10966/1096603/Inflatable-technology-using-flexible-materials-to-make-large-structures/10.1117/12.2500091.full?SSO=1

Edgecombe, John, Horacio de la Fuente, and Gerard Valle. "Damage Tolerance Testing of a NASA TransHab Derivative Woven Inflatable Module." 50th AIAA/ASMESCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference. 4-7 May 2009. https://arc.aiaa.org/doi/10.2514/6.2009-2167

Lead Center: MSFC

Participating Center(s): KSC, LaRC

Scope Title: Extraterrestrial Surface Construction

Scope Description:

Lunar and Martian construction of infrastructure from extraterrestrial materials and materials beneficiated or produced from in situ resources has the potential to radically reduce the cost and increase the scale of ambitious future space exploration. Technologies that support development of infrastructure structural elements are sought. Innovative materials and processes technology advancements are required to enable rapid advancement of a lunar or Martian village in a cost-effective manner.

Specific areas of technology development that are of interest include, but are not limited to, the following:

Construction technologies shall be based on the use of extraterrestrial materials and limit the need for any terrestrial materials. Development of lunar-construction-relevant materials and processes for infrastructure elements listed in point 2 below are highly encouraged.

Materials must have a defined application in a mission context.
Proposers are asked to define any consumable materials that must be brought from Earth for construction.

Fabrication and assembly of pressurized and unpressurized structural systems, including (for example) landing/launch pads, roads, blast shields, and habitats.

Both stationary and mobile fabrication/assembly systems shall be considered.
Novel fabrication and assembly methodologies shall be considered.
- Low-power methods and methods that benefit from the extraterrestrial surface environment are desired.

Technology development shall include design, analysis, fabrication, and testing of components, subsystems, and materials to enable full assessment and accountability of the technology product and fundamental findings with respect to their value toward reaching NASA's goals. Existing design and nondestructive evaluation (NDE) techniques are expected to be used when applicable. A relevant commercially available extraterrestrial simulant that mimics the silicate and oxide minerals in regolith and/or the volatiles in the lunar permanently shadowed regions or Martian surface and atmosphere is expected to be used for structure construction. Lunar materials, components, and systems that would be necessary for the proposed technology must be able to operate on the lunar surface (with thermal mitigations) in temperatures up to 110 °C (230 °F) during sunlit periods and as low as -170 °C (-274 °F) during periods of darkness. Martian materials, components, and systems must be able to operate on the Martian surface in a CO₂-rich atmosphere (with thermal mitigations, if necessary) in temperatures up to 20 °C (70 °F) and as low as -153°C (-225 °F). Systems must also be able to operate for at least 1 year with a goal of 5 years without substantial maintenance in the dusty regolith environment. Proposers should assume that operations involving other systems (e.g., robotics) and future astronauts will be ongoing not more than tens of meters away from the local fabrication and construction activities (i.e., minimization of dust generation is expected).

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:
Level 1: TX 12 Materials, Structures, Mechanical Systems, and Manufacturing
Level 2: TX 12.X Other Manufacturing, Materials, and Structures

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I deliverables may be a conceptual design with analysis to show feasibility at relevant scales and/or a small demonstration of the concept.

Phase II deliverables should be hardware demonstrations at a relevant scale.

State of the Art and Critical Gaps:

Planetary surface construction is not a current capability. The state of the art is terrestrial-based construction technology, e.g., cement, wood, and steel forms and terrestrial additive construction.

Relevance / Science Traceability:

The work desired applies to Technology Taxonomy (TX) Area 7: Exploration Destination Systems. It applies to 2018 NASA Strategic Plan Strategic Goal 2: Extend Human Presence Deeper into Space and to the Moon for Sustainable Long-Term Exploration and Utilization. It also applies to the Plan’s Strategic Objective 3.1: Develop and Transfer Revolutionary Technologies to Enable Exploration Capabilities for NASA and the Nation.

References:

Werkheiser, M. J., Fiske, M., Edmunson, J., & Khoshnevis, B. (2015). On the Development of Additive Construction Technologies for Application to Development of Lunar/Martian Surface Structures Using In-Situ Materials. In AIAA SPACE 2015 conference and exposition (p. 4451).
Moses, R. W., & Mueller, R. P. (2021). Requirements Development Framework for Lunar In Situ Surface Construction of Infrastructure. Earth and Space 2021 (pp. 1141-1155).
Gelino, N. J., Mueller, R. P., Moses, R. W., Mantovani, J. G., Metzger, P. T., Buckles, B. C., & Sibille, L. (2020). Off Earth Landing and Launch Pad Construction—A Critical Technology for Establishing a Long-Term Presence on Extraterrestrial Surfaces. Earth and Space 2021 (pp. 855-869).
Mueller, R. P., Fikes, J. C., Case, M. P., Khoshnevis, B., Fiske, M. R., Edmunson, J. E., ... & Andersen, C. (September 2017). Additive Construction with Mobile Emplacement (ACME). In 68th International Astronautical Congress (IAC), Adelaide, Australia (pp. 25-29).

Edmunson, J., Fiske, M., Alkhateb, H., Johnston, M., & Fikes, J. (2016). Additive Construction with Mobile Emplacement: Planetary.

Lead Center: LaRC

Participating Center(s): GSFC, MSFC

Scope Title: Nondestructive Evaluation (NDE) for In-Space and Additively Manufactured Materials/Structures

Scope Description:

NASA’s NDE SBIR subtopic will address a wide variety of NDE disciplines with a focus on in-space inspection. This SBIR solicitation will focus on aerospace structures and materials systems, including but not limited to Inconel, titanium, aluminum, carbon fiber, Avcoat, Alumina Enhanced Thermal Barrier (AETB), Phenolic Impregnated Carbon Ablator (PICA), and thermal blanket structures. Development efforts should target any set of these materials in common aerospace configurations, such as micrometeoroid and orbital debris (MMOD) shielding, truss structures, and stiffened structures. NDE can target material and material systems in a wrought state or additive manufacturing (AM). In-process or postproduction NDE techniques that could be used to inspect additively manufactured components will be favored. As NASA strives for longer duration space missions, these new tools need to be developed to support in-space manufacturing and assembly.

NDE Sensors and Data Analysis:

Technologies enabling the ability to perform automated inspections on large or complex structures are encouraged. Technologies should provide reliable rapid assessments of the location and extent of damage or defects. Methods are desired to perform inspections in areas with difficult access in pressurized habitable compartments and external environments for flight hardware. Many applications require the ability to see through assembled conductive and/or thermal insulating materials without contacting the surface.

Techniques that can dynamically and accurately determine position and orientation of the NDE sensor are needed to register NDE results to precise locations on the structure with little to no human intervention. Advanced processing and displays are needed to reduce the complexity of operation and interpretation for astronaut crews who need to make important assessments quickly. NDE inspection sensors are needed for potential use on free-flying inspection platforms. Integration of wireless systems with NDE may be of significant utility. It is strongly encouraged that proposals provide an explanation of how the proposed techniques and sensors will be applied to a complex structure. Examples of structural components include but are not limited to multiwall pressure vessels, batteries, tile, thermal blankets, micrometeoroid shielding, International Space Station (ISS) radiators, or aerospace structural components, including the lunar gateway.

Additionally, techniques for quantitative analysis of sensor data are desired. It is also considered highly desirable to develop tools for automating detection of material foreign object debris (FOD) such as lunar dust and/or defects and evaluation of bondline and in-depth integrity for ablative materials, like a heat shield. Typical internal void volume detection requirements for ablative materials are on the order of less than 6 mm, and bondline defect detection requirements are less than 25 mm.

Additive manufacturing is rapidly becoming a manufacturing method capable of producing fracture-critical components; as such, NDE requirements will become more stringent. Additively manufactured components represent a novel challenge for NDE due to the layering nature of the process and its effect on diffracting energy sources. Development of NDE techniques, sensors, and methods addressing these issues would be highly desired. Additionally, in situ inspection systems that support assessment of AM builds will be considered desirable. Most of the aerospace components will be metallic in nature, and critical flaws can be volumetric or fracture-like in nature.

In-Space Inspection:

Technologies sought under this SBIR include those related to in-space NDE. This includes on-orbit NDE (e.g., ISS or Gateway) as well as for future lunar, Mars, or other planetary missions. This could include new NDE tools for astronauts to use in a habitat or in the space environment (i.e., on an extravehicular activity (EVA)) or for automated inspection. Technologies may include fully functional NDE tools developed based on ground-use/laboratory equipment. Consideration will also be given to particularly promising technologies that may not provide turnkey operation but enable the advancement of future NDE inspection capabilities in space (i.e., enabling technologies). Fully functional NDE “tool” designs must address considerations related to size, mass, power, safety, environment, operation and/or automation, and data transfer related to their proposed application. For example, an NDE tool designed for ISS must ultimately be able to meet (after final development) ISS design requirements, launch mass/payload limitations, operational guidelines for crew, etc. If no specific application is outlined in the design, or if the proposal is for development of an enabling technology, then consideration must still be given to system size, mass, power, and data rate, to the extent that it makes the technology feasible in the within the next decade. To that end, consideration may be given to technology developments that are specifically focused on minimizing (or optimizing) these system parameters (e.g., low-mass, compact microfocus x-ray sources).

This solicitation is aimed at technologies for conventional NDE inspection of relevant components in space, meaning detection of commonly known defects in materials (cracks, pores, delamination, FOD, impact damage, etc.), rather than analytical tools aimed at determining chemistry, composition, or other properties of materials. Relevant components to be inspected may include (but are not limited to) spaceflight hardware, protective gear, core/rock samples, structural components, electronics/wiring, pressure vessels, thermal protection systems, etc. Of particular interest are technologies that advance the inspection of AM parts in space. These parts may be manufactured in an AM cabinet system that fits in an ISS EXPRESS (EXpedite the PRocessing of Experiments for Space Station) rack, which results in parts on the scale of 6 in. AM technologies used in such a payload could include fused deposition modeling, bound metal deposition, wire arc additive manufacturing, or other technologies using wire feedstock. Large-scale space structures may be manufactured or assembled in the space environment using AM techniques. Inspection technologies may involve x-ray technology (such as computed tomography), ultrasonic imaging, thermography, or any other NDE methods adapted for space use. NDE tools or enabling technologies that are compact, easy to carry (by astronauts), and work on low or accessible power will be considered.

Expected TRL or TRL Range at completion of the Project: 1 to 6

Primary Technology Taxonomy:
Level 1: TX 08 Sensors and Instruments
Level 2: TX 08.X Other Sensors and Instruments

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I Deliverables: For proposals focusing on NDE sensors: Lab prototype and feasibility study or software package, including applicable data or observation of a measurable phenomenon on which the prototype will be built. For proposals focusing on NDE modeling: Feasibility study, including demonstration simulations and data interpretation algorithms, proving the proposed approach to develop a given product (Technology Readiness Level (TRL) 2 to 4). Inclusion of a proposed approach to develop a given methodology to a TRL of 2 to 4. All Phase I proposals will include minimum of short description for Phase II prototype/software. It will be highly favorable to include a description of how the Phase II prototype or methodology will be applied to structures.

Phase II Deliverables: Working prototype or software of proposed product, along with full report of development, validation, and test results. Prototype or software of proposed product should be of TRL 5 to 6. Proposal should include plan of how to apply prototype or software on applicable structure or material system. Opportunities and plans should also be identified and summarized for potential commercialization.

State of the Art and Critical Gaps:

NASA and the SBIR program are preparing for the next phase of human deep space flight. As such, much of the materials, structures, and subsystem will have to be built or assembled in space. Quantitative and qualitative inspection of these components and structures will be critical to ensure safe spaceflight. Additionally, NDE sensors will be used to determine the health of structures as they age in space.

Relevance / Science Traceability:

Several missions could benefit from technology developed in the area of NDE. Currently, NASA is returning to manned spaceflight. The Artemis program's Orion spacecraft and Space Launch System have had inspection difficulties, and continued development and implementation of NDE tools will serve to keep our missions flying safely. Currently, Orion is using several techniques and prototypes that have been produced under the NDE SBIR topic. The Space Launch System is NASA’s next heavy-lift system, capable of sending hundreds of metric tons into orbit. Inspection of the various systems is ongoing and will continue to have challenges, such as verification of the friction stir weld on the fuel tanks. As NASA continues to push into deeper space, smart structures that are instrumented with structural health monitoring (SHM) systems can provide real-time mission-critical information on the status of the structure. NDE of spaceflight hardware and parts manufactured in space will be key enabling technologies for constant crew presence and long-duration missions.

References:

Burke, E. R.; Dehaven, S. L.; and Williams, P. A.: Device and Method of Scintillating Quantum Dots for Radiation Imaging. U.S. Patent 9,651,682, Issued May 16, 2017.
Burke, E. R.; and Waller, J.: NASA-ESA-JAXA Additive Manufacturing Trilateral Collaboration. Presented at Trilateral Safety and Mission Assurance Conference (TRISMAC), June 4-6, 2018, Kennedy Space Center, Florida.
Campbell Leckey, C. A.; Juarez, P. D.; Hernando Quintanilla, F.; and Yu, L.: Lessons from Ultrasonic NDE Model Development. Presented at 26th ASNT Research Symposium 2017, March 13-16, 2017, Jacksonville, Florida.
Campbell Leckey, C. A.: Material State Awareness: Options to Address Challenges with UT. Presented at World Federation of NDE Centers Short Course 2017, July 15-16, 2017, Provo, Utah.
Campbell Leckey, C. A.; Hernando Quintanilla, F.; and Cole, C.: Numerically Stable Finite Difference Simulation for Ultrasonic NDE in Anisotropic Composites. Presented at 44th Annual Review of Progress in Quantitative Nondestructive Evaluation, July 16-21, 2017, Provo, Utah.
Cramer, K. E.; and Klaassen, R.: Developments in Advanced Inspection Methods for Composites Under the NASA Advanced Composites Project. Presented at GE Monthly Seminar Series, April 13, 2017, Cincinnati, Ohio.
Cramer, K. E.; and Perey, D. F.: Development and Validation of NDE Standards for NASA's Advanced Composites Project. Presented at ASNT Annual Conference, October 30-November 2, 2017, Nashville, Tennessee.
Cramer, K. E.: Current and Future Needs and Research for Composite Materials NDE. Presented at SPIE Smart Structures and NDE 2018, March 4-8, 2018, Denver, Colorado.
Cramer, K. E.: Research Developments in Non-Invasive Measurement Systems for Aerospace Composite Structures at NASA. Presented at 2018 International Instrumentation and Measurement Technology Conference, May 14-18, 2018, Houston, Texas.
Dehaven, S. L.; Wincheski, R. A.; and Burke, E. R.: X-ray Transmission Through Microstructured Optical Fiber. Presented at QNDE - Review of Progress in Quantitative NDE, July 17-21, 2017, Provo, Utah.
Dehaven, S. L.; Wincheski, R. A.; and Burke, E. R.: X-ray Transmission Through Microstructured Optical Fiber. Presented at 45th Annual Review of Progress in Quantitative Nondestructive Evaluation (QNDE), July 15-19, 2018, Burlington, Vermont.
Frankforter, E.; Campbell Leckey, C. A.; and Schneck, W. C.: Finite Difference Simulation of Ultrasonic Waves for Complex Composite Laminates. Presented at QNDE 2018, July 15-19, 2018, Burlington, Vermont.
Gregory, E. D.; and Juarez, P. D.: In-situ Thermography of Automated Fiber Placement Parts: Review of Progress in Quantitative Nondestructive Evaluation. Presented at QNDE - Review of Progress in Quantitative NDE, July 17-21, 2017, Provo, Utah.

Gregory, E. D.; Campbell Leckey, C. A.; Schneck, W. C.; Swindell, P: A Versatile Simulation Framework for Elastodynamic Modeling of Structural Health Monitoring; https://ntrs.nasa.gov/citations/20190001865

Lead Center: LaRC

Participating Center(s): MSFC

Scope Title: Manufacturing of Materials from Lunar Surface Resources

Scope Description:

As humanity embarks on sustained deep space exploration, starting with the lunar surface, there will be a need for building infrastructure that is based on indigenous resources [1]. Usage of these resources will face limitations that include the available source materials, equipment, and power. Therefore, materials processing and manufacturing approaches are required that are operable within these constraints.

Operations on the lunar surface must consider types of materials available as well as their abundance. Various in situ resource utilization (ISRU) efforts are ongoing to extract and process the raw materials into usable forms. These include some SBIR topics that the prospective proposer is encouraged to investigate. Elements available for extraction from regolith include oxygen, silicon, iron, calcium, aluminum, magnesium, and titanium. From these, and from other materials that may be available in smaller quantities, manufacturing methods are needed to produce components for construction and for the building, replication, and repair of equipment.

Proposals are invited for approaches that utilize the resources available on the Moon to be able to produce structural girders, beams, and pipes that can withstand both tensile and bending forces. These are required in addition to compacted cementitious and sintered materials that can carry mostly compressive loads.

Concepts can include, but are not limited to, production using various metallic materials as well as basalt-fiber-reinforced geopolymers and other combinations that can be produced from lunar resources. Manufacturing methods that capitalize on the lunar environment are of particular interest.

The selection of the material system must consider the potential availability on the lunar surface and a demonstrated or projected ability to support tensile and bending loads. For example, proposed work may include an analysis of lunar material properties and processing methods that yield the required performance characteristics for relevant structures. An example beam would be a structural component for a crane with a 25-ft reach that can support one metric ton (2,200 lb) in lunar gravity. Proposers are pointed to the references provided [2-6] as well as ongoing ISRU activities for the latest and detailed information on the potential availability of various materials on the Moon.

Proposals to the current solicitation can assume the materials extracted and processed in the ISRU activities to be available and ready to use at levels of purity that range from as-dug regolith to separated and refined metals. The quantities available will depend on the lunar abundance of the materials and the effort needed for the processing. As-dug regolith can be expected to be available in large amounts; more refined materials can be expected to be available in quantities that decrease with the level of refinement and the requirements for that refinement, such as energy and any Earth-sourced ingredients.

Proposal elements of interest include but are not limited to:

Material concepts that can utilize various purities of feedstocks, e.g., concepts that might be able to use a metal at less than 100% purity.
Manufacturing processes that can take advantage of the lunar environment, such as vacuum, radiation, reduced gravity, etc.
Equipment required for the manufacturing, including the size scale, power requirements, production rates, and operating environments.
Preliminary proof-of-concept experiments for feasibility of the proposed material systems, processing methods, and equipment.

Expected TRL or TRL Range at completion of the Project: 4 to 5

Primary Technology Taxonomy:
Level 1: TX 12 Materials, Structures, Mechanical Systems, and Manufacturing
Level 2: TX 12.X Other Manufacturing, Materials, and Structures

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware

Desired Deliverables Description:

Phase I: Define the material system to be used for manufacturing of relevant components, the processes required, and the equipment needed to process that material. Provide one or more material systems, manufacturing processes, and equipment design concepts for the production of tensile- and bending-force-supporting components on the lunar surface using resources available from ISRU extraction and beneficiation activities. The concept will include analysis of how the material system(s) is/are able to meet the load-carrying requirements and the manufacturing parameters, and how the equipment that is required utilizes/succeeds in operating in the lunar environment.

Phase II would look at scaled/laboratory demonstrations of the material system(s), manufacturing processes, and equipment. These would include designing and building of relevant equipment and potential processing of commercially available regolith simulants or other materials that may match the materials expected to be available on the Moon, either in raw form or from other processes. Test coupons must be built and tested using as close an analog as possible of the lunar material system and a prototype of the proposed manufacturing equipment. Documentation of requirements for the manufacturing process and operation of the equipment, such as power and mass that can be used to evaluate feasibility in trade studies, shall be included.

State of the Art and Critical Gaps:

Sustainable long-term exploration of the Moon will be dependent on the utilization of lunar resources. While various efforts are looking at the excavation and extraction of those resources, there are currently gaps in manufacturing of the material feedstocks that may be available on the Moon into other useful products. Addressing these gaps requires understanding of the fabrication equipment and the full manufacturing cycle as well as the expected impact when the processes are run on the Moon.

Relevance / Science Traceability:

The Artemis program envisions the start of a long-term human presence on the lunar surface for the exploration and development of the Moon by Government as well as commercial companies and international partners. In order to support these missions, it will be essential to utilize resources that can be sourced from the lunar surface. The current solicitation calls for proposals that provide the support for these exploration and development activities. Technologies that are developed in this solicitation may also feature on preparatory missions for Artemis, such as the Commercial Lunar Payload Services Programs, depending on the readiness of the technology.

References:

NASA’s Plan for Sustained Lunar Exploration and Development. https://www.nasa.gov/sites/default/files/atoms/files/a_sustained_lunar_presence_nspc_report4220final.pdf [accessed 07/23/2021].
Grant H. Heiken, David T. Vaniman, Bevan M. French, eds. Lunar Sourcebook. Cambridge University Press, 1991. https://www.lpi.usra.edu/publications/books/lunar_sourcebook/ [accessed 07/23/2021].
R. D. Waldron. Lunar Manufacturing: A Survey of Products and Processes. Acta Astronautica. 1988; 17(7):691-708.
Dave Dietzler. Making It on the Moon: Bootstrapping Lunar Industry. NSS Space Settlement Journal, September 2016. https://space.nss.org/wp-content/uploads/NSS-JOURNAL-Bootstrapping-Lunar-Industry-2016.pdf [accessed 07/22/21].
I. A. Crawford. Lunar Resources: A Review. Progress in Physical Geography: Earth and Environment. 2015; 39(2):137-167.
United States Geological Survey (USGS). Unified Geologic Map of the Moon. https://astrogeology.usgs.gov/search/map/Moon/Geology/Unified_Geologic_Map_of_the_Moon_GIS_v2 [accessed 07/23/2021].

Scope Title: Welding Testbed for Space Manufacturing

Scope Description:

Technology development efforts are required to enable on-orbit servicing, assembly, and manufacturing (OSAM) for commercial satellites, robotic science, and human exploration. OSAM is an emerging national initiative to transform the way we design, build, and operate in space. The goal of the initiative is to develop a strategic framework to enable robotic servicing, repair, assembly, manufacturing, and inspection of space assets.

An in-space material welding capability is an important supporting technology for the long-duration, long-endurance space missions that NASA will undertake beyond the International Space Station (ISS). Historically, structures in space have been assembled using mechanical fastening techniques and modular assembly. Structural designs for crewed habitats, space telescopes, antennas, and solar array reflectors are primarily driven by launch considerations such as payload fairing dimensions and vibrational loads experienced during ascent. An in-space welding capability will greatly reduce constraints on the system imposed by launch, enabling the construction of larger, more complex, and more optimized structures. Welding is an essential complementary capability to large-scale additive manufacturing technologies being developed by NASA and commercial partners. Welding is also a critical capability for repair scenarios (e.g., repair of damage to a structure from micrometeorite impacts).

The development of welding processes for a variety of materials and thicknesses is carried out via a welding destructive testing and nondestructive testing feedback loop. This ensures that a weld procedure is well understood and that it produces welds that have sufficient material properties for their end-use application. While weld procedures are developed on the ground in simulated space environments, it is also necessary to further develop and validate these procedures in the true space environment where they will be applied. To achieve this need, a fully autonomous welding testbed must be created and deployed in space.

This subtopic seeks innovative engineering solutions—both fully autonomous and semiautonomous—to robotically weld materials for manufacturing in the unpressurized space environment. Current state-of-the-art (SOA) terrestrial welding methods such as laser beam, electron beam, and friction stir should be modified with an effort to reduce the footprint, mass, and power requirements for on-orbit applications.

Targeted applications for this technology include joining and repair of components at the subsystem level, habitat modules, trusses, solar arrays, and/or antenna reflectors. The need to repair a damaged structure or build new structures may require the need to not only weld material but to cut and remove material. A process that can weld material is the priority, but a robust process with cutting, removal, and testing capabilities adds value.

Expected TRL or TRL Range at completion of the Project: 3 to 6

Primary Technology Taxonomy:
Level 1: TX 12 Materials, Structures, Mechanical Systems, and Manufacturing
Level 2: TX 12.4 Manufacturing

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I is a feasibility study and laboratory proof of concept of a robotic welding process and system for on-orbit manufacturing applications. The Phase I effort should provide a laboratory demonstration of the welding process and its applicability to aerospace-grade metallic materials and/or thermoplastics, focusing on joint configurations that represent the priority in-space joining applications identified above. Work under Phase I will inform preliminary design of a mobile welding unit and/or in-space welding testbed. It will also inform a concept of operations for how the system would be deployed and operate in the space environment, with a focus on specific scenarios—for example, repair of a metal panel following micrometeorite damage, longitudinal welding of two metal curved panels, and welding of a truss to an adjacent truss. The Phase I effort should also provide an assessment of the proposed process operational capabilities (e.g., classes of materials that can be welded with the process, joint configurations that can be accommodated, and any expected impacts of the microgravity environment on joint efficiency relative to terrestrial system operation), volume, and power budget. A preliminary design and concept of operations are also deliverables under Phase I. Concepts for ancillary technologies such as postprocess inspection, in situ monitoring, mechanical testing, or robotic arms for manipulation of structures to be welded may also be included in the Phase I effort.

Development of a prototype with detailed analysis, initial testing, and associated software is desired for Phase I.

Phase II should further develop the prototype from Phase I and provide substantial test data using the prototype in an environment similar to the end application.

State of the Art and Critical Gaps:

A clear demonstrated understanding of the SOA is required. Any proposed technologies should not replicate the SOA and should instead advance the SOA or create an entirely different approach from the SOA. Welding in space has a multitude of applications, from repair to manufacturing, and is necessary to ensure a sustainable human presence in space. The development of space welding technologies is a substantial undertaking and requires years to perfect, so it is of the utmost importance that this process begins now. A welding testbed in space is an integral part of gaining weld property feedback data in an autonomous manner in a high-fidelity environment. The current SOA requires further advancement, and the growth of small business in the field of space welding is the best route to ensure that technological development is unique and that an array of technology providers exists in the future space economy.

Relevance / Science Traceability:

Space welding is necessary for the future sustainability of the space economy. To both build and repair structures in space, on the lunar surface, or on Mars, welding is a valuable tool that will provide agility for astronauts in a location where resources are highly limited. The development of space welding is a significant undertaking, so early development must begin now. The development of systems to autonomously weld structures in space and the ability to develop welding parameters through a closed-feedback-loop space testbed are both required to ensure that welding may be sufficiently applied in space.

References:

Tracie Prater et al. Overview of the In-Space Manufacturing Technology Portfolio. 2019. https://ntrs.nasa.gov/api/citations/20190030353/downloads/20190030353.pdf

Leigh M. Elrod et al. ISM In-Space Manufacturing. 2019. https://ntrs.nasa.gov/citations/20190033503

Lead Center: GRC

Participating Center(s): JSC, MSFC

Solicitation Year: 2022

Scope Title

Cryogenic Fluid Management (CFM)

Scope Description

This subtopic seeks technologies related to cryogenic propellant (e.g., hydrogen, oxygen, methane) storage and transfer to support NASA's space exploration goals. This includes a wide range of applications, scales, and environments consistent with future NASA missions. Such missions include but are not limited to upper stages, ascent and descent stages, refueling elements or aggregation stages, nuclear thermal propulsion, and in situ resource utilization. This subtopic solicits proposals in the following areas, in order of priority:

High-pressure-ratio compressor for on-orbit gas transfer: Design and develop concepts for a high-pressure-ratio compressor for supercritical xenon and helium, capable of increasing low-pressure fluid to 3,000 psia with continuous flow of up to 15 g/s, allowing for on-orbit transfer of gases for applications of refueling. The temperature range of the xenon is 17 to 40 °C. The compressor must be capable of surviving launch-load vibrations, be able to function accurately in microgravity and vacuum environment (10-5 torr), and be able to maintain gas cleanliness to Level A/10 for nonvolatile residue. For Phase I, the main deliverable should be a compressor design and performance analysis. For Phase II, the main deliverable should be a working engineering model of the compressor and the compressor itself.
Cryogenic flight-weight valves (minimum Cv >50, goal to Cv of ~100) for low-pressure (500 cycles with a goal of 5,000 cycles) to maximize the lifetime of the valve. Proposals can include metallic or nonmetallic sealing elements. Proposals should address the whole valve subsystem, including actuation and actuation mechanisms, with the goal of minimizing mass in Phase II. Phase I deliverable should be proof of concept of the valve with test data using liquid nitrogen, while the Phase II deliverable should be the valve.
Subgrid computational fluid dynamics (CFD) of the film condensation process for 1g and low gravity (lunar or martian) to be implemented into commercial industry standard CFD codes. The subgrid model should capture the formation and growth of the liquid layer as well as its movement along a wall boundary and should implement the volume of fluid (VOF) scheme. The condensation subgrid model should be validated against experimental data (with a target accuracy of 25%), with a preference for condensation data without a noncondensable. Emphasis should be placed on cryogenic fluid data, but noncryogenic data is acceptable. Phase I should be focused on simplified geometries (vertical plates/walls), while Phase II should be focused on complicated geometries (e.g., full cylindrical tank). The subgrid model and implementation scheme should be the final deliverable.
Development of heat flux sensors capable of measuring heat fluxes between 0.1 and 5.0 W/m2 for cryogenic applications. The sensors should have a target uncertainty of 2% full scale or less at temperatures as high as 300 K and at least as low as 77 K with a goal of 20 K. Proposers should target a demonstration of sensor operability in the 77-K temperature range in Phase I with a full demonstration of calibration and uncertainty in Phase II. Deliverable for Phase II should be the calibrated heat flux sensor.

Expected TRL or TRL Range at completion of the Project

2 to 4

Primary Technology Taxonomy

Level 1

TX 14 Thermal Management Systems

Level 2

TX 14.1 Cryogenic Systems

Desired Deliverables of Phase I and Phase II

Hardware Software Prototype

Desired Deliverables Description

Phase I proposals should at minimum deliver proof of the concept, including some sort of testing or physical demonstration, not just a paper study. Phase II proposals should provide component validation in a laboratory environment, preferably with hardware deliverable to NASA.

State of the Art and Critical Gaps

CFM is a crosscutting technology suite that supports multiple forms of propulsion systems (nuclear and chemical), including storage, transfer, and gauging, as well as liquefaction of ISRU-produced propellants. The Space Technology Mission Directorate (STMD) has identified that CFM technologies are vital to NASA's exploration plans for multiple architectures, whether hydrogen/oxygen or methane/oxygen systems, including chemical propulsion and nuclear thermal propulsion. Several recent Phase II projects have resulted from CFM subtopics, most notably for cryocoolers, liquid acquisition devices, phase separators, broad area cooling, and composite tanks.

Relevance / Science Traceability

STMD strives to provide the technologies that are needed to enable exploration of the solar system, both manned and unmanned systems; CFM is a key technology to enable exploration. Whether liquid oxygen/liquid hydrogen or liquid oxygen/liquid methane is chosen by Artemis as the main in-space propulsion element to transport humans, CFM will be required to store propellant for up to 5 years in various orbital environments. Transfer will also be required, whether to engines or other tanks (e.g., depot/aggregation), to enable the use of cryogenic propellants that have been stored. In conjunction with ISRU, oxygen will have to be produced, liquefied, and stored; liquefaction and storage are both CFM functions for the surface of the Moon or Mars. ISRU and CFM liquefaction drastically reduces the amount of mass that has to be landed.

References

No references for this subtopic.

Lead Center: JSC

Participating Center(s): ARC, GRC

Scope Title: Protective Medication Packaging Technologies Supporting Exploration Spaceflight Operations

Scope Description:

Successful long-duration space exploration missions will require robust crew support systems. These systems will rely on exponentially increasing crew autonomy, operate in low-to-no logistical resupply settings, and facilitate independent decision making within the context of challenging communication scenarios due to limited to no terrestrial-based support asset reach back. In addition, the long-duration spaceflight environment will require medically trained crew members who can assess, diagnose, and treat each other for a variety of illnesses and injuries. These medical events will require the preselection and long-term storage of various medications onboard human crewed spacecraft or pre-deployed in advance of human missions. Although currently there is no available method to sufficiently characterize or quantify the pharmaceutical stability, quality, or potency of repackaged medications (stored and eventually utilized for human consumption during long-duration space flight missions), available data shows that the median risk of drug failure (based on U.S. Pharmacopeia (USP) acceptance thresholds) for a 2-year exploration mission is approximately 59%. This risk increases to about 82% for a 3-year mission. These factors expose the distinct possibility that the provision of safe and effective drug treatment of long-duration crew may be at significant risk due to the current operationally derived need to repack crew medications to reduce resource "costs" (i.e., mass, volume, and power) possibly adversely impacting crew wellness, performance, and long-term health.

While baseline instability has not been experimentally investigated, most of the pharmaceuticals tested in spaceflight studies to date have been removed (due to mass, volume, and power considerations) from manufacturer's containers and repackaged into either polypropylene containers (Du et al. 2011) or lightweight, resealable plastic zipper storage bags. This type of repackaging remains the norm for supplying medications to the International Space Station (ISS). Unfortunately, such containers are not protective, therefore repackaged pharmaceuticals are exposed to ingress of atmospheric factors at concentrations in equilibrium with the ambient atmosphere (Putcha et al. 2016; Waterman et al. 2002). It is well established that such packaging is permeable to atmospheric factors such as moisture and oxygen and that prolonged exposure of susceptible medications is detrimental to shelf life (Roy et al. 2018; Waterman et al. 2002; Waterman et al. 2004).

Whereas exposure to spaceflight conditions (e.g., galactic cosmic radiation (GCR), microgravity or zero-gravity, etc.) is only a minor factor contributing to the cumulative risk of drug failure, with the significant factor being the baseline risk (observed in paired terrestrial controls under similar environmental conditions), repackaging of pharmaceuticals likely reduces medication effectiveness significantly (and increasingly, as "out of package" exposures extend in long-duration spaceflight), diminishes therapeutic effectiveness, thus potentially compromising crew health and performance.

In the past, repackaging methods have not been a significant limitation for missions where flight duration was much shorter than drug expiry (e.g., Apollo and Space Transport System (STS)) or where Low Earth Orbit permits regular replacement of expiring drugs (i.e., ISS). However, long-duration exposure of pharmaceuticals to atmospheric factors during exploration space missions will increase the risk of analytical drug failure over time, increasing the risk of therapeutic failure and potential exposure to toxicologically active impurities. Therefore, proven repackaging countermeasures are required to assure adequate stability of susceptible medications for the entire duration of exploration space missions.

This subtopic solicits proposals that address the critical need for exploring novel protective packaging technologies. Candidate technologies will retain or replicate (a) "initial" pharmaceutical packaging standards (i.e., minimization or elimination of atmospheric conditions), (b) acceptable shelf life (active pharmaceutical ingredient (API) minimums that meet or exceed Food and Drug Administration standards with respect to planned long-duration spaceflight timelines), (c) reduce reliance or need for cold storage/refrigeration of pharmaceuticals while, (d) preserving, optimizing, or reducing resource "costs" in regards to operational mass, volume, and power constraints (e.g., reducing power and mass requirements for an "in-vehicle" cold storage system), and (e) provide the potential for development of cross-cutting storage/repackaging technologies that integrate across, streamline, or expand the capabilities of multiple vehicle human support systems.

Expected TRL or TRL Range at completion of the Project: 3 to 6

Primary Technology Taxonomy:
Level 1: TX 06 Human Health, Life Support, and Habitation Systems
Level 2: TX 06.3 Human Health and Performance

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware
Software

Desired Deliverables Description:

Drug packaging that minimizes mass, volume, and material waste and protects contents from ingress of atmospheric factors, including moisture, oxygen, and carbon dioxide.

Drug packaging technologies that help preserve API integrity and efficacy across exploration spaceflight mission durations with minimal (or reduced) mass/volume/power resource cost(s).

Phase I Deliverable – Candidate packaging solutions.

Phase II Deliverable – Experimentally demonstrated effectiveness under long-term (2-year) and accelerated conditions.

State of the Art and Critical Gaps:

The state of the art of medication/pharmaceuticals packing technologies for exploration missions is uncertain. Foil packaging is an industry standard for pharmaceutical products and ensures low moisture transmission. Aclar® films have similar low moisture transmission and can be layered with other materials to increase the barrier to gas permeation. Mylar ® films have been used as a high-barrier packaging to protect foods and bulk pharmaceutical ingredients from the effects of oxygen, moisture, and light. Such materials—possibly combined with purging packaging headspace with inert gas (argon) or nitrogen—may be effective strategies to extend medication shelf life. Enclosing materials that scavenge oxygen, moisture (e.g., silica gels), and CO₂ may offer additional advantages.

Relevance / Science Traceability:

This subtopic seeks technology development that benefits the Exploration Medical Capability Element (ExMC) of the NASA Human Research Program (HRP). Pharmaceutical repackaging technologies are needed to address the following assigned risks:

Risk of ineffective or toxic medications during long-duration exploration spaceflight.
Risk of adverse health outcomes and decrements in performance due to inflight medical conditions.

This subtopic seeks technology development that supports the following identified HRP Gaps:

Pharm-101: "… determine the optimal packaging/storage strategy for medications in space that balances the needs of mitigating toxicity, preserving effectiveness, and minimizing resource "costs" (mass, volume, power, etc.)."
Pharm-401: “… perform further research to understand and characterize the active pharmaceutical ingredient and degradation profiles of medications for which we have low to moderate confidence in their safety and effectiveness for exploration missions.”
Pharm-601: “… characterize the extent to which spaceflight alters pharmacokinetics and pharmacodynamics.”

References:

Du B, Daniels V, Vaksman Z, Boyd J, Crady C, Putcha L (2011) Evaluation of Physical and Chemical Changes in Pharmaceuticals Flown on Space Missions. AAPS J 13:299-308. doi: 10.1208/s12248-011-9270-0.
Putcha L, Taylor PW, Daniels VR, Pool SL (2016) Clinical Pharmacology and Therapeutics. In: Anonymous Space Physiology and Medicine, pp 323-346. http://link.springer.com/10.1007/978-1-4939-6652-3_12.
Roy S, Siddique S, Majumder S, Mohammed Abdul MI, Ur Rahman SA, Lateef D, Dan S, Bose A (2018) A systemic approach on understanding the role of moisture in pharmaceutical product degradation and its prevention: challenges and perspectives 29. doi: 10.4066/biomedicalresearch.29-18-978.
Waterman KC, Adami RC, Hong JY (2004) Impurities in drug products. Sep Sci Technol 5:75-88. doi: https://doi.org/10.1016/S0149-6395(03)80006-5.
Waterman KC, Adami RC, Alsante KM, Antipas AS, Arenson DR, Carrier R, Hong J, Landis MS, Lombardo F, Shah JC, Shalaev E, Smith SW, Wang H (2002) Hydrolysis in Pharmaceutical Formulations 7:113-146. doi: 10.1081/PDT-120003494.
HRP Human Research Roadmap: Risks: Medication Stability Analysis: Device: https://humanresearchroadmap.nasa.gov/Tasks/task.aspx?i=1565

HRP Human Research Roadmap: Evidence Reports: https://humanresearchroadmap.nasa.gov/evidence/reports/Pharm.pdf?rnd=0.294621103114738

The Science Mission Directorate (SMD) will carry out the scientific exploration of our Earth, the planets, moons, comets, and asteroids of our solar system, and the universe beyond. SMD’s future direction will be moving away from exploratory missions (orbiters and flybys) into more detailed/specific exploration missions that are at or near the surface (landers, rovers, and sample returns) or at more optimal observation points in space. These future destinations will require new vantage points or would need to integrate or distribute capabilities across multiple assets. Future destinations will also be more challenging to get to, have more extreme environmental conditions and challenges once the spacecraft gets there, and may be a challenge to get a spacecraft or data back from. A major objective of the NASA science spacecraft and platform subsystems development efforts are to enable science measurement capabilities using smaller and lower cost spacecraft to meet multiple mission requirements thus making the best use of our limited resources. To accomplish this objective, NASA is seeking innovations to significantly improve spacecraft and platform subsystem capabilities while reducing the mass and cost that would in turn enable increased scientific return for future NASA missions. A spacecraft bus is made up of many subsystems such as: propulsion; thermal control; power and power distribution; attitude control; telemetry command and control; transmitters/antenna; computers/on-board processing/software; and structural elements. High performance space computing technologies are also included in this focus area. Science platforms of interest could include unmanned aerial vehicles, sounding rockets, or balloons that carry scientific instruments/payloads, to planetary ascent vehicles or Earth return vehicles that bring samples back to Earth for analysis. This topic area addresses the future needs in many of these sub-system areas, as well as their application to specific spacecraft and platform needs. For planetary missions, planetary protection requirements vary by planetary destination, and additional backward contamination requirements apply to hardware with the potential to return to Earth (e.g., as part of a sample return mission). Technologies intended for use at/around Mars, Europa (Jupiter), and Enceladus (Saturn) must be developed so as to ensure compliance with relevant planetary protection requirements. Constraints could include surface cleaning with alcohol or water, and/or sterilization treatments such as dry heat (approved specification in NPR 8020.12; exposure of hours at 115° C or higher, non-functioning); penetrating radiation (requirements not yet established); or vapor-phase hydrogen peroxide (specification pending). The National Academies’ Decadal Surveys for Astrophysics, Earth Science, Heliophysics, and Planetary Science discuss some of NASA’s science mission and technology needs and are available at https://www.nationalacademies.org/

Lead Center: JPL

Participating Center(s): GRC, GSFC, LaRC, MSFC

Scope Title: Critical Technologies for Sample-Return Missions

Scope Description:

This subtopic focuses on technologies for robotic sample-return (SR) missions that require landing on large bodies (e.g., the Moon, Mars, Vesta, Ceres, Phobos, Europa), as opposed to particulate-class SR missions (e.g., Genesis, Hayabusa) or touch-and-go (TAG) missions to relatively small asteroids or comets (e.g., OSIRIS-Rex, Hayabusa2). The mission destinations envisioned are dwarf planets (e.g., Vesta, Ceres) and planet or planet moons (e.g., Phobos, Europa). These are the most challenging missions in NASA's portfolio but also the most scientifically promising, given the vast array of instruments available on Earth to study the retrieved samples. Specifically, technologies are sought to address the following challenges associated with these SR missions: (1) Mass-efficient spacecraft architectures (e.g., efficient propulsion or materials that significantly reduce the mass of the launch payload required), (2) Sample integrity (e.g., surviving reentry), and (3) Planetary protection/contamination control (PP/CC) (e.g., preventing leakage into the Mars Sample Return (MSR) mission's orbital sample (OS) canister).

The heightened need for mass-efficient solutions in these SR missions stems from their extreme payload mass “gear ratio.” For example, the entire MSR campaign will probably require four heavy launch vehicle launches with rough spacecraft mass of 5,000 kg each to bring back multiple samples with an estimated total mass of 0.5 kg. Clearly, any mass savings in the ascent vehicle’s gross liftoff mass (GLOM) or in the mass of either the lander or the Earth Return Orbiter, for example, would yield many times more savings in the launch payload mass, enhancing the feasibility of these missions. Examples of propulsion technologies that may reduce overall mass include the development of lightweight, restartable ignition techniques for hybrid and solid rocket motors, lightweight spin motors, lightweight vectoring systems, lightweight insulation materials, and lightweight expandable nozzle designs to increase nozzle area ratios.

Once acquired, samples must be structurally and thermally preserved through safe landing and transport to Johnson Space Center (JSC) for analyses. Sample integrity technology solutions that address the long, high-radiation return trip, as well as the dynamic and high-temperature environment of reentry, are sought. Potential solutions include near-isotropic and crushable high-strength energy-absorbent materials that can withstand the ballistic impact landing. Materials that offer thermal isolation in addition to energy absorption are highly desirable given the reentry environment. In the case of cryogenically preserved samples, the technical challenge includes development of thermal control systems to ensure volatiles are conserved.

Finally, acquired samples must be chemically and biologically preserved in their original condition. Examples of PP/CC technology solutions sought include:

Materials selection: selection of metallic materials (non-organic) for the interior of the OS canister as well as materials that allow preferable surface treatments and bake-out sterilization approaches.
Surface science topics: Adsorber coatings/materials for contaminant adsorption (getter-type materials, such as aluminum oxide, porous polymer resin) and/or low-surface-energy materials to minimize contaminant deposition.
Characterization of contamination sources on lander, rover, capsule, ascent vehicle, and orbiter, for design of adequate mitigation measures.

Expected TRL or TRL Range at completion of the Project: 3 to 6

Primary Technology Taxonomy:
Level 1: TX 04 Robotics Systems
Level 2: TX 04.3 Manipulation

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype

Desired Deliverables Description:

A Phase I deliverable would be a final report that describes the requisite research and detailed design accomplished under the project.

A Phase II deliverable would be successful demonstration of an appropriate-TRL performance test, such as at representative scale and environment, along with all the supporting analysis, design, and hardware specifications.

State of the Art and Critical Gaps:

The kind of SR missions targeted in this solicitation are those that require landing on an extraterrestrial body. This most challenging kind of SR mission has only been successfully done in the Soviet Luna program that returned 326 g of Moon samples in three missions—out of eleven attempts—in the early 1970s. Hayabusa2 and OSIRIS-Rex are TAG SR missions. The former returned asteroid Ryugu samples to Earth in December 2020; the latter is expected to follow suit in September 2023 from asteroid Bennu. The first segment of NASA's MSR mission is the sample-collection rover Perseverance, which landed on Mars in February 2021. The MSR sample retrieval segment (lander, fetch rover, Mars Ascent Vehicle) is currently in Phase A development and expected to launch in 2028.

The content and breath of this solicitation is informed by lessons learned in MSR over the pre-Phase A years. Future SR missions are in need of technology improvements in each of the critical areas targeted: mass efficiency, sample integrity, and planetary protection.

This solicitation seeks proposals that have the potential to increase the TRL from 3 or 4 to 6 within 5 years, and are within the cost constraints of the Phases I, II, and III of this SBIR Program. Such progress would allow full flight qualification of the resulting hardware within 5 to 10 years.

Relevance / Science Traceability:

Medium- and large-class SR missions address fundamental science questions such as whether there is evidence of ancient life or prebiotic chemistry in the sampled body. Table S.1 of Vision and Voyages for Planetary Science in the Decade 2013-2022 (2011) correlates 10 "Priority Questions" drawn from three Crosscutting Science Themes, with "Missions in the Recommended Plan that Address Them." SR missions are shown to address 8 out of the 10 questions and cover every crosscutting theme, including Building New Worlds, Planetary Habitats, and Workings of Solar Systems.

References:

Vision and Voyages for Planetary Science in the Decade 2013-2022: http://nap.edu/13117
Visions into Voyages for Planetary Science in the Decade 2013-2022: A Midterm Review (2018): http://nap.edu/25186
Mars Sample Return (MSR): https://science.nasa.gov/science-pink/s3fs-public/atoms/files/07-GramlingMSR-PAC%2017Aug2020.pdf

Comet Nucleus Sample Return (CNSR): https://ntrs.nasa.gov/search.jsp?R=20180002990

Lead Center: JPL

Participating Center(s): GRC, GSFC, LaRC

Scope Title: Extreme Environments Technology

Scope Description:

This subtopic addresses NASA's need to develop technologies for producing space systems that can operate without environmental protection housing in the extreme environments of NASA missions. Key performance parameters of interest are survivability and operation under the following conditions:

Very low temperature environments (e.g., temperatures at the surfaces of Titan and of other ocean worlds as low as -180 °C; and in permanently shadowed craters on the Moon).
Combination of low-temperature and radiation environments (e.g., surface conditions at Europa of -180 °C with very high radiation).
Very high temperature, high pressure, and chemically corrosive environments (e.g., Venus surface conditions, having very high pressure and a temperature of 486 °C).

NASA is interested in expanding its ability to explore the deep atmospheres and surfaces of planets, asteroids, and comets through the use of long-lived (days or weeks) balloons and landers. Survivability in extreme high temperatures and high pressures is also required for deep-atmospheric probes to the giant planets. Proposals are sought for technologies that are suitable for remote-sensing applications at cryogenic temperatures and in situ atmospheric and surface explorations in the high-temperature, high-pressure environment at the Venusian surface (485 °C, 93 atm) or in low-temperature environments such as those of Titan (-180 °C), Europa (-220 °C), Ganymede (-200 °C), Mars, the Moon, asteroids, comets, and other small bodies.

Also, Europa-Jupiter missions may have a mission life of 10 years, and the radiation environment is estimated at 2.9 Mrad total ionizing dose (TID) behind 0.1-in-thick aluminum. Proposals are sought for technologies that enable NASA's long-duration missions to extreme wide-temperature and cosmic radiation environments. High reliability, ease of maintenance, low volume, low mass, and low outgassing characteristics are highly desirable. Special interest lies in development of the following technologies that are suitable for the environments discussed above:

Wide-temperature-range precision mechanisms: for example, beam-steering, scanner, linear, and tilting multi-axis mechanisms.
Radiation-tolerant/radiation-hardened low-power, low-noise, mixed-signal mechanism control electronics for precision actuators and sensors.
Wide-temperature-range feedback sensors with sub-arcsecond/nanometer precision.
Long-life, long-stroke, low-power, and high-torque force actuators with sub-arcsecond/nanometer precision.
Long-life bearings/tribological surfaces/lubricants.
High-temperature analog and digital electronics, electronic components, and in-circuit energy storage (capacitors, inductors, etc.) elements.
High-temperature actuators and gear boxes for robotic arms and other mechanisms.
Low-power and wide-operating-temperature radiation-tolerant/radiation-hardened radio-frequency (RF) electronics.
Radiation-tolerant/radiation-hardened low-power/ultralow-power, wide-operating-temperature, low-noise mixed-signal electronics for spaceborne systems such as guidance and navigation avionics and instruments.
Radiation-tolerant/radiation-hardened wide-operating-temperature power electronics.
Radiation-tolerant/radiation-hardened electronic packaging (including shielding, passives, connectors, wiring harness, and materials used in advanced electronics assembly).

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:
Level 1: TX 04 Robotics Systems
Level 2: TX 04.2 Mobility

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware
Research
Analysis

Desired Deliverables Description:

Provide research and analysis for Phase I as a final report. Deliverables for Phase II should include proof-of-concept working prototypes that demonstrate the innovations defined in the proposal and enable direct operation in extreme environments.

Research should be conducted to demonstrate technical feasibility during Phase I and show a path toward a Phase II hardware demonstration, and when possible, deliver a demonstration unit for functional and environmental testing at the completion of the Phase II contract.

State of the Art and Critical Gaps:

Future NASA missions to high-priority targets in our solar system will require systems that have to operate at extreme environmental conditions. NASA missions to the surfaces of Europa and other ocean worlds bodies will be exposed to temperatures as low as -180 °C and radiation levels that are at megarad levels. Operation in permanently shadowed craters on the Moon is also a region of particular interest. In addition, NASA missions to the Venus surface and deep atmospheric probes to Jupiter or Saturn will be exposed to high temperatures, high pressures, and chemically corrosive environments.

Current state-of-practice for development of space systems for the above missions is to place hardware developed with conventional technologies into bulky and power-inefficient environmentally protected housings. The use of environmental-protection housing will severely increase the mass of the space system and limit the life of the mission and the corresponding science return. This solicitation seeks to change the state of the practice by support technologies that will enable development of lightweight, highly efficient systems that can readily survive and operate in these extreme environments without the need for the environmental protection systems.

All proposals relevant to the scope described above would be eligible to be considered for award. For proposals featuring technologies intended for use in planetary science applications, this year a preference will be given to those proposals that would benefit in situ studies of icy ocean worlds, especially techniques that would be beneficial to systems that will descend through kilometers of cryogenic ice, acquire and communicate scientific observations during descent, and sample and concentrate meltwater and interior oceans.

Relevance / Science Traceability:

Relevance to SMD (Science Mission Directorate) is high.

Low-temperature survivability is required for surface missions to Titan (-180 °C), Europa (-220 °C), Ganymede (-200 °C), small bodies, and comets. Mars diurnal temperatures range from -120 °C to +20 °C. For the Europa Clipper baseline concept with a mission life of 10 years, the radiation environment is estimated at 2.9 Mrad TID behind 0.1-in-thick aluminum. Lunar equatorial region temperatures swing from -180 °C to +130 °C during the lunar day/night cycle, and shadowed lunar pole temperatures can drop to -230 °C.

Advanced technologies for high-temperature systems (electronic, electromechanical, and mechanical) and pressure vessels are needed to ensure NASA can meet its long-duration (days instead of hours) life target for its science missions that operate in high-temperature and high-pressure environments.

References:

Proceedings of the Extreme Environment Sessions of the IEEE Aerospace Conference: https://www.aeroconf.org/ or via IEEE Xplore Digital Library.
Proceedings of the meetings of the Venus Exploration Analysis Group (VEXAG): https://www.lpi.usra.edu/vexag/

Proceedings of the meetings of the Outer Planet Assessment Group (OPAG): https://www.lpi.usra.edu/opag/

Lead Center: JPL

Participating Center(s): GSFC

Scope Title: Contamination Control (CC) and Planetary Protection (PP) Implementation and Verification

Scope Description:

The CC and PP subtopic develops new technologies or supports new applications of existing technologies to clean spacecraft, instrumentation, or hardware, while assessing for molecular and biological contaminants to improve NASA's ability to prevent forward and backward contamination.

CC prevents the degradation of the performance of space systems due to particulate and molecular contamination. For CC efforts, understanding and controlling particulate and molecular contaminants supports the preservation of sample and science integrity and ensures spacecraft function nominally. NASA is seeking analytical and physics-based modeling technologies and techniques to quantify and validate submicron particulate contamination; low-energy surface material coatings to prevent contamination; modeling and analysis of particles and molecules to ensure hardware and instrumentation meet organic contamination requirements; and improved technologies for the detection and verification of low levels of organic compounds on spacecraft surfaces.

PP prevents forward and backward contamination to protect planetary bodies, including the Earth, during responsible exploration. Forward contamination is the transfer of viable organisms and bacterial endospores from Earth to another planetary body. Backward contamination is the transfer of biological material, with the potential to cause harm, from a planetary body to Earth's biosphere. Understanding potential CC and PP contaminants and preventing the contamination of our spacecraft and instruments in general also supports the integrity of NASA sample science and mitigates other potential impacts to spacecraft function.

NASA is seeking innovative approaches to address these challenges through:

Improvements to spacecraft cleaning and sterilization that are compatible with spacecraft materials and assemblies.
Prevention of recontamination and cross contamination throughout the spacecraft lifecycle.
Advanced technologies for the detection and verification of organic compounds and biologicals on spacecraft, specifically for microbial detection and assessments for viable organism and deoxyribonucleic-acid- (DNA-) based verification technologies and that may encompass sampling devices, sample processing, and sample analysis pipelines.
Active in situ recontamination/decontamination approaches (e.g., in situ heating of sample containers to drive off volatiles prior to sample collection) and in situ/in-flight sterilization approaches (e.g., UV or plasma) for surfaces.
Development of analytical and modeling-based methodologies to address bioburden and probabilistic risk assessment biological parameters to be used as alternatives to demonstrate requirement compliance.
Enabling end-to-end sample return functions to ensure containment and pristine preservation of materials gathered on NASA missions (e.g., development of technologies that support in-flight verification of sample containment or in-flight correctable sealing technologies).

Examples of outcomes:

End-to-end microbial reduction/sterilization technology for larger spacecraft subsystems.
Microbial reduction/sterilization technology for spacecraft components.
Ground-based biological contamination/recontamination mitigation system that can withstand spacecraft assembly and testing operations.
In-flight spacecraft component-to-component cross-contamination mitigation system.
Spacecraft sterilization systems for target body ground operations.
Viable organism and/or DNA sample collection devices, sample processing (e.g., low biomass extraction), and sample analysis (e.g., bioinformatics pipelines for low biomass).
Real-time, rapid device for detection and monitoring of viable organism contamination on low-biomass surfaces or in cleanroom air.
Bioburden spacecraft cleanliness monitors for assessing surface cleanliness throughout flight and surface operations during missions.
DNA-based system to elucidate abundance, diversity, and planetary protection relevant functionality of microbes present on spacecraft surfaces.
An applied molecular identification technology to tag/label biological contamination on outbound spacecraft.
Molecular mapping and detection technology for organic contamination on outbound and returned spacecraft and spaceflight hardware.
Low surface area energy coatings.
Molecular adsorbers (“getters”).
Technologies to assess human contamination vectors and safety for missions traveling to the Earth’s Moon and human missions traveling to Mars.
Experimental technologies for measurement of outgassing rates lower than 1.0×10^-15 g/cm²/sec with mass spectrometry, under flight conditions (low and high operating temperatures) and with combined exposure to natural environment (e.g., high-energy radiation, ultraviolet radiation, atomic oxygen exposure).
Physics-based technologies for particulate and molecular transport modeling and analysis for complex geometries with moving elements (e.g., rotating solar arrays, articulating robotic arms) in continuum, rarefied, and molecular flow environments, with additional physics (e.g., electrostatic, vibro-acoustic, particle detachment and attachment capabilities).
A ground-based containment system that protects the Earth from restricted Earth-return samples, protects the samples from terrestrial contamination and allows for hardware manipulation and preliminary characterization of samples (e.g., double-walled isolators).

Expected TRL or TRL Range at completion of the Project: 2 to 6

Primary Technology Taxonomy:
Level 1: TX 07 Exploration Destination Systems
Level 2: TX 07.3 Mission Operations and Safety

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I deliverable: As relevant to the proposed effort—proof-of-concept study for the approach to include data validation and modeling.
Phase II deliverable: As relevant to the proposed effort—detailed modeling/analysis or prototype for testing.
Areas to consider for deliverables: technologies, approaches, techniques, models, and/or prototypes, including accompanying data validation reports and modeling code demonstrating how the product will enable spacecraft compliance with PP and CC requirements.

State of the Art and Critical Gaps:

PP state of the art encompasses technologies from the 1960s to 1970s Viking spacecraft assembly and test era along with some more recent advancements in sterilization and sampling technologies. The predominant means to control biological contamination on spacecraft surfaces is to use some combination of heat microbial reduction processing and mechanical removal via solvent cleaning processes (e.g., isopropyl alcohol cleaning). Notably, vapor hydrogen peroxide is a NASA-approved process, but the variability of the hydrogen peroxide concentration, delivery mechanism, and material compatibility concerns still tends to be a hurdle to infuse it on a flight mission with complex hardware and multiple materials for a given component. Upon microbial reduction, during spacecraft integration and assembly, the hardware then is protected in a cleanroom environment (ISO 8 or better) using protective coverings when hardware is not being assembled or tested. For example, terminal sterilization has been conducted with recontamination prevention for in-flight biobarriers employed for the entire spacecraft (Viking) or a spacecraft subsystem (Phoenix spacecraft arm). In addition to the hardware approaches developed for compliance, environmental assessments are implemented to understand recontamination potential for cleanroom surfaces and air. Biological cleanliness is then verified through the NASA standard assay, which is a culture-based method. Although the NASA standard assay is performed on the cleanroom surfaces, DNA-based methodologies have been adopted by some spaceflight projects to include 16S and 18S ribosomal-ribonucleic-acid- (rRNA-) targeted sequencing, with metagenomic approaches currently undergoing development. Rapid cleanliness assessments can be performed, but are not currently accepted as a verification methodology, to inform engineering staff about biological cleanliness during critical hardware assembly or tests that include the total adenosine triphosphate (tATP) and limulus amoebocyte lysate (LAL) assays. Variability in detector performance thresholds in the low biomass limit remain a hurdle in the infusion of ATP luminometers for spaceflight verification and validation. Moreover, with recent missions leveraging probabilistic modeling for biological contamination, modeling has become a key tool in demonstrating compliance and helping to drive biological assurance cases for spacecraft cleanliness. Given the complexity of upcoming missions, this is rapidly becoming an emerging need in the discipline to help define parameters and develop upstream models for understanding biological cleanliness, distributions of biological contamination, behaviors of these biologicals on spacecraft surfaces, transport models, etc. In summary, the critical PP gaps include the assessment of DNA from low-biomass surfaces (<0.1 ng/µL DNA, using current technologies, from 1 to 5 m² of surface); sampling devices that are suitable for reproducible (at a certification level) detection of low biomass and compounds (e.g., viable organisms, DNA) but also compliant with spaceflight environmental requirements (e.g., cleanroom particulate generation, electrostatic discharge limits); quantification of the widest spectrum of viable organisms; enhanced microbial reduction/sterilization modalities that are compatible with flight materials and ground-/flight-/planetary-body-based recontamination prevention/mitigation systems.

CC requirements and practices are also evolving rapidly as mission science objectives targeting detection of organics and life are driving stricter requirements and improved characterization of flight-system- and science-instrument-induced contamination. State-of-the-art CC includes:

Testing and measurement of outgassing rates down to 3.0×10^-15 g/cm²/sec with mass spectrometry, under flight conditions (low and high operating temperatures) and with combined exposure to natural environment (high-energy radiation, ultraviolet radiation, atomic oxygen exposure).
Particulate and molecular transport modeling and analysis for forward contamination scenarios of simple and complex spacecraft geometries with electrostatic, vibro-acoustic, particle detachment and attachment capabilities in continuum, rarefied, and molecular flow environments.
Modeling and analysis of particulate flux for assessment of backward contamination scenarios using dynamic approaches (e.g., direct simulation Monte Carlo (DSMC) and Bhatnagar–Gross–Krook (BGK) formulations).

Relevance / Science Traceability:

With increased interest in investigating bodies with the potential for life detection such as Europa, Enceladus, Mars, and maybe other bodies, and the potential for sample return from such bodies, there is increased need for novel technologies associated with planetary protection and contamination control. The development of such technologies would enable missions to be able to be responsive to PP and CC requirements as they would be able to assess viable organisms and other particulate and organic contaminants; establish microbial reduction and protective technologies to achieve acceptable microbial bioburden and organic contamination levels for sensitive life detection in spacecraft and instruments to mitigate risk and inadvertent “false positives”; ensure compliance with sample return planetary protection and science requirements; and support model-based assessments of planetary protection requirements for biologically sensitive missions (e.g., outer planets and sample return).

References:

Planetary Protection: https://planetaryprotection.nasa.gov/
JPL Planetary Protection Center of Excellence: https://planetaryprotection.jpl.nasa.gov/
Handbook for the Microbial Examination of Space Hardware: https://explorers.larc.nasa.gov/2019APSMEX/SMEX/pdf_files/NASA-HDBK-6022b.pdf

McCoy, K. et. al.: "Europa Clipper planetary protection probabilistic risk assessment summary," Planetary and Space Science, Vol. 196, February 2021, 105139.

Lead Center: ARC

Participating Center(s): AFRC, GSFC, JPL, LaRC

Scope Title: Unpiloted Aerial Platforms for High-Altitude, Long-Endurance (HALE) Missions

Scope Description:

NASA is interested in increased utilization of innovative, cost-effective, unpiloted, aerial platforms, including ones that are heavier and lighter than air, to perform NASA missions in the stratosphere in order to supplement current piloted and satellite platforms. Unmanned aerial platforms are especially suited for HALE missions that occur at or above 50,000 to 90,000 ft and can support continuous flights of 30 days or more at altitude.

HALE missions enable new Earth and space science applications and an opportunity for testing spaceborne-like measurements in the stratosphere. High spatial and temporal resolution observations from HALE can improve measurements of Earth system processes or phenomena requiring sustained observations, including: air quality monitoring, coastal zone and ocean imaging and monitoring, mapping of geologically active regions, forest and agricultural monitoring, and imaging of polar regions. The NASA Surface Biology and Geology mission, for example, is anticipating the need for measurements of leaf canopy chemistry during the growing season, and significant changes can happen between overpasses of polar orbiting satellites. Similarly, the Surface Topography and Vegetation Incubation team recently released a report citing the need for more frequent observations of areas prone to landslides and other ephemeral or episodic events where time series observations can improve Earth system models.

HALE Platforms offer several key challenges, including solar/battery technologies, operation in regions of harsh radiation and temperatures, vehicle health monitoring, and mission deployment/support at remote locations. Methods for accurate stationkeeping in areas of interest would also have to be developed.

Proposals are solicited for both heavier- and lighter-than-air innovative stratospheric platforms that can operate at an altitude of 60,000 ft or above, for a mission of 30 or more days in duration. The proposed vehicle must be able to carry a scientific instrumentation payload of 22lbs or more on all science missions. The combined system must be able to maintain position within 100 nautical miles of a fixed point on the ground and be able to provide at least 100+ W of sustained power (28 Vdc) to payloads. The platform must also have high band width, line-of-sight payload telemetry, and SATCOM capability to enable beyond visual line of site command and control. Proposals can be based on new design platforms or extensively modified existing platforms to meet the above HALE mission requirements.

The primary focus of proposal should be on vehicle design towards a flight test prototype in Phase II. Other aspects such as concept of operations, vehicle maintenance, vehicle transport and deployment, ground station design, and flight-test planning should also be addressed.

Expected TRL or TRL Range at completion of the Project: 2 to 6

Primary Technology Taxonomy:
Level 1: TX 04 Robotics Systems
Level 2: TX 04.2 Mobility

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware

Desired Deliverables Description:

It is expected that a Phase I effort will consist of a system-level design and a proof-of-concept experiment on one or more key components. Deliverable items for Phase I shall be a final report describing the results of the concept analysis and demonstration of any key component technology developed.

The Phase II effort will focus on the development of a concept prototype and feasibility testing. The Phase II deliverable should include a final report on design concept documentation, test reports, and photos of any prototypes that were built and tested.

State of the Art and Critical Gaps:

NASA Global Hawk unmanned aircraft system (UAS) previously provided HALE capability for NASA Earth Science missions but was retired from this mission in 2020. NASA presently has no UAS platforms serving in this role and is reliant on satellites and piloted aircraft to fly these missions. Currently, NASA Earth Science has needs but no platforms to meet this.

While NASA continues development of super-pressure balloons with extended duration, several lighter-than-air vehicles have recently been developed that can provide capabilities to meet NASA science needs.

Several prototype and proof-of-concept HALE vehicles are under development and flight testing. These next-generation HALE vehicles under development have had a focus on communications, and so payloads relevant to Earth Science have not been demonstrated. These existing platforms, which include both heavier- and lighter-than-air platforms could be modified to meet requirements of this solicitation, or new designs could meet them as well.

Relevance / Science Traceability:

As the impacts of climate change become more pronounced through long-term drought, more frequent and intense wildfires, and an increase in severe weather occurrences, there is increased emphasis on Earth Science missions by NASA, other Government Agencies, and private industry. This includes new technologies and capabilities to enhance our ability to observe and predict effects on the environment and the economy of these more frequently occurring events.

NASA, other Government Agencies, and private companies have also shown increased interest in utilizing UAS platforms, both heavier and lighter than air, for Earth Science data collection, supplementing satellite and piloted Earth Science aircraft. This is largely because of the ability of UAS to perform dull, dirty, difficult, and dangerous missions more easily than other platforms.

There is interest from the highest levels of Government to invest in the domestic UAS manufacturing base to reduce reliance on foreign manufacturers as well as security concerns with foreign UAS platforms and technologies.

References:

Airborne Platforms to Advance NASA Earth System Science Priorities: Assessing the Future Need for a Large Aircraft Committee on Future Use of NASA Airborne Platforms to Advance Earth Science Priorities 2021: https://www.nap.edu/catalog/26079/airborne-platforms-to-advance-nasa-earth-system-science-priorities-assessing - see page 142.
Thriving on Our Changing Planet: A Decadal Strategy for Earth Observation from Space National Academies of Sciences, Engineering 2018: https://www.nap.edu/catalog/24938/thriving-on-our-changing-planet-a-decadal-strategy-for-earth
Observing Earth's changing surface topography and vegetation structure 2021: https://science.nasa.gov/science-red/s3fs-public/atoms/files/STV_Study_Report_20210622.pdf - see page 122.

Scope Title: Unpiloted Aerial Platforms for Extreme Environment Missions on Earth

Scope Description:

NASA is interested in increased utilization of unpiloted aerial platforms, both lighter and heavier than air, for Earth Science missions to supplement current piloted and satellite platforms, taking advantage of unpiloted aerial platforms to perform dull, dirty, difficult, and dangerous missions. These platforms are especially suited for extreme environment missions such as volcano, storm, and wildfire penetration as there would be no risk to humans compared to piloted aircraft.

Numerous Earth Science missions require aircraft to operate in situ or in close proximity to extreme environments. This includes flights into volcano plumes to compare sulfur dioxide concentration measurements with those measured by satellites. Another application is storm penetration where unmanned aircraft system (UAS) platforms are flown into thunderstorms and hurricanes to obtain measures of air pressure, wind conditions, temperatures, and other data used for storm forecasting and weather model development. A third example is operation in wildfires where unmanned aerial vehicles (UAVs) can gather information on emissions and fire behavior.

UAS Platforms designed for operation in extreme environments offer several key challenges to developers. Strong winds in the area of these missions usually ground smaller UAS platforms. Turbulence could cause vehicle upset and loss of control in addition to structural damage. Many UAS platforms are not weather resistant and so cannot operate in visible precipitation. For operation in extremely cold conditions, icing could cause loss of the aircraft.

Proposals are solicited for both heavier- and light-than- air aerial platforms that can operate in the extreme environments described above. Proposed platforms should address the operational challenges described to enable missions to be accomplished with minimal vehicle loss. The proposed vehicle must be able to carry a scientific instrumentation payload on all science missions. The proposal can be based on new design platforms or extensively modified existing platforms to meet the above mission requirements.

The primary focus of the proposal should be on vehicle design. However, other aspects such as concept of operations, vehicle maintenance, vehicle transport and deployment, ground station design, and flight test planning should also be addressed.

Expected TRL or TRL Range at completion of the Project: 2 to 6

Primary Technology Taxonomy:
Level 1: TX 04 Robotics Systems
Level 2: TX 04.2 Mobility

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware

Desired Deliverables Description:

It is expected that a Phase I effort will consist of a system-level design and a proof-of-concept experiment on one or more key components or enabling technologies. Deliverable items for Phase I shall be a final report describing the results of the concept analysis and demonstration of any key component technology developed.

The Phase II effort will focus on the development of a concept prototype and feasibility testing. Phase II deliverables should include a final report on design concept documentation, test reports, and photos of any prototypes that were built and tested.

State of the Art and Critical Gaps:

Currently, most UAS platforms can operate only in the proximity of extreme environments but do not have capability for actual penetration other than with a high probability of vehicle loss. The strong winds and turbulence associated with these environments usually grounds smaller UAS platforms or could cause vehicle upset and loss of control as well as structural damage.

Many UAS platforms are not weather resistant and so cannot operate in precipitation. For operation in extremely cold conditions such as polar regions, icing could cause loss of the vehicle.

Because of the capability and operational limits of current UAS platforms, it may not be possible to capture important Earth Science data in hazardous environments.

NASA ARMD (Aeronautics Research Mission Directorate) and NASA ARMD SBIR technologies as well as technologies developed by universities could be utilized by the proposed to address some of these challenges such as icing detection and removal; gust load alleviation; upset prevention, detection, and recovery; see-and-avoid systems; technologies for beyond visual line of sight operation; and others.

Relevance / Science Traceability:

Because of global warming and associated effects such as long-term drought, more frequent and intense wildfires, and an increase in severe weather occurrences, there is an increased priority of Earth Science missions by NASA, other Government Agencies, and private industry. This includes prediction of, detection of, response to, and measurement of effects on the environment and the economy of these more frequently occurring events.

NASA, other Government Agencies, and private companies have also shown increased interest in utilizing unmanned aircraft system (UAS) platforms, both heavier and lighter than air, for Earth Science data collection, supplementing satellite and piloted Earth Science aircraft. This is largely due to the ability of UAS to perform dull, dirty, difficult, and dangerous missions more easily than other platforms. In addition, simpler UAS platforms could be more easily deployed to quickly respond to events of interest.

In addition, there is interest from the highest levels of government to invest in the domestic UAS manufacturing base to reduce reliance on foreign manufacturers such as DJI as well as security concerns with foreign UAS platforms and technologies.

Historically it has been difficult to operate UAS platforms in the National Airspace, primarily because of safety concerns. A large amount of planning, coordination, and approvals were required, making a quick response nearly impossible. Less restrictive operational requirements as developed by the NASA UAS in the NAS program and advances in UAS Air Traffic Technologies developed under the NASA UAS Traffic Management (UTM) Project have enabled simpler, safer, and more efficient UAS flight operations both for private companies and for NASA Earth Science missions.

Advances in UAS technologies, developed under the NASA Aeronautics Research Mission Directorate (ARMD), have enables more capable, less expensive, and higher performing platforms, resulting in an increase of small, innovative, domestic UAS manufacturers. This pool of UAS companies have the expertise and capabilities to develop UAS platforms for future NASA Earth Science missions as well as to commercialize these platforms for non-NASA users.

References:

Airborne Platforms to Advance NASA Earth System Science Priorities: Assessing the Future Need for a Large Aircraft Committee on Future Use of NASA Airborne Platforms to Advance Earth Science Priorities 2021: https://www.nap.edu/catalog/26079/airborne-platforms-to-advance-nasa-earth-system-science-priorities-assessing
Thriving on Our Changing Planet: A Decadal Strategy for Earth Observation from Space National Academies of Sciences, Engineering 2018: https://www.nap.edu/catalog/24938/thriving-on-our-changing-planet-a-decadal-strategy-for-earth

Scope Title: Lighter-than-air platform subsystems for Earth and Venus

Scope Description:

1. Venus lighter-than-air platform:

NASA is interested in scientific exploration of Venus using aerial vehicles to perform in situ investigations of its atmosphere and is currently developing concepts for variable-altitude balloons operating at an altitude range between 52 and 62 km.

One concept for a variable-altitude balloon features a super-pressure (SP) balloon located within a zero-pressure (ZP) balloon. The configuration can be described as one small balloon inside a large balloon that are co-located at the bottom. Altitude changes are made by transfer of helium between the two balloons. Pumping helium from the ZP balloon into the SP balloon reduces buoyancy to descend in altitude. Venting helium from the SP balloon into the ZP balloon increases buoyancy to ascend in altitude. Isolating the ZP and SP balloons when neither the pump or vent is operated enables the balloon to float at constant altitude. Details on the variable-altitude balloon system concept can be found in [Hall 2021].

Proposals for an innovative balloon altitude modulation system featuring a lightweight, high-efficiency pump, isolation valves, and venting orifices are desired. The performance requirements of the balloon altitude modulation system will vary depending on the size of the balloon system and payload. For the purposes of adequately scaling this effort, the following specifications represent the requirements for a current Venus balloon concept (the fluid medium is helium gas):

The pump shall have a nominal flow rate of 250 liters per minute at a pressure rise of 30 kPa.
The vent shall have a nominal flow rate of 1,000 liters per minute at a pressure drop of 5 kPa.

For reliability purposes, the mission operating lifetime is about 100 days of continuous operations.

Typical commercial pumps with this pressure rise and flow rate have a mass around 15 kg and require 250 W of power. Ground-breaking solutions to reduce pump mass to <7 kg and reduce power to <120 W are goals for the specified flow rate and operating pressure.

The specified pressure and flow requirements are current best estimates and will not change during the Phase I proposal development period but may be updated for Phase II.

Venus features a challenging atmospheric environment that significantly impacts the design and operation of devices on aerial vehicles. Proposers should be familiar with the properties of the Venus atmosphere as described in this call. Additional information on the Venus atmospheric environment can be found in the References section.

2. Earth Lighter-than-air platform:

NASA is also looking for an innovative way to reduce the termination dispersions from a few miles to within 1/2 to 1/4 mile of the predicted termination point by the use of a steerable parachute recovery system (SPRS). The SPRS will need to be able to maneuver around infrastructure (e.g., oil wells, power lines, wind mills), protected areas (e.g., national parks, special habitats), natural resources (e.g., rivers, mountains, lakes), and other areas of interest (e.g., farm land). The SPRS will need to provide real-time maneuverability for a science gondola from a remote operations control room using the communications and telemetry systems provided by the Columbia Scientific Balloon Facility (CSBF). The system should be lightweight—no more than 75 lb—including power.

Expected TRL or TRL Range at completion of the Project: 2 to 6

Primary Technology Taxonomy:
Level 1: TX 15 Flight Vehicle Systems
Level 2: TX 15.6 Vehicle Concepts

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware

Desired Deliverables Description:

The deliverables for Phase I include a trade study of the potential systems, a simulation of how each system should work, and a report on the recommendation of one to two systems to be further developed in Phase II. It is anticipated that these products are achievable given the SBIR time and funding constraints.

The deliverables for Phase II include an engineering development unit and flight testing with a report of the results.

State of the Art and Critical Gaps:

1. Venus lighter-than-air platform state of the art:

There are few commercially available pumps in the market today that have the pressure rise and flow rate capabilities needed for a Venus balloon. Most pumps are not built to be lightweight or efficient, which are of critical importance on a balloon mission. Commercial pumps with the targeted flow rate and pressure capability typically have a mass around 15 kg and require 250 W of power. Isentropic pumping power analysis shows that only 80 W of power are required to achieve the desired flow rate and pressure rise. Therefore, the thermodynamic efficiency of commercial pumps is only about 33%. The Venus balloon system desires a system that is at least 65% efficient (2x over commercial products) and half the mass of commercial pump systems to maximize resource availability on the balloon system.

2. Earth lighter-than-air platform state of the art:

A scientific balloon floats at an average altitude of 110,000 ft or more and carries science payloads up to 8,000 lb. At the end of a scientific balloon mission, the science payload on the gondola (“science gondola” from this point on) is separated from the balloon and falls to Earth on a parachute, following the wind currents at the time of release, and then lands on cardboard crush pads. In most cases this allows recovery of the science gondola, although the payload and gondola may be in areas that are hard to reach using conventional recovery trucks. However, there are rare cases where the science gondola falls either into water or in areas that require special equipment or are difficult for recovery (i.e., inaccessible area). Currently, trajectory predictions for termination are within a few miles and are dependent on models, map overlays (showing restricted air space, national/state parks), and observations from a plane on areas along the trajectory to determine the best area to terminate the balloon and bring the science gondola safely to the ground. Some items that are considered during the termination discussions are science mission minimums, trajectory predications (e.g., national or state parks, lakes, mountains, rivers, infrastructure, crop lands), weather conditions, and risk to the public.

Current state of the art does not include steerable systems in balloon parachutes. Success in this endeavor will primarily entail steerability, and will also frequently result in a safety analysis, which will allow more “green lights” for launch than would otherwise be the case.

Relevance / Science Traceability:

1. Venus Lighter-than-air platform relevance:

The Mars Helicopter, Ingenuity, and the Titan Dragonfly mission show there is significant interest in planetary aerial vehicles for science investigations. It is in NASA's interest through the SBIR program to continue fostering innovative ideas to extend our exploration capabilities by developing technologies for Venus aerial mission concepts.

The NASA Jet Propulsion Laboratory's (JPL's) Solar System Mission Formulation Office and Science Mission Directorate's (SMD's) Planetary Science Division advocate Venus aerial vehicle platform development. NASA recently completed the Venus Flagship Mission concept study, which included a balloon system for the Planetary Decadal Survey [Gilmore, 2020].

Science traceability: The 2019 VEXAG Venus Strategic Plan identified several key science investigations that are ideally suited to aerial platforms. The areas of scientific interest include Atmospheric Gas Composition, Cloud and Haze Particle Characterization, Atmospheric Structure, Surface Imaging, and Geophysical Investigations. The variable-altitude aerial vehicle platform is ideal for investigating these science goals and objectives. Building the variable-altitude balloon requires the development of several key components such as the helium transfer system identified in this call.

2. Earth Lighter-than-air platform relevance:

This subtopic will be relevant to any mission directorate, commercial entity, or other government agency that drops payloads from an altitude, including the Balloon Program. Other potentially interested projects include NASA sounding rockets, unmanned aerial vehicles (UAVs), and aircraft programs.

References:

Venus Lighter-than-air platform:

Crisp, D.: “Radiative forcing of the Venus mesosphere I: Solar fluxes and heating rates,” Icarus, 67, pp. 484-514, 1986.
Gilmore, M., et al.: “Venus Flagship Mission Planetary Decadal Study,” Planetary Mission Concept Studies Virtual Workshop, 2020.
Hall, J., et al.: “Prototype Development of a Variable Altitude Venus Aerobot”, AIAA Aviation Forum, 2021.
Knollenberg and Hunten: "The microphysics of the clouds of Venus: Results of the Pioneer Venus Particle Size Spectrometer Experiment," JGR, 85, pp. 8039-8058, 1980, doi:10.1029/JA085iA13p08039.
Oschlisniok, J. et al.: "Microwave absorptivity by sulfuric acid in the Venus atmosphere: First results from the Venus Express Radio Science experiment VeRa," Icarus, 221, p. 940, 2012.
Titov, D., Ignatiev, N. I., K. McGouldrick, Wilquet, V., and Wilson, C. F.: “Venus III: Clouds and hazes of Venus,” Space Sci. Rev., 214, 126, 2018.
The VEXAG Strategic Plan 2019: https://www.lpi.usra.edu/vexag/documents/reports/Combined_VEXAG_Strategic_Documents_2019.pdf
The Venus Atmospheric Properties are available in Kliore, A. J., Moroz, V. I., Keating G. M., Eds.: “The Venus International Reference Atmosphere,” Adv. Space Res., Vol. 5, No. 11, pp 8+305, 1985, ISBN 0-08-034631-6.

Earth Lighter-than-air platform:

JPADS: circumventing GPS for next-gen precision airdrops: https://patents.google.com/patent/EP1463663A4/en https://www.airforce-technology.com/features/featurejpads-circumventing-gps-for-next-gen-precision-airdrops-4872436/

Lead Center: GSFC

Participating Center(s): JPL, LaRC, MSFC

Scope Title: Analog-to-Digital Conversion Components

Scope Description:

NASA's space-based observatories, flyby spacecraft, orbiters, landers, and robotic and sample-return missions require robust command and control capabilities. Advances in technologies relevant to command and data handling and instrument electronics are sought to support NASA's goals and several missions and projects under development.
The 2022 subtopic goals are to develop platforms for the implementation of miniaturized highly integrated avionics and instrument electronics that:

Are consistent with the performance requirements for NASA missions.
Minimize required mass/volume/power as well as development cost/schedule resources.
Can operate reliably in the expected thermal and radiation environments.

Successful proposal concepts should significantly advance the state of the art. Furthermore, proposals developing hardware should indicate an understanding of the intended operating environment, including temperature and radiation. Note that environmental requirements vary significantly from mission to mission. For example, some low-Earth-orbit missions have a total ionizing dose (TID) radiation requirement of less than 10 krad(Si), whereas planetary missions can have requirements well in excess of 1 Mrad(Si).

Specific technologies sought by this scope include:

Radiation-hardened mixed-signal structured application-specific integrated circuit (ASIC) platforms to enable miniaturized and low-power science sensor readout and control, with sufficient capability to implement 12-bit digital-to-analog converters (DACs), monotonic and 12- to 16-bit analog-to-digital converters (ADCs) (<100 kHz 16-bit and 1 to 2 MHz 12-bit), and also charge-sensitive amplifiers for solid-state detectors and readout integrated circuit (ROIC) for silicon photomultipliers.
Low-power, radiation-hardened ASIC devices to enable direct capture of analog waveforms.

Expected TRL or TRL Range at completion of the Project: 1 to 4

Primary Technology Taxonomy:
Level 1: TX 02 Flight Computing and Avionics
Level 2: TX 02.1 Avionics Component Technologies

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware
Software
Analysis
Research

Desired Deliverables Description:

Desired Phase I deliverables include the design, simulation, and analysis to demonstrate viability of proposed component.

Desired Phase II deliverables include a prototype mixed-signal ASIC implemented with a proof-of-concept end-user design. The proof-of-concept design should demonstrate the stated performance capabilities of the ASIC.

State of the Art and Critical Gaps:

There is a need for a broader range of mixed-signal structured ASIC architectures. This includes the need for viable options for mixed ASICs with high-resolution, low-noise analog elements, especially 12-bit DACs and 12- to 16-bit ADCs. The current selection of mixed-signal structured ASICs is limited to 10-bit designs, which do not provide the accuracy or resolution to perform the science required of many of the instruments currently being flown. Mixed-signal structured ASICs can integrate many functions and therefore can save considerable size, weight, and power over discrete solutions—significantly benefiting NASA missions. The lack of parts with high-precision analog is greatly limiting their current application.

Relevance / Science Traceability:

Mixed-signal structured ASIC architectures are relevant to increasing science return and lowering costs for missions across all Science Mission Directorate (SMD) divisions. However, the benefits are most significant for miniaturized instruments and subsystems that must operate in harsh environments. These missions include interplanetary CubeSats and SmallSats, outer-planet instruments, and heliophysics missions to harsh radiation environments. For all missions, the higher accuracy would provide better science or allow additional science through the higher density integration.

References:

The following resources may be helpful for descriptions of radiation effects in electronics:

NASA Technical Reports Server: https://ntrs.nasa.gov/
NASA Electronic Parts and Packaging Program: https://nepp.nasa.gov/
NASA/GSFC Radiation Effects and Analysis Home Page: https://radhome.gsfc.nasa.gov/top.htm

Scope Title: Low-Cost Data Acquisition System

Scope Description:

Destinations such as Mars, Venus, and Titan pose many challenges for entry, descent, and landing (EDL) data acquisition systems, including radiation, g-loading, and volume constraints. Recent notable examples of such systems are the Mars Entry, Descent, and Landing Instrumentation (MEDLI) and MEDLI2 sensor suites, which successfully acquired EDL data in 2012 and 2021, respectively. The NASA MEDLI and MEDLI2 data acquisition systems were very well designed and robust to the extreme environments of space transit and EDL but came at a great financial burden to these missions. The high cost prohibits smaller mission classes such as Discovery and New Frontiers from using MEDLI-like systems, therefore limits the EDL science that can be conducted by NASA. In an effort to bring EDL instrumentation to all missions, NASA seeks a low-cost, robust, high-accuracy data acquisition system. Wireless data acquisition capability would eliminate external radio-frequency interference coupling effects and represents a significant cost and mass savings opportunity on future NASA missions. For example, the sensor cable mass for the Orion Exploration Flight Test 1 (EFT-1) Developmental Flight Instrumentation (DFI) suite was 700 lb. of the entire 1200-lb DFI system. A wireless option for the low-cost data acquisition system is therefore highly desirable.

Data acquisition requirements:

Compatibility: Minimum 15 thermocouples (minimum of 2 Type R and minimum of 8 Type K) and 8 pressure transducers (120- or 350-ohm bridge).
Power: 16 W or less.
Size: Modularity encouraged, max. module size of 10 cm³, four modules max.
Measurement resolution: 12 bit or higher.
Acquisition rate: 8 Hz or higher.
Weight: 5 kg or less.
Accuracy: +/-0.5% of FSR (full scale range).
Radiation tolerant by design: Minimum of 10 krad (30 krad or better desired).
Axial loading capability: minimum 15g (Venus missions could require 100g to 400g).
Temperature capability: -40 to +85 °C.
Cost: Fully qualified target of ~$1M (recurring).

Optional wireless capability:

Centralized or distributed architecture.
Scalable architecture.
0.0% packet loss.
Capable of operating independently for a minimum of 2 years.
Completely wireless: data acquisition and communication powered by a battery or harvested energy (e.g., solar, thermal).

Expected TRL or TRL Range at completion of the Project: 1 to 4

Primary Technology Taxonomy:
Level 1: TX 02 Flight Computing and Avionics
Level 2: TX 02.2 Avionics Systems and Subsystems

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I deliverables would include electrical system design, trade studies, component selections, requirements definitions, and systems analysis to result in a modeled and analyzed data acquisition system architecture. Early breadboard circuits or prototypes may be included.

Phase II deliverables would include production of a prototype low-cost data acquisition system and results from electrical performance testing. Testing may include some environmental and stress testing.

State of the Art and Critical Gaps:

The NASA MEDLI and MEDLI2 data acquisition systems were very well designed and robust to the extreme environments of space transit and EDL, but this comes at a great financial burden to these missions. The high cost prohibits smaller mission classes such as Discovery and New Frontiers from using MEDLI-like systems, therefore limiting the EDL science that can be conducted by NASA. To bring EDL instrumentation to all missions, NASA seeks a low-cost, robust, high-accuracy data acquisition system.

Relevance / Science Traceability:

This technology would be especially relevant to upcoming Science Mission Directorate (SMD) planetary missions, such as DAVINCI and VERITAS, but low-cost data acquisition systems with these capabilities would also be relevant to the other science lines of business, especially for future cost and volume-constrained and distributed-systems missions.

References:

MEDLI2: https://www.nasa.gov/directorates/spacetech/game_changing_development/projects/MEDLI-2
MEDLI: https://mars.nasa.gov/msl/spacecraft/instruments/medli/
NASA Selects 2 Missions to Study ‘Lost Habitable’ World of Venus: https://www.nasa.gov/press-release/nasa-selects-2-missions-to-study-lost-habitable-world-of-venus
NASA to Explore Divergent Fate of Earth’s Mysterious Twin with Goddard’s DAVINCI+: https://www.nasa.gov/feature/goddard/2021/nasa-to-explore-divergent-fate-of-earth-s-mysterious-twin-with-goddard-s-davinci
VERITAS: https://www.jpl.nasa.gov/missions/veritas

Scope Title: Printed High Density Interconnects

Scope Description:

As the size of circuit boards continues to shrink and electronic component sizing continues to approach bare die form factors, NASA's need for high-reliability, high-density interconnection solutions is increasing. The ability to connect components or even larger assemblies together without the need for conventional connectors and harnessing stands to offer significant advantages to the size and weight requirements of command, data handling, and electronics systems. High-reliability interconnect methodologies that can operate in space environments (vacuum, vibration) and deliver hundreds of signal/power connections while using as little physical board area as possible are desired.

Chip-scale interconnection methodologies such as wirebonding are size and volume efficient, but present manufacturing, reliability, and handling challenges when applied in an exposed manner on otherwise conventional circuit board assemblies. NASA seeks manufacturing technologies that could be applied at the circuit board assembly level to create high-reliability, high-density electrical connections across three-dimensional (3D) topologies, such as connecting to the top surface of microcircuit die adhered to a substrate. Emerging additively manufactured and printed hybrid electronics technologies offer potential solutions that also address the reliability and handling challenges present with larger assembly implementations, but further development is needed to demonstrate performance and reliability for NASA applications.

Specifically, NASA is seeking:

Capability to reliably print or produce 400 or more conductive traces, on the order of 50 to 100 µm width, with 100 to 200 µm pitch.
Capability to reliably print or produce traces that can traverse a vertical shift to an elevated topology shifts of up to 1.5 mm in height. Printing or producing fillets or ramps to accommodate smooth transitions for the vertical topology shifts is acceptable.
Electrical resistivity of the traces shall be no more than 300 ohm/mm, and isolation to adjacent traces shall be on the order of gigaohms.
Printed or produced traces shall demonstrate alignment to target features on the substrates and the elevated topology surfaces.
Demonstrated reliability and workmanship testing performance, such as vibration and thermal cycling.

Expected TRL or TRL Range at completion of the Project: 1 to 4

Primary Technology Taxonomy:
Level 1: TX 02 Flight Computing and Avionics
Level 2: TX 02.2 Avionics Systems and Subsystems

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I deliverables would include development of prototype design, materials selection and trade studies, production of necessary equipment fixtures and tooling, and ultimately demonstration of the proposed interconnect manufacturing.

Phase II deliverables would include refinement of prototype designs, demonstration of consistent print production across multiple samples, electrical performance, and results of workmanship and reliability tests of produced designs.

State of the Art and Critical Gaps:

The current assembly process for arrays of die and sensors is wire bonding. However, as die become smaller and die pads become smaller and denser, this pushes the limits of wire bonding capabilities. The next generation of NASA science missions have needs for higher density interconnect solutions. Printed hybrid electronics technologies are emerging; however, they have not yet demonstrated suitable repeatability and reliability for use in NASA applications.

Relevance / Science Traceability:

These technologies would be broadly beneficial to command and data handling (C&DH) architectures on many NASA missions. There is also a crossover need for this technology on high-density detector systems that will be needed for NASA's next-generation science missions.

Missions/Programs/Projects that could use the technology:
- Large UV/Optical/IR Surveyor (LUVOIR).
- Habitable Exoplanet Observatory (HabEx).
- Cosmic Evolution Through UV Spectroscopy (CETUS).

References:

The following resources may be helpful for descriptions of radiation effects in electronics:

NASA Technical Reports Server: https://ntrs.nasa.gov/
NASA Electronic Parts and Packaging Program: https://nepp.nasa.gov/
NASA/GSFC Radiation Effects and Analysis Home Page: https://radhome.gsfc.nasa.gov/top.htm
LUVOIR: https://asd.gsfc.nasa.gov/luvoir/
Habitable Exoplanet Observatory (HabEx): https://www.jpl.nasa.gov/habex/
Cosmic Evolution Through UV Spectroscopy (CETUS): https://cor.gsfc.nasa.gov/copag/AAS_Jan2018/Heap_UVisSig_8Jan18.pdf

Scope Title: Intelligent Hardware Supervisors

Scope Description:

The space radiation environment and single-event effects (SEEs) are known to cause errors and interruptions in electronics circuitry. NASA has an increasing need to achieve higher performance processing and microcircuits, and this often requires infusion of commercial electronic parts, which may not be explicitly designed for radiation tolerance. One critical aspect to successfully using these commercial technologies in a space system is being able to recognize when a component has been hit by a SEE and commanding that component to reset itself, without causing major disruption to the entire system.

To this goal, NASA seeks responsive or intelligent hardware supervisor components for SEEs. Ideally, a microcircuit that can monitor the operational profile of other components and intelligently determine what is and is not a latchup or other event versus a computationally intense processing state, especially for consumer/COTS (commercial-off-the-shelf) electronics.

Expected TRL or TRL Range at completion of the Project: 1 to 4

Primary Technology Taxonomy:
Level 1: TX 02 Flight Computing and Avionics
Level 2: TX 02.1 Avionics Component Technologies

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I deliverables would include system design, trade studies, component selections, requirements definitions, and systems analysis to result in a modeled and analyzed system architecture. Early breadboard circuits or prototypes may be included.

Phase II deliverables would include production of a prototype(s) and electrical performance testing. Testing may include some environmental and stress testing.

State of the Art and Critical Gaps:

Existing hardware supervisors do exist, but they do not fully address the needs of NASA missions seeking to infuse modern COTS components. Supervisor methodologies are either too conservative, and overly reset devices causing undue downtime and data loss or are more intelligent to distinguish upsets but require computationally intense processing and power resources to implement. NASA needs supervisor components that can intelligently determine latchups or other events without a computationally intense processing state.

Relevance / Science Traceability:

These technologies would be relevant to increasing science return and lowering costs for missions across all Science Mission Directorate (SMD) divisions. However, the benefits are most significant for miniaturized instruments and subsystems that must operate in harsh environments. These missions include interplanetary CubeSats and SmallSats, outer planets instruments, and heliophysics missions to harsh radiation environments.

References:

The following resources may be helpful for descriptions of radiation effects in electronics:

NASA Technical Reports Server: https://ntrs.nasa.gov/
NASA Electronic Parts and Packaging Program: https://nepp.nasa.gov/

NASA/GSFC Radiation Effects and Analysis Home Page: https://radhome.gsfc.nasa.gov/top.htm

Lead Center: JPL

Participating Center(s): GSFC

Scope Title: High-Performance Space Computing Technology

Scope Description:

Most current NASA missions utilize 20-year-old space computing technology that is inadequate for future missions. Newer processors with improved performance are becoming available from industry but still lack the performance, power efficiency, and flexibility needed by the most demanding mission applications. The NASA High-Performance Spaceflight Computing (HPSC) project is addressing these needs. This subtopic solicits technologies that can enable future high-performance, multicore processors, along with the supporting technologies needed to fully implement avionics systems based on these processors.

Fault-tolerant internet protocol (IP) core supporting Ethernet, Time-Sensitive Networking (TSN), Time-Triggered Ethernet (TTE), and remote direct memory access (RDMA) over converged Ethernet (RoCE) to support processor clustering.
Compilers that support software-implemented fault tolerance (SIFT) capabilities (e.g., control flow checking, coordinated checkpoint/rollback, recovery block) for multicore processors are desired.
- Compile-time fault tolerance is desired by NASA for reorganizing execution code to automatically build redundancy in stall cycles without requiring additional development from the user; this would be exceptional for performance optimization of code without putting additional burden on the developers. This is increasingly important with the adoption of more complex and commercial processors in future missions.
Radiation-tolerant, point-of-load (POL) converters that feature multiple outputs, intelligent communication, or high power.
- Modern and next-generation processors require multiple voltage supply levels, requiring multiple discrete POL converters occupying valuable processor card real estate. A multiple-output POL would enable smaller and more powerful spaceflight processing platforms.
- Future spaceflight systems have increased needs for fault detection, tolerance, and command ability. A POL converter capable of communicating with command-and-control architectures to report health status, telemetry, or to adjust parameters is desired.
- Future high-powered spaceflight processing applications will have a need for high-power POL converters. Specifically, converters capable of providing low voltage, but high currents (tens of amps) are desired.
Coprocessors to (a) accelerate onboard artificial intelligence applications, or (b) perform digital signal processing (DSP) functions. Specifically, technologies are sought that either enable the reliable use of commercial off-the-shelf (COTS) coprocessors in space systems, or fault-tolerant IP cores that can be implemented in a radiation-hardened field-programmable gate array (FPGA).
Radiation-tolerant solid-state memory drives (minimum 1-TB capacity) with Peripheral Component Interconnect Express (PCIe) interface, supporting file systems with industry-standard Non-Volatile Memory Express (NVMe) software stack.
Checkpointing and recovery mechanism for single-process flight software applications.
- Especially with increased use of COTS processors, single-event functional interrupts (SEFIs) have a high chance the processor will need to reset or incur a kernel panic. NASA desires a way for the flight software to be automatically checkpointed or have some sort of functional save-state to recover before the upset.
- Current methodologies for resetting and recovering processors and flight software applications can incur considerable downtime and data loss. A more intelligent, rapid method for resetting and recovering is desired. NASA's Core Flight Software (cFS) would be an ideal candidate software to implement this capability.

Expected TRL or TRL Range at completion of the Project: 1 to 4

Primary Technology Taxonomy:
Level 1: TX 02 Flight Computing and Avionics
Level 2: TX 02.X Other Flight Computing and Avionics

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware
Software
Research

Desired Deliverables Description:

Phase I Deliverables:

For software and hardware elements, a solid conceptual design, plan for full-scale prototyping, and simulations and testing results to justify prototyping approach. Detailed specifications for intended Phase II deliverables.

Phase II Deliverables:

For software and hardware elements, a prototype that demonstrates sufficient performance and capability and is ready for future development and commercialization.

State of the Art and Critical Gaps:

Most NASA missions utilize processors with in-space-qualifiable high-performance computing that has high power dissipation (approximately 18 W), and the current state-of-practice Technology Readiness Level 9 (TRL-9) space computing solutions have relatively low performance (between 2 and 200 DMIPS (Dhrystone million instructions per second) at 100 MHz). A recently developed radiation-hardened processor provides 5.6 GOPS (giga operations per second) performance with a power dissipation of 17 W. Neither of these systems provides the desired performance, power-to-performance ratio, or flexibility in configuration, performance, power management, fault tolerance, or extensibility with respect to heterogeneous processor elements. Onboard network standards exist that can provide >10 Gbps bandwidth, but not everything is available to fully implement them.

Relevance / Science Traceability:

The high-performance spaceflight computing (HPSC) ecosystem is enhancing to most major programs in the Human Exploration and Operations Mission Directorate (HEOMD). It is also enabling for key Space Technology Mission Directorate (STMD) technologies that are needed by HEOMD, including the Safe and Precise Landing - Integrated Capabilities Evolution (SPLICE) project. Within the Science Mission Directorate (SMD), strong mission pull exists to enable onboard autonomy across Earth science, astrophysics, heliophysics, and planetary science missions. There is also relevance to other high-bandwidth processing applications within SMD, including adaptive optics for astrophysics missions and science data reduction for hyperspectral Earth science missions.

References:

RISC-V: https://riscv.org/news/2019/09/risc-v-gains-momentum-as-industry-demands-custom-processors-for-new-innovative-workloads/
Next Generation Space Interconnect Standard: http://www.rapidio.org/wp-content/uploads/2014/10/RapidIO-NGSIS-Seminar-July-23-2014.pdf
He, J., et al. Provably Correct Systems. Formal Techniques in Real-Time and Fault-Tolerant Systems. pp. 288-335. ProCoS. 1994.
Reis, G.A. SWIFT: Software Implemented Fault Tolerance. International Symposium on Code Generation and Optimization. IEEE. 2004.
Wessman, N., et al. De-RISC: The First RISC-V Space-Grade Platform for Safety-Critical Systems. pg. 17-26. IEEE Space Computing Conference Proceedings. 2021.
Franconi, N., et al. Signal and Power Integrity Design Methodology for High-Performance Flight Computing Systems. pg. 17-26. IEEE Space Computing Conference Proceedings. 2021.
Yanguas-Gil, A., et al. Neuromorphic Architectures for Edge Computing under Extreme Environments. pg. 39-45. IEEE Space Computing Conference Proceedings. 2021.

Sabogal, S., et al. A Methodology for Evaluating and Analyzing FPGA-Accelerated, Deep-Learning Applications for Onboard Space Processing. pg. 143-154. IEEE Space Computing Conference Proceedings. 2021.

Lead Center: JSC

Participating Center(s): N/A

Scope Title: Display Systems

Scope Description:

NASA's vision for human spaceflight requires the crew to execute increasingly complex tasks in more demanding and dangerous environments. As a result, advances in avionics technologies relevant to human interfaces for space systems are sought that can be infused into current and future human spaceflight programs, including orbiting spacecraft, surface habitats, surface mobility vehicles, and spacesuits. The 2022 subtopic goals are to advance technologies that increase the reliability of crew interface systems in the radiation environment beyond Low Earth Orbit (LEO), while also increasing the crew’s capabilities and effectiveness in performing mission tasks. Standards-based interfaces are of particular interest to promote interoperability and equipment reuse across spacecraft.

Successful proposal concepts should significantly advance the state of the art. Furthermore, proposals should indicate an understanding of the safety-critical operations performed by spaceflight crews, as well as the intended radiation environment. Note that environmental requirements vary significantly between space systems and missions, with some spacecraft and surface vehicles supporting human operations for days and others supporting periodic crewed missions for 15 or more years.

Specific technologies sought by this subtopic include display systems capable of supporting long-duration human spaceflight beyond low Earth orbit. Multifunctional visual displays provide the highest bandwidth and most versatile means for crew to receive complex information, but unique component technologies with limited radiation performance data prevent high-reliability displays from being developed. The following design parameters and data are sought for display panel and pixel technologies:

A scalable architecture that permits different levels of performance
Radiation test data, analysis of failure modes, radiation-tolerant designs, and prototype hardware/software solutions
A display panel diagonal measurement of at least 14 in. with the capability to render complex graphics, including high-definition video, at a frame rate of at least 20 frames per second.

Design and performance parameters are driven by use cases requiring crewmembers to directly control the spacecraft using live streaming video, such as in-space docking, controlled landing, robotic operations, and surface mobility.

Expected TRL or TRL Range at completion of the Project: 3 to 7

Primary Technology Taxonomy:
Level 1: TX 02 Flight Computing and Avionics
Level 2: TX 02.2 Avionics Systems and Subsystems

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware
Software
Analysis

Desired Deliverables Description:

The desired Phase I deliverables include designs, simulations, and analyses to demonstrate the viability of proposed designs and components.

The desired Phase II deliverables for display systems include a prototype demonstration of a custom or modified display panel technology that mitigates radiation failure modes of electronic components. The proof-of-concept design should consider scalability and integration with other display components.

State of the Art and Critical Gaps:

Commercial display technologies have been used in LEO on the International Space Station for decades, but radiation test data for complex electronics beyond LEO are very limited, and existing test data indicate displays may be more susceptible to radiation than other electronic components. As a result, spacecraft designers are forced to take an unquantified risk of equipment failure due to radiation effects and to include backup crew interface systems that take up valuable mass, volume, and power on the spacecraft. While ongoing Government and industry investments seek to improve processor and graphics processing unit (GPU) performance, quantifying and improving the radiation tolerance of display panel components remains unaddressed.

Relevance / Science Traceability:

This subtopic is relevant to human spaceflight programs in the development and planning phases, including Gateway, HLS (Human Landing System), Orion, and xEMU (Exploration Extravehicular Mobility Unit), as well as to lunar and Martian surface habitation systems and rovers. Technology solutions developed under this subtopic have the potential for a direct infusion path as these spacecrafts are designed and developed.

Electronic visual displays are required for human spaceflight (NPR 8705.2C, NASA Human-Rating Requirements for Space Systems) and will be at the center of any spacecraft’s crew interface architecture. By quantifying and improving the reliability of radiation-tolerant displays, spacecraft designers will be able to simplify this architecture by reducing the need for redundancy, sparing, and operational constraints while also reducing mass, volume, and power needs.

References:

NASA Electronic Parts and Packaging Program: https://nepp.nasa.gov/NASA/GSFC
Radiation Effects and Analysis Homepage: https://radhome.gsfc.nasa.gov/top.htm
NASA Cross-Program Design Specification for Natural Environments (DSNE): http://ntrs.nasa.gov/citations/20200000867
The Past, Present, and Future of Display Technology in Space: https://arc.aiaa.org/doi/10.2514/6.2010-8915
NASA Active Matrix Organic Light Emitting Diode (AMOLED) Environmental Test Report: https://ntrs.nasa.gov/citations/20140003471
OLED Technology Evaluation for Space Applications: https://ntrs.nasa.gov/citations/20150016975

Scope Title: Audio Systems

Scope Description:

NASA's vision for human spaceflight requires the crew to execute increasingly complex tasks in more demanding and dangerous environments. As a result, advances in avionics technologies relevant to human interfaces for space systems are sought that can be infused into current and future human spaceflight programs, including orbiting spacecraft, surface habitats, surface mobility vehicles, and spacesuits. The 2022 subtopic goals are to advance technologies that increase the reliability of crew interface systems in the radiation environment beyond Low Earth Orbit (LEO), while also increasing the crew’s capabilities and effectiveness in performing mission tasks. Standards-based interfaces are of particular interest to promote interoperability and equipment reuse across spacecraft.

Successful proposal concepts should significantly advance the state of the art (SOA). Furthermore, proposals should indicate an understanding of the safety-critical operations performed by spaceflight crews, as well as the intended radiation environment. Note that environmental requirements vary significantly across space systems and missions, with some spacecraft and surface vehicles supporting human operations for days and others supporting periodic crewed missions for 15 or more years.

Specific technologies sought by this subtopic include audio systems that provide two-way voice communication between crew members and mission personnel on Earth through all mission phases and crew activities. These systems also must annunciate alarms and may provide a means of controlling systems by voice or record field notes. Robust audio system technologies are sought with the following design and performance parameters:

Low-latency G.711 and G.729 audio encoding/decoding and routing from multiple simultaneous sources.
Integrate with Ethernet-based spacecraft networks to route multiple simultaneous audio streams to each user.
Support ad hoc addition and removal of end systems and in-flight configuration and extensibility.
Leverage modular and standards-based hardware and software.
Provide radiation tolerance and fault mitigation.
Incorporate SOA microphones, speakers, and acoustic echo-canceling technologies that improve speech quality and intelligibility for voice communication and speech recognition in acoustically challenging environments, such as noisy habitable modules and spacesuits. NASA human spaceflight programs typically require a speech intelligibility score of 90% per the ANSI S3.2 standard using the Modified Rhyme Test (MRT) method.

Expected TRL or TRL Range at completion of the Project: 4 to 7

Primary Technology Taxonomy:
Level 1: TX 02 Flight Computing and Avionics
Level 2: TX 02.2 Avionics Systems and Subsystems

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware
Software

Desired Deliverables Description:

The desired Phase I deliverables include designs, tabletop hardware/software prototypes, and analyses to demonstrate the viability of proposed designs and components.

The desired Phase II deliverables for display systems include a prototype hardware and software audio system that can be tested in NASA network test facilities with at least three simultaneous audio endpoints. The audio system should be tested for radiation tolerance.

State of the Art and Critical Gaps:

Audio systems are not currently available that meet NASA’s basic functional requirements and can perform reliably in the spaceflight radiation and acoustic environments.

Relevance / Science Traceability:

This subtopic is relevant to human spaceflight programs in the planning phases, including human landing systems (HLSs) and lunar and Martian surface habitation systems and rovers. Technology solutions developed under this subtopic have the potential for a direct infusion path as these spacecrafts are designed and developed.

Voice communication and auditory alarms have been included in NASA spacecraft since the Mercury Program, but this has not been sufficient to sustain a robust commercial market for space-rated audio systems. As NASA and commercial partners have increased new spacecraft development, the dearth of vendors has resulted in substantial schedule, cost, and technical integration risk.

References:

NASA Electronic Parts and Packaging Program: https://nepp.nasa.gov/
Radiation Effects and Analysis Home Page: https://radhome.gsfc.nasa.gov/top.htm
NASA Cross-Program Design Specification for Natural Environments (DSNE): https://ntrs.nasa.gov/citations/20200000867
Space Shuttle Orbiter Audio Subsystem: https://ntrs.nasa.gov/citations/19790056509

Technical Aspects of Acoustical Engineering for the ISS: https://ntrs.nasa.gov/api/citations/20090009764/downloads/20090009764.pdf

Lead Center: JSC

Participating Center(s): GRC, JPL, KSC, MSFC

Scope Title: Oxygen from Regolith

Scope Description:

Lunar regolith is approximately 45% oxygen by mass. The majority of the oxygen is bound in silicate minerals. Previous efforts have shown that it is possible to extract oxygen from regolith using various techniques. NASA is interested in developing novel oxygen extraction systems that can be proven to handle large amounts of lunar regolith throughput while minimizing consumables, mass, and energy. NASA is also interested in developing the supporting technologies that may enable or enhance the ability to extract oxygen from lunar regolith. Each of the following specific areas of technology interest may be proposed as individual efforts or combined.

Novel Silicate Reduction Methods: Proposed concepts should describe a reduction method for highland anorthosite that avoids reduction of the regolith in the molten liquid state (i.e., gas/granular, liquid/granular, or vacuum/granular material processing). If reactants are utilized in the reduction process, and multiple reaction products are generated, all steps in regenerating the reactants and separating the products need to be considered. Proposed concepts must include a method to move regolith through the reaction zone (e.g., regolith inlet/outlet valves capable of passing abrasive granular material through the valve for hundreds of cycles). The target production rate for a pilot plant system is 1,000 kg of oxygen per year. The target production rate for a full-scale system is 10,000 kg of oxygen per year. Since access to continuous power is not initially planned, proposers will need to consider how to stop and restart their reduction method periodically throughout the year.
Noncontact High-Temperature Measurement: Proposed concepts should be capable of determining temperatures up to 2,000ºC without contacting the material being measured (e.g., pyrometer). Compatibility with multiple oxygen extraction methods is desired. Instruments must be capable of operating inside of a vacuum chamber.
Regolith Feed/Removal Systems and Mineral Measurement for Oxygen Removal: For oxygen extraction from regolith systems, it is anticipated that hardware will be required to transfer regolith from excavators to the reduction reactor and to transfer processed regolith from the reduction reactor to a holding hopper or the lunar surface. To better understand and control oxygen/metal extraction processes, NASA would like to examine the regolith material before and after the reduction process. Proposed concepts should describe a regolith feed/removal system that includes instrumentation to determine the amount of oxygen in the regolith upstream and downstream of an oxygen extraction zone. Measurements should be taken at a frequency that accounts for the regolith feed rate. The target regolith feed rate for a pilot plant system is 2.5 kg/hr. The target regolith feed rate for a full-scale system is 25 kg/hr. Compatibility with multiple oxygen extraction methods is desired. Instruments must be capable of operating inside of a vacuum chamber.

Expected TRL or TRL Range at completion of the Project: 4 to 5

Primary Technology Taxonomy:
Level 1: TX 07 Exploration Destination Systems
Level 2: TX 07.1 In-Situ Resource Utilization

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware

Desired Deliverables Description:

Phase I efforts should provide a feasibility study and/or proof of concept. Phase II efforts should demonstrate the technology using lunar regolith simulant where applicable. Phase II efforts should be demonstrated no less than 1/10 of the pilot plant production rate where applicable and should describe how the process can be applied at the full-scale production rate.

State of the Art and Critical Gaps:

Some oxygen-from-regolith methods have been demonstrated at relevant scales and are progressing toward TRL 6. Many other methods have been demonstrated at the bench scale, but current designs lack a means to move regolith in and out of the oxygen extraction zone. Many of these processes are used terrestrially, but industrial designs do not provide a means to keep gases from escaping to the vacuum of space.

Relevance / Science Traceability:

STMD (Space Technology Mission Directorate) has identified the need for oxygen extraction from regolith. The alternative path, oxygen from lunar water, currently has much more visibility. However, we currently do not know enough about the concentration and accessibility of lunar water to begin mining it at a useful scale. Lunar water prospecting missions are required to properly assess the utilization potential of water on the lunar surface. Until more water prospecting data becomes available, NASA recognizes the need to make progress on the technology required to extract oxygen from dry lunar regolith.

References:

Lomax, B. A., Conti, M., Khan, N., Bennett, N. S., Ganin, A. Y., & Symes, M. D. (2020). Proving the viability of an electrochemical process for the simultaneous extraction of oxygen and production of metal alloys from lunar regolith. Planetary and Space Science, 180, 104748.
Schwandt, C., Hamilton, J. A., Fray, D. J., & Crawford, I. A. (2012). The production of oxygen and metal from lunar regolith. Planetary and Space Science, 74(1), 49-56.
Fox, E. T. (2019). Ionic liquid and in situ resource utilization. https://ntrs.nasa.gov/citations/20190027398
Cardiff, E. H., Pomeroy, B. R., Banks, I. S., & Benz, A. (2007, January). Vacuum pyrolysis and related ISRU techniques. In AIP Conference Proceedings (Vol. 880, No. 1, pp. 846-853). American Institute of Physics. https://ntrs.nasa.gov/citations/20070014929
Gustafson, R. J., White, B. C., & Fidler, M. J. (2009). Oxygen production via carbothermal reduction of lunar regolith. SAE International Journal of Aerospace, 4(2009-01-2442), 311-316.
Gustafson, R. J., White, B. C., Fidler, M. J., & Muscatello, A. C. (2010). The 2010 field demonstration of the solar carbothermal reduction of regolith to produce oxygen. https://ntrs.nasa.gov/citations/20110005526
Gustafson, R., White, B., & Fidler, M. (2011, January). 2010 field demonstration of the solar carbothermal regolith reduction process to produce oxygen. In 49th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition (p. 434).
Muscatello, T. (2017). Oxygen extraction from minerals. https://ntrs.nasa.gov/citations/20170001458
Paley, M. S., Karr, L. J., & Curreri, P. (2009). Oxygen production from lunar regolith using ionic liquids. https://ntrs.nasa.gov/citations/20090017882
Sibille, L., Sadoway, D. R., Sirk, A., Tripathy, P., Melendez, O., Standish, E., ... & Poizeau, S. (2009). Production of oxygen from lunar regolith using molten oxide electrolysis. https://ntrs.nasa.gov/citations/20090018064

Scope Title: Lunar Ice Mining

Scope Description:

We now know that water ice exists on the poles of the Moon from data obtained from missions like the Lunar Prospector, Chandrayaan-1, Lunar Reconnaissance Orbiter (LRO), and the Lunar Crater Observation and Sensing Satellite (LCROSS). We know that water is present in permanently shadowed regions (PSRs), where temperatures are low enough to keep water in a solid form despite the lack of atmospheric pressure. Many efforts are now underway to develop technologies needed to extract and capture lunar water ice. However, many other volatiles may be co-located with the water ice that may have additional in situ resource utilization (ISRU) applications. NASA is interested in developing technologies to capture and utilize other volatiles that may be located in PSRs. NASA is also interested in developing the supporting technologies that may enhance efforts to excavate water ice. Each of the following specific areas of technology interest may be proposed as individual efforts or combined.

Nonwater Volatile Capture: Proposed concepts should define a target volatile (e.g., H₂S, NH₃, SO₂, C₂H₄, CO₂, CH₃OH, CH₄) to be captured from lunar regolith and describe how it may be utilized in a way that reduces the cost of landing consumables on the lunar surface. Concepts need to operate in PSRs of the lunar poles (<100K) and collected products will be removed from the PSR and processed in a near-permanently lit location nearby. Concepts to minimize electrical power usage are highly encouraged.
Regolith/Ice Crushing: Proposed concepts should be able to crush frozen regolith simulant with a water ice content of 90% by mass while minimizing temperature increase in the material. The target production rate for a pilot-plant-scale ice-crushing system is 10 kg of regolith per hour. The target production rate for a full-scale system is 100 kg of regolith per hour. Concepts should consider how volatiles released during crushing may be minimized or captured if a significant fraction are lost.

Expected TRL or TRL Range at completion of the Project: 4 to 5

Primary Technology Taxonomy:
Level 1: TX 07 Exploration Destination Systems
Level 2: TX 07.1 In-Situ Resource Utilization

Desired Deliverables of Phase I and Phase II:

Prototype
Analysis
Hardware

Desired Deliverables Description:

Phase I efforts should provide a feasibility study and/or proof of concept. Phase II efforts should demonstrate the technology using lunar regolith simulant where applicable. Phase II efforts should be demonstrated no less than 1/10 of the pilot plant production rate where applicable and should describe how the process can be applied at the full-scale production rate.

State of the Art and Critical Gaps:

Multiple efforts are now underway to extract, purify, and capture lunar water ice. However, little work has been performed on developing technologies to capture and utilize other useful volatiles that may be co-located within a PSR. Ice-crushing technology was developed at a small scale to support the Regolith and Environment Science and Oxygen and Lunar Volatiles Extraction (RESOLVE) project, but little work has been performed for larger scale applications.

Relevance / Science Traceability:

NASA has referenced water ice as one of the reasons we have chosen the lunar poles as the location to establish a sustained human presence. The Space Technology Mission Directorate (STMD) has identified the need for water extraction technologies. The Science Mission Directorate (SMD) is currently funding the Volatiles Investigating Polar Exploration Rover (VIPER) mission to investigate lunar water ice.

References:

Colaprete, A., Schultz, P., Heldmann, J., Wooden, D., Shirley, M., Ennico, K., ... & Goldstein, D. (2010). Detection of water in the LCROSS ejecta plume. Science, 330(6003), 463-468.
Gladstone, G. R., Hurley, D. M., Retherford, K. D., Feldman, P. D., Pryor, W. R., Chaufray, J. Y., ... & Stern, S. A. (2010). LRO-LAMP observations of the LCROSS impact plume. Science, 330(6003), 472-476.
Hibbitts, C. A., Grieves, G. A., Poston, M. J., Dyar, M. D., Alexandrov, A. B., Johnson, M. A., & Orlando, T. M. (2011). Thermal stability of water and hydroxyl on the surface of the Moon from temperature-programmed desorption measurements of lunar analog materials. Icarus, 213(1), 64-72.
Poston, M. J., Grieves, G. A., Aleksandrov, A. B., Hibbitts, C. A., Darby Dyar, M., & Orlando, T. M. (2013). Water interactions with micronized lunar surrogates JSC‐1A and albite under ultra‐high vacuum with application to lunar observations. Journal of Geophysical Research: Planets, 118(1), 105-115.
Mortimer, J., Lécuyer, C., Fourel, F., & Carpenter, J. (2018). D/H fractionation during sublimation of water ice at low temperatures into a vacuum. Planetary and Space Science, 158, 25-33.
Environment Science and Oxygen and Lunar Volatiles Extraction (RESOLVE), https://ntrs.nasa.gov/citations/20150022136
Volatiles Investigating Polar Exploration Rover (VIPER), https://www.nasa.gov/viper

Scope Title: Size Sorting, Beneficiation, and Metal Production

Scope Description:

Size sorting and beneficiation can be applied to ice mining and to oxygen extraction from regolith. Size sorting is a necessary step in any in situ resource utilization (ISRU) process involving regolith to ensure that the regolith delivered to an ISRU plant does not include objects large enough to cause mechanical failures within the system. Beneficiation allows for improved efficiency of ISRU processes that involve heating regolith to acquire a specific resource. NASA is also interested in processes where the primary product is metal—specifically metals other than iron, since iron extraction from regolith is a fairly advanced technology. Each of the following specific areas of technology interest may be proposed as individual efforts or combined.

Size Sorting: Proposed concepts should demonstrate a means to remove particles larger than 1 mm from a feedstock of lunar regolith simulant. The target production rate for a pilot-plant-scale system is 10 kg of regolith per hour. The target production rate for a full-scale system is 100 kg of regolith per hour.
Beneficiation of Water Ice: Proposed concepts should describe a method for separating water ice from bulk regolith without causing the water ice to change phase. The described method should address how sublimation losses can be minimized. The target production rate for a pilot-plant-scale system is 10 kg of regolith per hour. The target production rate for a full-scale system is 100 kg of regolith per hour.
Mineral Beneficiation: Proposed concepts should define a target mineral to be concentrated from lunar regolith feedstock and describe how it will be utilized in a way that reduces the cost of landing consumables on the lunar surface.
Metal Production: Proposed concepts should define a target metal (e.g., aluminum) to be extracted from lunar regolith and describe how it will be utilized in a way that reduces the cost of landing consumables on the lunar surface. Proposed concepts must include a method to move regolith through the reaction zone (e.g., regolith inlet/outlet valves capable of passing abrasive granular material through the valve for hundreds of cycles.) Near-pure metals or metal alloys are acceptable. Properties of metals extracted for manufacturing should be considered and provided.

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:
Level 1: TX 07 Exploration Destination Systems
Level 2: TX 07.1 In-Situ Resource Utilization

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware

Desired Deliverables Description:

Phase I efforts should provide a feasibility study and/or proof of concept. Phase II efforts should demonstrate the technology using lunar regolith simulant where applicable. Phase II efforts should be demonstrated at no less than 1/10 of the pilot plant production rate where applicable and should describe how the process can be applied at the full-scale production rate.

State of the Art and Critical Gaps:

The Moon to Mars Oxygen and Steel Technology (MMOST) SBIR Phase II sequential project is currently implementing size sorting and beneficiation of minerals containing iron at a relevant scale and is also producing iron as the main product. There has been little advancement toward the production of other metals such as aluminum. The Aqua Factorem project funded through the NASA Innovative Advanced Concepts (NIAC) program represents the state of the art for ice beneficiation.

Relevance / Science Traceability:

NASA has referenced water ice as one of the reasons we have chosen the lunar poles as the location to establish a sustained human presence. The Space Technology Mission Directorate (STMD) has identified the need for water extraction technologies. The Science Mission Directorate (SMD) is currently funding the Volatiles Investigating Polar Exploration Rover (VIPER) mission to investigate lunar water ice.

References:

Trigwell, S., Captain, J., Weis, K., & Quinn, J. (2012). Electrostatic beneficiation of lunar regolith: Applications in in situ resource utilization. Journal of Aerospace Engineering, 26(1), 30-36. https://ntrs.nasa.gov/api/citations/20110016173/downloads/20110016173.pdf
Quinn, J. W., Captain, J. G., Weis, K., Santiago-Maldonado, E., & Trigwell, S. (2013). Evaluation of tribocharged electrostatic beneficiation of lunar simulant in lunar gravity. Journal of Aerospace Engineering, 26(1), 37-42. https://ntrs.nasa.gov/api/citations/20110016172/downloads/20110016172.pdf
Schwandt, C., Hamilton, J. A., Fray, D. J., & Crawford, I. A. (2012). The production of oxygen and metal from lunar regolith. Planetary and Space Science, 74(1), 49-56.

Volatiles Investigating Polar Exploration Rover (VIPER), https://www.nasa.gov/viper

The SBIR focus area of Entry, Descent and Landing (EDL) includes the suite of technologies for atmospheric entry as well as descent and landing on both atmospheric and non-atmospheric bodies. EDL mission segments are used in both robotic planetary science missions and human exploration missions beyond Low Earth Orbit, and many technologies have application to emerging commercial space capabilities such as lunar landing, low-cost space access, small spacecraft, and asset return.

Robust, efficient, and predictable EDL systems fulfill the critical function of delivering payloads to lunar and planetary surfaces through challenging environments, within mass and cost constraints. Future NASA Artemis and planetary science missions will require new technologies to break through historical constraints on delivered mass, enable sustained human presence, or to go to entirely new planets and moons. Even where heritage systems exist, no two planetary missions are exactly “build-to-print,” leading to frequent challenges from environmental uncertainty, risk posture, and resource constraints that can be dramatically improved with investments in EDL technologies. EDL relies on validated models, ground tests, and sensor technologies for system development and certification. Both new capabilities and improved assessment and prediction of state-of-the-art systems are important facets of this focus area.

The subtopics in this Focus Area generally align with the Entry, Descent, and Landing flight regimes, including the flight instrumentation area. In future solicitations, the intent is to maintain these subtopic titles, and to rotate the content within the subtopics as Agency needs and priorities change and as technologies are matured.

The subtopics and their overarching content descriptions are:

Z7.01 Entry, Descent and Landing Flight Sensors and Instrumentation: Seeks sensors and components for precision landing and hazard detection, as well as heatshield instrumentation and other EDL flight systems diagnostics and electronics.
Z7.03 Entry and Descent Systems Technologies: Contains hypersonic materials, aeroshell systems, and modeling advances, including deployable aeroshells for EDL and asset return and recovery. Includes smaller-scale systems appropriate for small spacecraft applications.
Z7.04 Landing Systems Technologies: Covers landing engines, plume-surface interaction modeling, testing, and instrumentation, and landing attenuation systems.

Please refer to the subtopic write-ups for the specific content and scope solicited this year.

Lead Center: MSFC

Participating Center(s): AFRC, JSC, LaRC

Scope Title: Hot Structure Technology for Aerospace Vehicles

Scope Description:

This subtopic deals with the development of hot structure technology for aerospace vehicle structural components that are exposed to extreme heating environments. The hot structure technologies proposed for development must be for reusable, nonmetallic, oxidation-resistant, fiber-reinforced composite structures. Hot structure is an enabling technology for reusability, thus facilitating the development of advanced propulsion systems requiring multiple engine firings and vehicles requiring aerocapture/aerobraking followed by entry, descent, and landing. The development of hot structure technology for (a) combustion-device liquid rocket engine propulsion systems and (b) aerodynamic structures for aeroshells, control surfaces, wing leading edges, and heatshields is of great interest.

Desired hot structure systems encompass multifunctional structures that can reduce or eliminate the need for active cooling, and in the case of aerodynamic structures, separate thermal protection system (TPS) materials. The potential advantages of using hot structure systems in place of actively cooled structures or a TPS with underlying cool structure include reduced mass, increased mission performance (such as reusability and greater thermal efficiency), improved aerodynamics for aeroshell components, improved structural efficiency, and increased ability for nondestructive inspections. These aerospace vehicle applications are unique in requiring the hot structure to carry primary structure vehicle loads and to be reusable after exposure to extreme temperatures during liquid rocket engine firings and/or atmospheric entry. Examples of prior flight-proven hot structures include: (a) the composite nozzle extensions for the Centaur RL10 family of upper-stage rocket engines, and (b) the wing leading edges and control surfaces for the Space Shuttle Orbiter, Hyper-X (X-43A), and/or X-37B.

This subtopic seeks to develop innovative, low-cost, damage-tolerant, reusable, lightweight fiber-reinforced composite hot structure technology adhering to the following:

At a minimum, the subject hot structures must be capable of operating at temperatures of at least 1,510 °C (2,750 °F)—higher temperatures are of even greater interest, such as up to 2,204+ °C (4,000+ °F).
Constructed from composite fiber-reinforced materials, such as carbon-carbon (C-C) and ceramic matrix composite (CMC) materials.
Potential applications of interest for hot structure technology include: (a) propulsion system components (hot gas valves, combustion chambers, nozzles, and nozzle extensions) and (b) primary load-carrying aeroshell structures, control surfaces, leading edges, and heatshields.

Proposals should present approaches to address the current need for improvements in operating temperature capability, toughness/durability, reusability, and material system properties, as well as the need to reduce cost and manufacturing time requirements. Technology focus areas for submitted proposals should address one or more of the following:

Repeatable materials properties: Improvements in manufacturing processes and/or material designs to achieve repeatable uniform material properties, while minimizing data scatter, that are representative of actual vehicle components: specifically, material property data obtained from flat-panel test coupons should correlate directly to the properties of prototype and flight test articles.
Improved toughness/durability: Material/structural architectures and multifunctional systems providing significant toughness and/or durability improvements over typical 2D interlaminar mechanical properties while maintaining in-plane and thermal properties when compared to state-of-the-art C-C or CMC materials. Examples include incorporating through-the-thickness stitching, braiding, or 3D woven preforms. Advancements in oxidation resistance that enhance durability are also of interest, and may include matrix inhibition, oxidation resistant matrices, functionally graded material systems, and exterior environmental coatings. The goals here are to eliminate/reduce discontinuities in material properties and to provide robust material systems.
Reduced cost and/or delivery time: Manufacturing process methods that enable a significant reduction in the cost and time required to fabricate materials and components. There is a great need to reduce cost and processing time for hot structure materials and components—current state-of-the-art materials are typically expensive and have fabrication times often in the range of 6 to 12 months (or longer), which can limit or exclude the use of such materials. Approaches enabling reduced costs and manufacturing times should not lead, however, to significant reductions in material properties. Advanced manufacturing methods may include but are not limited to the following: (a) rapid densification cycles, (b) high char-yield resins, (c) additive manufacturing (AM), and (d) automated weaving, braiding, layup, etc.

Expected TRL or TRL Range at completion of the Project: 2 to 4

Primary Technology Taxonomy:
Level 1: TX 12 Materials, Structures, Mechanical Systems, and Manufacturing
Level 2: TX 12.1 Materials

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware
Research
Analysis

Desired Deliverables Description:

Research, testing, and analysis should be conducted to demonstrate technical feasibility during Phase I and show a path towards Phase II hardware or prototype demonstrations. Phase I feasibility studies should also address cost and the risks associated with the hot structure technology.

In addition to delivery of a Phase I final report, a representative sample(s) of the material and/or technology addressed by the Phase I project should be provided at the conclusion of the Phase I contract. Examples of representative Phase I sample deliverables include:

Coupons appropriate for thermal and/or mechanical material property tests.
Arc-jet test specimens.
Subelement or subcomponent structures.

Plans for potential follow-on Phase II contracts should include the delivery of manufacturing demonstration units to NASA or a commercial space industry partner during Phase II. Testing of such demonstration articles should be a part of the anticipated Phase II effort. Depending upon the primary application addressed by the Phase II contract, such test articles may include subscale nozzle-extensions, arc-jet specimens, or other representative hot structure components. Opportunities and plans should also be identified and summarized for potential commercialization with at least one aerospace company. Vehicle integration issues (attachment, joining, etc.) should be addressed.

State of the Art and Critical Gaps:

The current state of the art for composite hot structure components is limited primarily to applications with maximum use temperatures in the 1,093 to 1,593 °C (2,000 to 2,900 °F) range. While short excursions to higher temperatures are possible, considerable degradation may occur. Reusability is limited and may require considerable inspection before potential reuse. Critical gaps or technology needs include:

Increasing operating temperatures to 1,649 to 2,204+ °C (3,000 to 4,000+ °F).
Increasing resistance to environmental attack (primarily through oxidation).
Increasing manufacturing technology capabilities to improve reliability, repeatability, and quality control.
Increasing durability/toughness and interlaminar mechanical properties (for 2D reinforcement) or introducing 3D architectures.
Decreasing manufacturing cost.
Decreasing overall manufacturing time requirements.

Relevance / Science Traceability:

Hot structure technology is relevant to the Human Exploration and Operations Mission Directorate (HEOMD), where the technology can be infused into spacecraft and launch vehicle applications. Such technology should provide either improved performance or enable advanced missions requiring reusability, increased damage tolerance, and the durability to withstand long-duration space exploration missions. The ability to allow for delivery and/or return of larger payloads (and crewed vehicles) to various space destinations, such as the lunar South Pole and Mars, is also of great interest.

The Advanced Exploration Systems (AES) Program (https://www.nasa.gov/directorates/heo/aes/index.html) would be ideal for further funding a prototype hot structure system and technology demonstration effort. Commercial space programs, such as the Commercial Resupply Services (CRS) Program, the Commercial Crew Program (CCP), the Commercial Lunar Payload Services (CLPS) Program, and Next Space Technologies for Exploration Partnerships (NextSTEP), are also interested in this technology for flight vehicles. Additionally, NASA HEOMD programs that could use this technology for propulsion upgrades or block changes in the future include the Artemis Space Launch System (SLS), Orion, and Human Landing System (HLS). Hot structure technology is also highly relevant to the NASA Aeronautics Research Mission Directorate’s (ARMD’s) Hypersonic Technology (HT) Project (https://www.nasa.gov/aeroresearch/programs/aavp/ht). Other relevant efforts include the work done by NASA and the Defense Advanced Research Projects Agency (DARPA) in developing nuclear thermal propulsion (NTP) systems, both for reactor materials and nozzle extensions.

Potential NASA users of this technology exist for a variety of propulsion systems and other applications requiring the use of similar materials, including the following:

Upper-stage engine systems, such as those for the Artemis SLS.
In-space propulsion systems, including nuclear thermal propulsion systems.
Lunar/Mars lander descent/ascent propulsion systems.
Propulsion systems for the commercial space industry, which is partnering with and supporting NASA efforts.
Atmospheric entry vehicle aeroshells, such as those for use at Earth, Mars, or other planets and their moons.
Related applications include the structures required for hypersonic flight vehicles.

Finally, the U.S. Air Force is interested in such technology for its National Security Space Launch (NSSL), ballistic missile, and hypersonic vehicle programs. Other non-NASA users include the U.S. Army, the U.S. Navy, the U.S. Space Force, the Missile Defense Agency (MDA), and DARPA. The subject technology can be both enhancing to systems already in use or under development, as well as enabling for applications that may not be feasible without further advancements in high-temperature composites technology.

References:

Liquid Rocket Propulsion Systems:

“Carbon-Carbon Nozzle Extension Development in Support of In-Space and Upper-Stage Liquid Rocket Engines;” Paul R. Gradl and Peter G. Valentine; 53^rd AIAA/SAE/ASEE Joint Propulsion Conference, Atlanta, GA; AIAA-2017-5064; July 2017;https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170008949.pdf.
“Extreme-Temperature Carbon- and Ceramic-Matrix Composite Nozzle Extensions for Liquid Rocket Engines;” Peter G. Valentine and Paul R. Gradl; 70^th International Astronautical Congress (IAC), Washington DC; IAC-19-C2.4.9; October 2019; https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20190033315.pdf.

Hypersonic Flight Vehicle Structures:

"Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles;" David E. Glass; 15^th AIAA International Space Planes and Hypersonic Systems and Technologies Conference, Dayton, OH; AIAA-2008-2682; April-May 2008;https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080017096.pdf.
"A Multifunctional Hot Structure Heatshield Concept for Planetary Entry;" Sandra P. Walker, Kamran Daryabeigi, Jamshid A. Samareh, Robert Wagner, and Allen Waters; 20^th AIAA International Space Planes and Hypersonic Systems and Technologies Conference, Glasgow, Scotland; AIAA 2015-3530; July 2015; https://arc.aiaa.org/doi/pdf/10.2514/6.2015-3530.

Note: The above references are all open literature references. Other references exist regarding this technology, but they are International Traffic in Arms Regulations (ITAR) restricted. Numerous online references exist for the subject technology and projects/applications presented here, both foreign and domestic.

Lead Center: JSC

Participating Center(s): ARC, GSFC, JPL, LaRC

Scope Title: Air Data Sensors to Support Entry, Descent, and Landing (EDL) Environment Characterization

Scope Description:

Current NASA state-of-the-art air data sensors for EDL applications are very expensive to incorporate on planetary missions because they must meet functional and performance requirements during and after exposure to loads and environments associated with long-duration spaceflight and atmospheric entry. The dynamic loads and thermal environments encountered prior to arrival at the destination make flight qualification of air data sensors challenging and costly. Scarce commercial options exist for off-the-shelf products with the potential to meet NASA’s requirements for accuracy and survivability. To bring more commercial options for air data sensors that can be flown on EDL missions as part of an air data system, NASA seeks proposals in two distinct areas: pressure transducers and lidar sensors.

1. Air Data Pressure Transducers

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suites flew supersonic-range pressure transducers on the heat shield that were developed in house at NASA Langley Research Center because there was no commercially available pressure sensor that met the mission’s requirements. The hypersonic pressure transducer on the heat shield was a flight spare from the first MEDLI suite, which flew on the Mars Science Laboratory Mission in 2012. In situ pressure measurements on the aerodynamic surfaces of an EDL vehicle capsule—such as the heat shield and backshell—are primarily used to reconstruct the free-stream density and vehicle attitude (angles of attack and sideslip) to isolate aerodynamic performance. NASA seeks pressure transducers that can meet the following requirements:

Configuration: The pressure transducer shall be hermetically sealed. The design space should consider a nonamplified output configuration or a configuration with embedded electronics for an amplified output. The internal temperature shall be monitored. The pressure transducer shall measure absolute pressure, and the housing must be able to be connected to a flared tube fitting. The pressure transducer should have the capability of being mechanically mounted in a 2- or 3-point configuration.
Mass: Less than 300 g if no active electronics; less than 400 g for a unit with active electronics.
Size: Less than 442 c³.
Electrical connections: Electrical interface/connector should be configured for power, ground, analog signal, analog return, and temperature sensor accommodation. Electrical connector pin configurations should allow for interchangeability of mating connectors for all pressure transducers.
Parts, material, and processes used in the construction of the pressure transducer should be controlled by specification or procedure per AS9100 or equivalent. Any soldering should meet NASA-STD-8739.3 or IPC J-STD-001 with space addendum, and any fusion welding should follow AWSD17.1.
The pressure transducer should meet MIL-STD-461 for electromagnetic interference (EMI) compliance (amplified units only).
Axial loading capability: Minimum 15 g (Venus missions could require 100 g or higher).
Temperature capability: Operating temperature range of -120 to 80 °C. It is desired that the unit can survive temperatures as cold as -130 °C in a nonoperating condition. It is also desired that the unit can survive a dry heat microbial reduction temperature of 104 °C for 200 hr, or 110 °C for 100 hr.
Functional characteristics:
- Input voltage: Up to 10 Vdc for a nonamplified unit or 12 to 36 Vdc for an amplified unit.
- Input current: Should not exceed 7 mA for a nonamplified unit or 30 mA for an amplified unit.
- Output impedance: Should not exceed 10 kilohms for a nonamplified unit or 1 kilohm for an amplified unit.
- Output voltage: Minimum output of 1.2 mV/V (nonamplified unit).
- Measurement range: 0 to 1.0 psia for a supersonic range pressure transducer; 0 to 5.0 psia for hypersonic range pressure transducer. Zero psia is considered to be less than 10^-5 torr.
- Accuracy: Provide a description of the approach to quantify and demonstrate accuracy of the pressure transducer.
- Static error band should be no greater than +/- 0.3% of full scale based on an unweighted least-squares straight-line fit. The static error band includes errors due to nonlinearity, hysteresis, and non-repeatability.
Cost: Fully qualified first-unit target of ~$500K.

2. Air Data Lidar Sensors

Air data lidar sensors have the potential of providing more accurate velocimetry data than pitot tubes for Mars landing and Earth reentry vehicles. Furthermore, a lidar-based air data sensor can eliminate the aerodynamic influences of pitot tubes, particularly in the supersonic velocity regime. NASA seeks proposals for air data lidar sensors that can provide critical air-vector velocity data during the atmospheric entry and descent phases of the spacecraft. An ability to provide other relevant air data, such as atmospheric pressure, is viewed favorably if it can enhance and/or complement the air velocity measurement capabilities. The proposed lidar sensor must be compact and efficient with a clear path to spaceflight units meeting physical and environmental constraints of landing vehicles.

Expected TRL or TRL Range at completion of the Project: 2 to 4

Primary Technology Taxonomy:
Level 1: TX 09 Entry, Descent, and Landing
Level 2: TX 09.X Other Entry, Descent, and Landing

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware

Desired Deliverables Description:

Phase I Goals: Design and proof of concept, including the production approach to achieve the cost goals.

Phase II Goals: Prototype/breadboard validation in laboratory environment.

State of the Art and Critical Gaps:

NASA now requires instrumentation on all EDL missions, including competed science missions, and these cost- and mass-constrained missions cannot use the state-of-the-art instrumentation. Very few commercial options exist for air data sensors that can meet accuracy and survivability requirements.

Relevance / Science Traceability:

EDL instrumentation directly informs and addresses the large performance uncertainties that drive the design, validation, and in-flight performance of planetary entry systems. Improved understanding of entry environments and real-time measurement knowledge could lead to reduced design margins, enabling a greater payload mass-fraction, and smaller landing ellipses for placing advanced payloads onto the surface of atmospheric and airless bodies.

References:

H. Hwang, et al. (2016), "Mars 2020 Entry, Descent and Landing Instrumentation 2 (MEDLI2)," 46th AIAA Thermophysics Conference, Washington, D.C., June 2016. AIAA paper No. AIAA 2016-3536.
J. Santos, K. Edquist, H. Hwang, et al. (2020), "Entry, Descent, and Landing Instrumentation," White Paper for the Planetary Sciences Decadal Survey, 2023-2032. September 2020.

Scope Title: Novel Lidar Component Technologies Applicable to Guidance, Navigation, and Control (GN&C) for Precise Safe Landing

Scope Description:

NASA is seeking the development of component technologies for advanced lidar sensors that will be utilized within entry, descent, and landing (EDL) and deorbit, descent, and landing (DDL) GN&C systems for precise safe landing on solid solar system bodies, including planets, moons, and small celestial bodies (e.g., asteroids and comets). The EDL phase applies to landings on bodies with atmospheres, whereas DDL applies to landings on airless bodies. For many of these missions, EDL/DDL represents one of the riskiest flight phases. NASA has been developing technologies for precision landing and hazard avoidance (PL&HA) to minimize the risk of the EDL/DDL phase of a mission and to increase the accessibility of surface science targets through precise and safe landing capabilities. One flight instrumentation focus of PL&HA technology has been in the development of lidar technologies that provide either terrain mapping (range point cloud) capability or direct velocity measurement. The continued maturation of these technologies is targeting (1) multimodal operation (i.e., combining mapping and velocimetry functions); (2) reduction of size, mass, and power; and (3) multicomponent integration.

This solicitation is requesting specific lidar system components and not complete lidar solutions. To be considered, all component technologies proposed must show a development path to operation within the applicable EDL/DDL spaceflight environment (radiation, thermal, vacuum, vibration, etc.). The specific lidar component technologies desired include the following (proposals can be to either or both):

Dense focal plane arrays for simultaneous ranging and Doppler velocimetry with the following characteristics:

Simultaneous measurements from each pixel or from subsets of pixels.
Functionality (when integrated into a lidar system) for measuring range up to 8 km.
- Range precision less than 5 cm, 1-sigma, for 3D image frames up to 1 km.
- Range precision less than 1 m, 1-sigma, for ranges up to 8 km.
Functionality (when integrated into a lidar system) for measuring velocity from 0 to 200 m/sec (or greater) along the line of sight (LOS).
- Doppler velocity precision on order of 1 cm/sec, 1-sigma, from ranges of 4 km or greater.
Rejection of false locks on dust or plumes from the spacecraft exhaust.
Implementation for low power, mass, and size.

Readout integrated circuit (ROIC) consisting of preamplifiers and switching fabric, capable of operating at cryogenic temperatures, with the following characteristics:

Preamplifiers: Array of low-noise transimpedance preamplifiers, one for each detector element.
- Electrical bandwidth: >150 MHz.
- Transimpedance gain: >300 kV/A.
- Input current noise: <1.5 pA/Hz^1/2.
- Input voltage noise: <10 nV/Hz^1/2.
- Output: Analog pulse waveforms, DC coupling.
- Electrical power: <2 mW per element.
Network switching fabric: Connection of a subarray of the detector elements to the output terminals.
- Switch speed: >5 MHz with settling time <40 ns.
- Number of input channels: Up to 2x320.
- Number of output channels: Subarray of the input signals up to 16 channels.
- Interchannel isolation: < -37 dB @ 1 GHz.
- Insertion loss: <1 dB.
- Total electrical power: <0.05 W.

Expected TRL or TRL Range at completion of the Project: 4 to 6

Primary Technology Taxonomy:
Level 1: TX 09 Entry, Descent, and Landing
Level 2: TX 09.X Other Entry, Descent, and Landing

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

The following deliverables are desired for Phase I: (1) Hardware demonstrations of sensor components and applicable support hardware and/or (2) Analysis and software simulations of component proofs of concept within simulated environments. Responses must show a path for the proposed capabilities to be compatible with the environmental conditions of spaceflight.

The following deliverables are desired for Phase II: (1) Hardware demonstrations of sensor components and applicable support hardware and (2) Analysis of components in laboratory or relevant environment (depending on TRL). Phase II products will need to demonstrate a path for the capabilities to be compatible with the environmental conditions of spaceflight.

State of the Art and Critical Gaps:

For more than a decade, the EDL GN&C and sensors community has been developing the technologies to enable precise safe landing. Infusion of these capabilities into spaceflight missions and spinoff into the commercial sector remains the critical gap. Bridging this gap requires additional component technology advancements for specific lidar sensors that enhance operational performance, increase dynamic envelope, reduce size/mass/power/cost, and enable spaceflight qualification.

Relevance / Science Traceability:

GN&C/PL&HA technologies for precise safe landing are critical for future robotic science and human exploration missions to locations with hazardous terrain and/or pre-positioned surface assets (e.g., cached samples or cargo) that pose significant risks to successful spacecraft touchdown and mission surface operations. The PL&HA technologies enable spacecraft to land with minimum position error from targeted surface locations, and they implement hazard-avoidance diverts to land at locations safe from lander-sized or larger terrain hazards (e.g., craters, rocks, boulders, sharp slopes, etc.). PL&HA has maintained consistent prioritization within the NASA and National Research Council (NRC) space technology roadmaps for more than a decade, and multiple planetary landers such as Mars 2020 and upcoming Commercial Lunar Payload Services (CLPS) are starting to infuse some of the PL&HA capabilities.

References:

A. Martin, et al. (2018), "Photonic integrated circuit-based FMCW coherent LiDAR," in Journal of Lightwave Technology, vol. 36, no. 19, 4640-4645, Oct.1, 2018, doi: 10.1109/JLT.2018.2840223.
C.V. Poulton, A. Yaacobi, D.B. Cole, M.J. Byrd, M. Raval, D. Vermeulen, and M.R. Watts (2017), "Coherent solid-state LIDAR with silicon photonic optical phased arrays," Opt. Lett. 42, 4091-4094.
F. Amzajerdian, G.D. Hines, D.F. Pierrottet, B.W. Barnes, L.B. Petway, and J.M. Carson (2017), “Demonstration of coherent Doppler lidar for navigation in GPS-denied environments,” Proc. SPIE 10191, Laser Radar Technology and Applications XXII, 1019102.
X. Sun, J.B. Abshire, J.D. Beck, P. Mitra, K. Reiff, and G. Yang (2017), “HgCdTe avalanche photodiode detectors for airborne and spaceborne lidar at infrared wavelengths,” Optics Express, 25: 16589-16602.
X. Sun, J. Abshire, M. Krainak, W. Lu, J. Beck, W. Sullivan III, P. Mitra, D. Rawlings, R. Fields, D. Hinkley, B. Hirasuna (2019), “HgCdTe avalanche photodiode array detectors with single photon sensitivity and integrated detector cooler assemblies for space applications,” Optical Engineering, 58, pp. 067103.

W. Sullivan III, M. Goodwin, J. Beck, C. Kmilar, D. Rawlings, M. Skokan, P. Mitra (2018), “A 5 MHz frame rate 32x30 HgCdTe APD focal plane array with photon counting sensitivity for infrared laser radar technology,” Military Sensing Symposium (MSS), March 20, 2018 (unclassified, U.S. citizen only).

Lead Center: LaRC

Participating Center(s): ARC

Scope Title: Entry and Descent System Technologies

Scope Description:

NASA is advancing deployable aerodynamic decelerators to enhance and enable robotic and human space missions. Applications include Mars, Venus, and Titan as well as payload return to Earth from low Earth orbit. The benefit of deployable decelerators is that the entry vehicle structure and thermal protection system are not constrained by the launch vehicle shroud. Deployable decelerators have the flexibility to more efficiently use the available shroud volume and can be packed into a much smaller volume for Earth departure, addressing potential constraints for payloads sharing a launch vehicle. For Mars, this technology enables delivery of a very large (20 metric tons or more) usable payload, which may be needed to support human exploration. The technology also allows for reduced-cost access to space by enabling the recovery of launch vehicle assets. Development of efficient gas generator technology is needed for inflation of large inflatable decelerators. NASA is also seeking development of domestic capability for fabricating custom stretch-broken carbon and polymer blended yarns for traditional thermal protection systems for other planetary entry missions. This subtopic area solicits innovative technology solutions applicable to both deployable and traditional entry concepts. Specific technology development areas include (1) gas generators for hypersonic inflatable aerodynamic decelerators (HIAD) and (2) blended phenolic/carbon yarn for 3D woven ablative thermal protection systems.

1. Gas Generators for HIAD

Development of gas generator technologies used as inflation systems that result in improved mass efficiency and system complexity over current pressurized cold gas systems for inflatable structures is desired. Inflation gas technologies can include warm or hot gas generators, sublimating powder systems, or hybrid systems; however, the final delivery gas temperature must not exceed 200 °C. Lightweight, high-efficiency gas inflation technologies capable of delivering gas at 250 to 10,000 standard liters per minute (SLPM) are sought. This range spans a number of potential applications. Thus, a given response need not address the entire range. Additionally, the final delivery gas and its byproducts must not harm aeroshell materials such as the fluoropolymer liner of the inflatable structure. Minimal solid particulate is acceptable as a final byproduct. Water vapor as a final byproduct is also acceptable for lower flow (250 to 4,000 SLPM) and shorter duration missions, but it is undesirable for higher flow (8,000 to 10,000 SLPM) and longer duration missions. Chillers and/or filters can be included in a proposed solution, but they will be included in assessing overall system mass versus amount of gas generated. Gas delivery configurations that rely on active flow control devices are not desired. Long-term mission applications will have inflatable volumes in the range of 1,200 to 4,000 ft³ with final inflation pressures in the range of 15 to 30 psid. Initial concepts will be demonstrated with small-scale volumes to achieve the desired inflation pressures and temperatures. Focus of Phase I development can be subscale manufacturing demonstrations that demonstrate proof of concept and lead to Phase II manufacturing scaleup for applications related to human-scale Mars entry, Earth return, launch vehicle asset recovery, or the emergent small-satellite community.

2. Blended Phenolic/Carbon Yarn for 3D Woven Ablative Thermal Protection Systems

Development of domestic capability for fabricating custom stretch-broken carbon and polymer blended yarns is desired. Specifically, NASA is interested in the ability to twist and ply stretch-broken fibers into a 4-ply blended yarn of varying carbon/phenolic/thermoset resin ratios (phenolic or other nonbrittle fibers preferred). Challenges include maintaining an intimate blend ratio to maintain consistent linear weight while also fabricating a high-quality yarn free from breaks and large yarn defects (e.g., slubs and flames in the resin phase), with uniformity in diameter such that yarns are capable of being processed into 3D woven preforms for advanced thermal protection systems. Phase I effort shall identify the ability to fabricate these custom yarns and establish the characterization processes and controls that will be necessary to eventually fine-tune the blended yarn properties. Final composition of interest to NASA would be a carbon/phenolic blended yarn—any surrogate polymeric yarn should have similar stretch-breaking and blending performance such that any successful process shown with surrogate yarn is extensible to a carbon/phenolic blend with low risk. Notional Phase II effort would demonstrate blending of stretch-broken carbon/kynol fibers and detailed yarn testing—char, strength, yield, etc.—to meet the following established NASA specifications:

Carbon to phenolic ratio in the yarn by mass shall be 63 ± 4% carbon to 37 ± 4% phenolic
Blended yield shall be 1,140 yd/lb +/- 10%.
Yarn shall have a minimum strength of >13,000 cN and elongation of >1%.
Yarn shall have a twist in the “S” direction and shall be 115 +/- 15% T/m (twists per meter) (2.92 T/in.).
Yarn shall be manufactured so as to reduce presence of surface features such as slubs or flames.

Expected TRL or TRL Range at completion of the Project: 1 to 4

Primary Technology Taxonomy:
Level 1: TX 09 Entry, Descent, and Landing
Level 2: TX 09.1 Aeroassist and Atmospheric Entry

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Reports documenting analysis and development results, including description of any hardware or prototypes developed. Focus of Phase I development can be material coupons and/or subscale manufacturing demonstrations that demonstrate proof of concept and lead to Phase II scaleup and testing in relevant environments for applications related to Mars and other planetary entry, Earth return, launch asset recovery, or the emergent small-satellite community.

State of the Art and Critical Gaps:

The current state of the art for deployable aerodynamic decelerators is limited due to novelty of this technology. Development of gas generator technologies that improve mass efficiency over current pressurized cold gas systems for inflatable structures is needed. Domestic capability for producing blended phenolic/carbon yarn for 3D woven thermal protection systems is nonexistent, and NASA is interested in developing this domestic capability for future missions.

Relevance / Science Traceability:

NASA needs advanced deployable aerodynamic decelerators to enhance and enable robotic and human space missions. Applications include Mars, Venus, and Titan as well as payload return to Earth from low Earth orbit. NASA also needs domestic supply of blended phenolic/carbon yarn for 3D woven traditional thermal protection systems. HEOMD (Human Exploration and Operations Mission Directorate), STMD (Space Technology Mission Directorate), and SMD (Science Mission Directorate) can benefit from this technology for various exploration missions.

References:

Hughes, S. J., et al., “Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview,” AIAA Paper 2011-2524.
Bose, D. M, et al., “The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Mission Applications Study,” AIAA Paper 2013-1389.
Hollis, B. R., “Boundary-Layer Transition and Surface Heating Measurements on a Hypersonic Inflatable Aerodynamic Decelerator with Simulated Flexible TPS,” AIAA Paper 2017-3122.
Olds, A. D., et al., “IRVE-3 Post-Flight Reconstruction,” AIAA Paper 2013-1390.
Del Corso, J. A., et al., “Advanced High-Temperature Flexible TPS for Inflatable Aerodynamic Decelerators,” AIAA Paper 2011-2510.
Cassell, A., et al., “ADEPT, A Mechanically Deployable Re-Entry Vehicle System, Enabling Interplanetary CubeSat and Small Satellite Missions,” SSC18-XII-08, 32nd Annual AIAA/USU Conference on Small Satellites.
Cassell, A., et al., “ADEPT Sounding Rocket One Flight Test Overview,” AIAA Paper 2019-2896.

Ellerby, D., et al., “Heatshield for Extreme Entry Environment Technology (HEEET) Thermal Protection System (TPS),” Materials Science and Technology (MS&T) 2019, September 29-October 3, 2019, Portland, Oregon.

Lead Center: MSFC

Participating Center(s): GRC, LaRC

Scope Title: Plume-Surface Interaction (PSI) Instrumentation, Ground Testing, and Analysis

Scope Description:

As NASA and commercial entities prepare to land robotic and crewed vehicles on the Moon, and eventually Mars, characterization of landing environments is critical to identifying requirements for landing systems and engine configurations, instrument placement and protection, and landing stability. The ability to predict the extent to which regolith is liberated and transported in the vicinity of the lander is also critical to understanding the effects on precision landing sensor requirements and landed assets located in close proximity. Knowledge of the characteristics, behavior, and trajectories of ejected particles and surface erosion during the landing phase is important for designing effective sensor systems and PSI risk mitigation approaches. Mission needs to consider include landers with single and multiple engines, both pulsed and throttled systems, landed mass from 400 to 40,000 kg, and both lunar and Mars destinations.

NASA is seeking support in the following areas:

Ground test data, test techniques, and diagnostics across physical scales and environments, with particular emphasis on nonintrusive approaches and methodologies.
PSI-specific flight instrumentation, with particular emphasis on in situ measurements of particle size and particle velocity during the landing phase.
Solutions to alleviate or mitigate the PSI environments experienced by propulsive landers—not vehicle-specific solutions.
Validated, robust, and massively parallel computational fluid dynamics (CFD) models and tools for predicting PSI physics for plumes in low-pressure and rarefied environments, time-evolving cratering and surface erosion, and near-field and far-field ejecta transport.

NASA has plans to purchase services for payload delivery to the Moon through the Commercial Lunar Payload Services (CLPS) contract. Under this subtopic, proposals may include efforts to develop payloads for flight demonstration of relevant PSI technologies in the lunar environment. The CLPS payload accommodations will vary depending on the particular service provider and mission characteristics, but the data to be obtained or mitigations to be demonstrated should be broadly applicable to other future landing systems. Additional information on the CLPS program and providers can be found at this link: https://www.nasa.gov/content/commercial-lunar-payload-services. CLPS missions will typically carry multiple payloads for multiple customers. Smaller, simpler, and more self-sufficient payloads are more easily accommodated and would be more likely to be considered for a NASA-sponsored flight opportunity. Commercial payload delivery services are currently under contract, and flight opportunities are expected to continue well into the future. In future years, it is expected that larger and more complex payloads will be accommodated. Selection for award under this solicitation will not guarantee selection for a lunar flight opportunity.

Expected TRL or TRL Range at completion of the Project: 3 to 6

Primary Technology Taxonomy:
Level 1: TX 09 Entry, Descent, and Landing
Level 2: TX 09.3 Landing

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Deliverables of all types can be infused into the prospect missions due to early design maturity.

For PSI ground test data, flight instrumentation, diagnostics, and mitigation approaches, Phase I deliverables should include detailed test plans, with prototype and/or component demonstrations as appropriate. Phase II deliverables should include complete data products, fully functional hardware, and validated performance in relevant environments.

For PSI modeling and simulation, Phase I deliverables should demonstrate proof of concept and a minimum of component-level verification, with detailed documentation on future data needs to complete validation of the integrated model and uncertainty quantification methodology. Phase II deliverables must demonstrate verification and validation beyond the component level, with validation demonstrated through comparisons with relevant data and documented uncertainty quantification. Significant attention should be applied to create highly robust and extremely high-performance computational simulation tool deliverables, exploiting leading-edge computational architectures to achieve this performance.

State of the Art and Critical Gaps:

The characteristics and behavior of airborne particles during descent is important for designing descent sensor systems that will be effective. Furthermore, although the physics of the atmosphere and the characteristics of the regolith are different for the Moon, the capability to model PSIs on the Moon will feed forward to Mars, where it is critical for human exploration.

Currently, flight data are collected from early planetary landing, and those data are fed into developmental tools for validation purposes. The validation data set, as well as the expertise, grows as a result of each mission and is shared across and applied to all other missions. We gain an understanding of how various parameters, including different types of surfaces, lead to different cratering effects and plume behaviors. The information helps NASA and industry make lander design and operations decisions. Ground testing (“unit tests”) is used early in the development of the capability in order to provide data for tool validation.

The current postlanding analysis of planetary landers (on Mars) is performed in a cursory manner with only partially empirically validated tools, because there has been no dedicated fundamental research investment in this area. Flight test data does not exist in the environments of interest.

Relevance / Science Traceability:

Current and future lander architectures will depend on knowledge of PSI, such as:

Artemis human landing system (HLS).
Commercial robotic lunar landers (CLPS or other).
Planetary mission landers (Mars Sample Retrieval Lander and others).
Human Mars landers.

References:

Lander Technologies: https://www.nasa.gov/content/lander-technologies
Metzger, Philip, et al. (2009). ISRU implications for lunar and Martian plume effects. 47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition.
Plemmons, D. H., et al. (2008). Effects of the Phoenix Lander descent thruster plume on the Martian surface. Journal of Geophysical Research: Planets, 113(E3).
Mehta, M., et al. (2013). Thruster plume surface interactions: Applications for spacecraft landings on planetary bodies. AIAA Journal, 51(12), 2800-2818.
Vangen, Scott, et al. (2016). International Space Exploration Coordination Group Assessment of Technology Gaps for Dust Mitigation for the Global Exploration Roadmap. AIAA SPACE 2016. 5423.

Scope Title: Landing Shock Attenuation, Reusability, and In Situ Landing Sensors

Scope Description:

Novel and creative solutions will be required to attenuate the structural loads induced by the landing of crewed spacecraft, commercial cargo payloads, scientific payloads, critical surface assets, and surface habitats on the Moon and Mars. In principle, the mass and scale of these spacecraft, payloads, and assets could range from something akin to a small-satellite class, roughly 10 to 500 kg, to masses on the order of thousands of kg. This capability is critical for landing larger spacecraft near assets already in place.

Current landing system solutions include legs, shock absorbers, inflatables, crushables, sky cranes, pallets, etc., but new technologies, novel combinations of existing technologies, and/or the repurposing of current Earth-based technologies could enable new mission design and feasibility.

Mission concepts requiring the sustainability and reusability of assets and payloads on the surfaces of celestial bodies (including the Moon, Mars, moons of Mars, comets, and/or asteroids) will benefit from the development of reusable landing systems, including consideration of launch plumes for ascent vehicles. Reusability can also be interpreted to include the postlanding adaptation of landing systems to enable mobility or augmented capabilities (e.g., "touch-and-go" mobility, grappling, maneuverability, etc.).

In situ landing sensors that measure the induced loads and shocks experienced within these challenging environments will provide engineers and researchers with valuable in situ data, which will enable improved environmental modeling, landing structure design, and sensor design. Possible applications include advanced touchdown sensors, measurement of payload orientation, stability, and/or landing loads.

Also, of interest are approaches for achieving multifunctional components, repurposing landing structures for postflight mission needs such as payload placement or mobility, and incorporating design features that reduce operating complexity.

Under this subtopic, proposals may include efforts to develop prototypes for flight demonstration of relevant technologies in the lunar environment or in terrestrial testbeds. The Commercial Lunar Payload Services (CLPS) accommodations will vary depending on the particular service provider and mission characteristics. Additional information on the CLPS program and providers can be found at this link: https://www.nasa.gov/content/commercial-lunar-payload-services. CLPS missions will typically carry multiple payloads for multiple customers. Smaller, simpler, and more self-sufficient payloads are more easily accommodated and would be more likely to be considered for a NASA-sponsored flight opportunity. Commercial payload delivery services may begin as early as 2022, and flight opportunities are expected to continue well into the future. In future years, it is expected that larger and more complex payloads will be accommodated. Selection for award under this solicitation will not guarantee selection for a lunar flight opportunity.

Expected TRL or TRL Range at completion of the Project: 3 to 6

Primary Technology Taxonomy:
Level 1: TX 09 Entry, Descent, and Landing
Level 2: TX 09.3 Landing

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Deliverables and/or prototypes of all types can be infused into the prospective missions due to early design maturity.

Phase I deliverables should include preliminary designs, end-product test plans, and component-level testing and/or demonstrations as appropriate, and Phase II should include a working prototype demonstration in a relevant environment.

State of the Art and Critical Gaps:

Robust landing structures can enable lunar and Mars global access with 20-ton payloads to support human missions.

Mission risks related to hazard avoidance may be partially mitigated by robust landing system accommodation of landing hazards.

Development of exploration technologies to enable a vibrant space economy can be partially addressed with respect to landing technologies related to landing pads and protective and robust landing structures.

Construction and outfitting of assets on the Moon and Mars could be addressed by technologies related to multifunctional and adaptive landing structures for use after landing.

Relevance / Science Traceability:

Current and future lander architectures will depend on landing shock attenuation, reusability, and intelligent landing sensors, such as:

Artemis human landing system (HLS).
Commercial robotic lunar landers (CLPS or other).
Planetary mission landers (Mars Sample Retrieval Lander and others).
Human Mars landers.
Scientific investigations of comets and asteroids.

References:

Lander Technologies: https://www.nasa.gov/content/lander-technologies
Commercial Lunar Payload Services: https://www.nasa.gov/content/commercial-lunar-payload-services
Lunar Exploration and Transportation Services: https://www.nasa.gov/nextstep/humanlander3
JAXA Hyabusa2: https://www.hayabusa2.jaxa.jp/en/
JPL ATHLETE Rover: https://www-robotics.jpl.nasa.gov/systems/system.cfm?System=11
SLS-SPEC-159 Cross-Program Design Specification for Natural Environments (DSNE)

Lead Center: GSFC

Participating Center(s): JPL, JSC, LaRC, MSFC

Scope Title: Coatings for Lunar Regolith Dust Mitigation for Thermal Radiators and Extreme Environments

Scope Description:

Thermal coatings are an integral part of a space mission and are essential to the survivability of the spacecraft and instrument. Radiator surface coatings with desired emissivity and absorptivity provide a passive means for instrument temperature control. The utilization of variable-emittance devices further enables active control of the instrument temperature when the heat output from the instrument or the thermal environment of the radiator changes. With NASA’s new initiative to return to the Moon, a new coating technology that will keep surfaces clean and sanitary is needed. New coating formulations utilizing durable, anticontamination, and self-cleaning properties that will disallow the accumulation of dust, dirt, and foreign materials are highly desirable. These coatings can have low absorptance and high infrared (IR) emittance properties or be transparent for use on existing thermal coating systems. The goal of this technology is to preserve optimal long-term performance of spacecraft and habitation components and systems. Furthermore, coatings that can survive and operate in extreme environments (cryogenic or high temperature) are desirable.

Expected TRL or TRL Range at completion of the Project: 2 to 5

Primary Technology Taxonomy:
Level 1: TX 14 Thermal Management Systems
Level 2: TX 14.3 Thermal Protection Components and Systems

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware

Desired Deliverables Description:

Phase I Deliverables:

Successful development of coating formulations that lead to the desired dust mitigation.
Deliverable of coupon.
Samples of the hardware for further testing at NASA facilities.
Final report.

Phase II Deliverables:

Results of performance characterization tests.
Results of stability test of the coating formulations and their mechanical durability test under the influence of simulated space and lunar environmental conditions.
Test coupon.
Final report.

State of the Art and Critical Gaps:

There are limited options for durable, stable thermal control coatings that are dust shedding in charging environments. Current state-of-the-art, sprayable radiation-stable coatings are able to coat complex, irregular surfaces, but they are porous and will become imbedded with dust and particulates. Other surface films tend to be less optically stable and may charge in the plasma environment, thereby attracting lunar regolith to their surfaces. Mirrors have the limitations of requiring flat surfaces and are not conformal in nature. Currently, no single thermal control surface appears to provide stability, durability, and meet optical property requirements for sustained durations in space and lunar environments.

Relevance / Science Traceability:

Many Science Mission Directorate (SMD) missions will greatly benefit from this dust mitigation thermal coating technology: any lunar-related project and projects involved with robotic science rovers and landers.

References:

References for dust mitigation coatings such as lotus thermal coatings: https://ntrs.nasa.gov/search.jsp?R=20150020486
References for extreme environment coatings: https://vfm.jpl.nasa.gov/files/EE-Report_FINAL.pdf
References in Subtopic Z13.01, Active and Passive Dust Mitigation Surfaces.

Scope Title: Heat Pumps for High-Temperature Sink Environments

Scope Description:

Operations in extreme environments where the environment sink temperature exceeds spacecraft hardware limits will require active cooling if long-duration survivability is expected. Robotic science rovers operating on the lunar surface over diurnal cycles face extreme temperature environments. Landers with clear views of the sky can often achieve sufficient heat rejection with a zenith or, if sufficiently far from the equator, an anti-Sun-facing radiator. However, science rovers must accommodate random orientations with respect to the surface and Sun. Terrain features can then result in hot environment sink temperatures beyond operating limits, even with shielded and articulated radiator assemblies. Lunar dust degradation on radiator thermo-optical properties can also significantly affect effective sink temperatures. During the lunar night, heat rejection paths must be turned off to preclude excessive battery mass or be properly routed to reclaim nuclear-based waste heat.

Science needs may drive rovers to extreme terrains where steady heat rejection is not otherwise possible. The paradigm of swarms or multiple smaller rovers enabled by commercial lander opportunities will need to leverage standard rover bus designs to permit flexibility. A heat pump provides the common extensibility for thermal control over the lunar diurnal. Active cooling systems or heat pumps are commonly used on spacecraft. Devices used include mechanical cryocoolers and thermoelectric coolers. For higher loads, vapor compression systems have been flown, and more recently, reverse turbo-Brayton-cycle coolers are being developed under NASA's Game Changing program for high-load, high-temperature-lift cryocoolers. However, technology gaps exist for midrange heat pumps that are suitable for small science rovers where internal heat dissipation may range from 20 to 100 W.

Expected TRL or TRL Range at completion of the Project: 2 to 5

Primary Technology Taxonomy:
Level 1: TX 14 Thermal Management Systems
Level 2: TX 14.X Other Thermal Management Systems

Desired Deliverables of Phase I and Phase II:

Research
Analysis
Prototype

Desired Deliverables Description:

Conceptual design (Phase I).
Physics-based analysis or model (Phase I).
Proof-of-concept hardware (Phase I).
Proof-of-concept hardware tested against simulated loads in proposed environments (Phase II).
Final report (Phase I, Phase II).

State of the Art and Critical Gaps:

Specifically, heat pump systems are needed with the following:

Temperature lift from a cold side at <50 °C to an environmental sink temperature as high as 75 °C (temperature lift of 50 °C or heat rejection rate of 230 W/m²), with a system coefficient of performance >2.5.
Tolerance to being powered down during the lunar night and restarted during the day reliably over multiple diurnals.
Minimal exported vibrations, if any, for compatibility with science instruments.

Novel heat-pump systems are desired. Enabling improvements to state-of-the-art systems are also welcome.

Relevance / Science Traceability:

NASA's lunar initiative and Planetary Science Division form the primary customer base for this technology. Missions that directly address the National Research Council Planetary Science Decadal Survey may be users of this technology.

References:

Apollo Lunar Roving Vehicle Documentation: https://www.hq.nasa.gov/alsj/alsj-LRVdocs.html
Apollo Experience Report—Thermal Design of Apollo Lunar Surface Experiments Package: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19720013192.pdf
Thermal Considerations for Designing the Next Lunar Lander: https://aip.scitation.org/doi/10.1063/1.2437438

Scope Title: Advanced Manufacturing of Loop Heat Pipe Evaporator

Scope Description:

A loop heat pipe (LHP) is a very versatile heat transport device that has been used on many spacecrafts. At the heart of the LHP is the evaporator and reservoir assembly. During the manufacturing, tedious processes are required to machine the porous primary wick and insert it into the evaporator, and both ends of the wick need to be sealed for liquid and vapor separation. One commonly used method for vapor seal is to use a bimetallic knife-edge joint, which is more prone to failure over long-term exposure to thermal cycles and shock and vibration. These tedious manufacturing processes add to the cost of the traditional LHP. A new manufacturing technique that will allow the primary wick to be welded directly to the reservoir without the use of a knife-edge seal is needed to reduce the cost and enhance the reliability.

Expected TRL or TRL Range at completion of the Project: 4 to 6

Primary Technology Taxonomy:
Level 1: TX 14 Thermal Management Systems
Level 2: TX 14.X Other Thermal Management Systems

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware

Desired Deliverables Description:

Successfully develop advanced techniques to manufacture the LHP evaporator and reservoir assembly (Phase I).
Demonstrate the performance of the evaporator/reservoir performance in an LHP setup (Phase I).
Demonstrate the performance of the evaporator/reservoir performance in an LHP setup optimized to operate in simulated realistic environments with appropriate cycling (Phase II).
Final report (Phase I, Phase II).

State of the Art and Critical Gaps:

The LHP evaporator contains a porous wick, which provides the capillary pumping capability to sustain the fluid flow in the loop. The smaller the pore size of the wick, the higher its capillary pumping capability. However, a smaller pore size results in a higher flow resistance that must be overcome by the capillary force. Traditional sintered metal wicks have a pore size on the order of 1 µm and porosity around 0.4 to 0.6. To replace the traditional porous wick, the new wick produced by the advanced manufacturing technology must have comparable pore size and porosity. The smallest pore size currently produced by direct metal laser sintering is on the order of 10 µm.

Relevance / Science Traceability:

Traditional LHPs are used on many NASA missions including the Ice, Cloud, and Land Elevation Satellite (ICESat), ICESat-2, Swift, Aura, Geostationary Operational Environmental Satellite (GOES), Geostationary Operational Environmental Satellite-R Series (GOES-R), and Surface Water and Ocean Topography (SWOT). Similar future Science Mission Directorate (SMD) missions, especially those using small satellites, can greatly benefit from this technology.

References:

Richard, Bradley, et al.: "Loop Heat Pipe Wick Fabrication via Additive Manufacturing," NASA Thermal & Fluid Analysis Workshop, August 21-25, 2017, Marshall Space Flight Center, Huntsville, AL.

Scope Title: Approaches and Techniques for Lunar Surface Payload Survival

Scope Description:

The lunar environment poses significant challenges to small, low-power (~100 W or less) payloads, rovers, and landers required for lunar science. The lunar day/night cycle is approximately one Earth month. During that time, surface temperatures on the lunar surface can reach 400 K at local solar noon or drop to below 100 K during the lunar night—and even colder in permanently shadowed regions. These hot and cold conditions can last several Earth days, because of the slow rotation of the Moon, or permanently in shadowed craters. Lunar dust deposited on heat-rejection surfaces and coatings will increase the heat absorbed from the Sun, thus reducing the effectiveness of radiators for heat rejection. The lunar gravity, which is 1/6th of the Earth's, will limit the ability of typical low-power heat transport devices, but the gravity field may provide advantages that could be utilized. Higher heat dissipation capacity should be addressed in Z2.01. This call seeks to solicit innovative proposals to enable lunar science in the difficult lunar environment. Example technologies may include, but are not limited to, active loops that may be turned off and are freeze tolerant, zero- or low-power non-consumable/regenerative heat generation sources, high-thermal-capacitance thermal storage, advanced insulation, and passive switching with high turndown ratios (e.g., >400:1). Furthermore, small form factors are also desired. Technologies should show substantial increase over the state of the art. Technology proposals should address power usage in day and night/shadow, mass, heat transport when turned on, heat leak when turned off, temperature drops through the system, heat storage/release amount, sensitivity to lunar topography and orientation, and so forth.

Expected TRL or TRL Range at completion of the Project: 3 to 4

Primary Technology Taxonomy:
Level 1: TX 14 Thermal Management Systems
Level 2: TX 14.2 Thermal Control Components and Systems

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware

Desired Deliverables Description:

Thermal management approaches, techniques, and hardware components to enable the accommodation of temperature extremes encountered in the lunar environment. Concept model deliverable for Phase I and prototype demonstration in relevant environment in Phase II.

State of the Art and Critical Gaps:

Missions like Surveyor and Lunokhod hibernated during the night or reduced operational power near noon, in attempts to survive single or multiple lunar cycles. ALSEPs (Apollo Lunar Surface Experiments Packages) were deployed on several Apollo missions and had select experiments that operated for many lunar cycles. However, both Lunokhod and ALSEP benefited from radioisotope heat and power sources, which are either too expensive or not likely to be available for near-term future lunar science experiments. In fact, most modern lunar surface mission planning is based on solar power and batteries and typically avoids the challenges associated with surviving the full lunar cycle or shadowed regions. Because interest in lunar science and the development of abilities to deliver payloads to the lunar surface is resurgent, the capability to operate through the entire lunar environment is critical. In the absence of perpetual power supplies like radioisotope thermoelectric generators (RTGs), thermal management approaches to accommodate the lunar extremes, extended day/night cycles, and shadowed regions are seen as enabling.

Relevance / Science Traceability:

Science Mission Directorate (SMD) lunar surface science investigations will employ small, low-power payloads that will require advanced thermal control approaches and techniques to survive and operate for extended duration through extreme thermal environments on the lunar surface.

NASA has plans to purchase services for delivery of payloads to the Moon through the Commercial Lunar Payload Services (CLPS) contract. Under this subtopic, proposals may include efforts to develop payloads for flight demonstration of relevant technologies in the lunar environment. The CLPS payload accommodations will vary depending on the particular service provider and mission characteristics. Additional information on the CLPS program and providers can be found at this link: https://www.nasa.gov/content/commercial-lunar-payload-services. CLPS missions will typically carry multiple payloads for multiple customers. Smaller, simpler, and more self-sufficient payloads are more easily accommodated and would be more likely to be considered for a NASA-sponsored flight opportunity. Commercial payload delivery services may begin as early as 2021, and flight opportunities are expected to continue well into the future. In future years it is expected that larger and more complex payloads will be accommodated. Selection for award under this solicitation will not guarantee selection for a lunar flight opportunity.

References:

NASA Prepares for Performing New Science on the Moon: https://www.jpl.nasa.gov/news/news.php?release=2007-068
The Surveyor Program: https://history.nasa.gov/TM-3487/ch2-1.htm
The Surveyor Program: https://www.lpi.usra.edu/lunar/missions/surveyor/(link is external)
Missions - Lunokhod 01: https://solarsystem.nasa.gov/missions/lunokhod-01/in-depth/

Missions - Lunokhod 02: https://solarsystem.nasa.gov/missions/lunokhod-02/in-depth/

Lead Center: JSC

Participating Center(s): GRC, GSFC, JPL, MSFC

Scope Title: Spacecraft Thermal Management

Scope Description:

NASA seeks new technologies that will facilitate low-mass and highly reliable thermal control systems for the exploration of our solar system. This solicitation specifically targets proposals for new technologies and methods that clearly address one of the following areas:

Lunar surface habitat thermal technologies
High-temperature heat acquisition and transport for nuclear electric propulsion (NEP)
Topology optimization of thermal control systems

These areas are considered of equal priority, and no award preference is expected for one area over another.

1. Lunar Surface Habitat Thermal Technology Development

NASA is seeking focused efforts to develop thermal control technologies that will enable crewed habitats for extended stays on the lunar surface. Technologies should address a gap associated to long-duration habitation on the lunar surface, where temperatures range from -193 °C or lower in shadow regions (including night) to 120° C at the equatorial subsolar point. Technologies are needed that allow a single mobile habitat to operate in all these environments. Technologies should address reduction in mass, volume, and power usage relative to current solutions. The addition of heaters can lead to increased vehicle mass due to additional power generation and storage requirements and is not considered a novel architecture approach. Proposed radiator technologies should also address micrometeoroid and orbital debris (MMOD) robustness and protection potential where appropriate.

Examples of other challenges to address in this area include the deposition of dust on radiators leading to degraded optical properties, contamination-insensitive evaporators/sublimators to enable long mission life, self-healing coolant tubes for MMOD-impact resilience, and passive gas traps for removing gas bubbles from internal thermal control system loops that use low-surface-tension non-water coolants. Technologies should be suitable for use with habitats having variable heat loads averaging between 2 and 6 kW. All technologies should support a minimum operational duration of 5 years and be compatible with encountered environments.

Alternatively, technologies that utilize the conditions provided by the lunar environment to provide a critical function may also be considered; for example, air-water separator technologies that leverage the gravity field of the lunar surface, or concepts that explore the viability of utilizing the lunar surface regolith to provide long-duration thermal control function. As appropriate, such systems should also address functional capability in the microgravity environment that will be experienced prior to lunar surface operations.

2. High-Temperature Heat Acquisition and Transport for Nuclear Electric Propulsion (NEP)

NASA is seeking the development of thermal transport systems for NEP. This application requires the transfer of large amounts of thermal energy from a nuclear reactor to a power conversion system. NASA desires a high-temperature heat transfer system capable of transferring 4 to 10 MW of thermal power from a nuclear reactor, at a supply temperature of 1,200 to 1,400 K and a flux on the order of 0.3 MW/m² with a goal of 1 MW/m², to the hot-end heat exchangers of an electric power conversion system. The target distance for the power conversion system is 5 m from the reactor, but transport distances up to 10 m may be required. The system will need to be gamma- and neutron-radiation tolerant, be single-fault tolerant (a single leak should not render the system inoperable) and have an operating life of 15+ years. System mass and reliability should be addressed as part of the proposal.

Example solutions include, but are not limited to, liquid metal heat pipes or pumped fluid loops. Special consideration should be given to interfaces (both at the nuclear reactor and at the power conversion system) to maximize heat transfer. Integration with the reactor may include solutions that run through the reactor core. For integration with the power conversion system, a helium-xenon working fluid in a Brayton cycle system may be assumed but is not required.

3. Topology Optimization of Thermal Control Systems

Advanced design and manufacturing are rapidly transforming engineered systems. The advent of reliable additive manufacturing techniques coupled with robust optimization algorithms is facilitating the development of new high-performance systems. To date, the advanced design community has primarily focused on optimized structural systems that minimize mass and volume while meeting structural performance requirements. While some work has been done to develop advanced design tools for thermal control systems, considerable work remains to make it standard practice. This solicitation requests the development of a topology optimization (TO) tool that can optimize a thermal-fluid component (e.g., a heat exchanger). Specific goals include minimizing component (heat exchanger) mass, minimizing pressure drop, and maximizing heat transfer efficiency. Because of the inherent multiphysics characteristics of the problem (coupled structural/thermal/fluids behavior), proposals are encouraged to leverage existing TO software (e.g., see Watkins (2019) and other TO references below) that can already handle structural and thermal conduction optimization, and extend the code to handle systems that include single-phase laminar convective heat transfer.

This solicitation requests the development of TO software capable of minimizing heat exchanger mass while meeting envelope volume, heat transfer, and pressure drop targets. The initial target is optimization for laminar single-phase flow. An extended goal is to be able to optimize a heat exchanger for turbulent single-phase flow while accommodating manufacturing constraints to ensure the heat exchanger design is manufacturable.

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:
Level 1: TX 14 Thermal Management Systems
Level 2: TX 14.2 Thermal Control Components and Systems

Desired Deliverables of Phase I and Phase II:

Analysis
Prototype
Hardware
Software

Desired Deliverables Description:

Phase I awards in this area are expected to demonstrate analytical and/or empirical proof-of-concept results that demonstrate the ability of the organization to meet the goals stated in the solicitation.

At the conclusion of a Phase II contract, deliverables are expected to include a functioning prototype (or better) that demonstrates the potential to meet the performance goals of the technology or software. Any delivered math models should include supporting data that validates the assumptions used within the model.

State of the Art and Critical Gaps:

These focus areas strive to reduce mass, volume, and power of a thermal control system in the next generation of robotic and human-class spacecraft and to enable long-term missions to the Moon and Mars. These improvements may come through either novel hardware solutions or modernization of software tools. The current state of the art in thermal control systems is vehicle power and mass impact of greater than 25 to 30% due to old technologies still in use. Furthermore, as missions become more variable (dormancy, environments, etc.), the need for intelligent design and control (both actively and passively) within the thermal control system becomes more apparent. For topology optimization (TO) in particular, it has become a well-established structural design tool, but it has yet to penetrate the thermal design community. Multiple research efforts have shown that TO of thermal-fluid systems is possible and can be successfully implemented to obtain optimized designs; however, a robust commercial code that can do this is yet to be demonstrated. Additionally, science payloads will continue to decrease in size, increase in power, and require precise temperature control, all of which cannot be readily provided by traditional thermal control methods due to vehicle-level impacts of overall performance, mass/volume, and power.

Relevance / Science Traceability:

Long-duration habitats (Moon, Mars, etc.).
Lunar surface power.
Mars transit vehicles.
SmallSats/CubeSats.
Rovers and surface mobility.
Nuclear electric propulsion (NEP) systems.

References:

Stephan, R. "Overview of the Altair Lunar Lander Thermal Control System Design and the Impacts of Global Access," AIAA 2011-5001, 2011.
Ewert, M. K. "Investigation of Lunar Base Thermal Control System Options," SAE Transactions, J. of Aerospace, 102(1), 829-840, 1993.
Leimkuehler, T. O., et al. “Operational Experience with the Internal Thermal Control System Dual-Membrane Gas Trap,” 33rd International Conference on Environmental Systems (ICES), Vancouver, British Columbia, Canada, July 2003.
Leimkuehler, T. O., et al. “Effects of Surfactant Contamination on the Next Generation Gas Trap for the ISS Next Generation Gas Trap for the ISS Internal Thermal Control System,” 34th International Conference on Environmental Systems (ICES), Colorado Springs, CO, July 2004.
Wetch, J. R., et al. “Megawatt Class Nuclear Space Power Systems (MCNSPS) Conceptual Design and Evaluation Report,” Volumes I-IV, NASA CR-179614, September 1988.
General Atomics Project 3450. “Thermionic Fuel Element Performance Final Test Report, TFE Verification Program,” GA-A21596 (UC-224), Prepared under Contract DE-AC03-86SF16298, Department of Energy, 1994.
Ashcroft, J., and Eshelman, C. “Summary of NR Program Prometheus Efforts,” LM-05K188, 2006.
Aerojet, “SNAP-8 Performance Potential Study, Final Report,” NASA CR-72254, 1967.
Horner-Richardson, K., et al. “Fabrication and Testing of Thermionic Heat Pipe Modules for Space Nuclear Power Systems,” 27th IECEC, San Diego, CA, Paper Number 929075, 1992.
Ernst, D. M., and Eastman, G. Y. “High Temperature Heat Pipe Technology at Thermacore – An Overview,” AIAA-85-0981, 1985.
Voss, S. S., and Rodriguez, E. A. “Russian System Test Program (1970-1989),” American Institute of Physics Conference Paper 94-0101, 1994.
Stone, J. R. “Alkali Metal Rankine Cycle Boiler Technology Challenges and Some Potential Solutions for Space Nuclear Power and Propulsion Applications,” NASA Technical Memorandum 106593, July 1994.
Demuth, S. F. “SP 100 Space Reactor Design,” Progress in Nuclear Energy, Volume 42, Number 3, 2003.
Ashcroft, J. and Eshelman, C. “Summary of NR Program Prometheus Efforts,” LM-05K188, 2006.
Davis, J. E. “Design and Fabrication of the Brayton Rotating Unit,” NASA CR-1870, March 1972.
Richardson-Hartenstein, K., et al. “Fabrication and Testing of Thermionic Heat Pipe Modules for Space Nuclear Power Systems,” 27th IECEC, Paper Number 929075, 1992.
Watkins, Ryan. "Designing Optical Instruments for Space Applications: Multiphysics Topology Optimization," 2019.
Watkins, Ryan. "Topology Optimization: A Shift Towards Computational Design," 2016.
Kambampati, Sandilya, and Hyunsun A. Kim. "Level Set Topology Optimization of Load Carrying Heat Dissipation Devices," AIAA Aviation 2019 Forum, 2019.
Kambampati, Sandilya, and H. Alicia Kim. "Level Set Topology Optimization of Cooling Channels Using the Darcy Flow Model," Structural and Multidisciplinary Optimization (2020): 1-17.

Feppon, Florian, et al. "Topology Optimization of Thermal Fluid–Structure Systems Using Body-Fitted Meshes and Parallel Computing," Journal of Computational Physics 417 (2020): 109574.

NASA is pursuing rapid identification, development, and testing of capabilities that exploit small spacecraft platforms and responsive launch capabilities to increase the pace of space exploration, scientific discovery, and the expansion of space commerce in a sustainable manner. These emerging capabilities have the potential to enable new mission architectures, enhance conventional missions, and promote development and deployment on faster timelines. This will, in turn, allow NASA and other space mission operators to achieve their objectives at significantly lower programmatic risk and cost than traditional approaches.

Small spacecraft are typically defined as those weighing 180 kg or less and are often designed for shared launch using standardized form factors and interfaces and containerized deployment (e.g. CubeSats). Small spacecraft and responsive launch capabilities are proving to be disruptive innovations for exploration, discovery, and commercial applications. NASA seeks technical innovations that enable small spacecraft to rival the capabilities of their larger, more expensive counterparts, while also striving to make them cheaper and quicker to build, and easier to launch and operate. In addition, NASA seeks innovations to help address the looming concern of space debris growth in Low Earth Orbit (LEO) following the expected launch of constellations consisting of thousands of satellites, whilst also further expanding the reach of small spacecraft beyond LEO. Greatly improved capabilities are needed for lunar exploration missions, lunar communications and navigation infrastructure, and exploration at Mars and other deep space destinations. Technology and capability investment will be needed to meet these upcoming mission needs while keeping overall costs low, mission cadence high, and retaining the agile aerospace approach that has fueled what has been termed the “smallsat revolution”.

Specific improvements required are long-range high-bandwidth optical and RF communications; novel navigation devices and navigation references for use well beyond Earth; improved power management; and robust tolerance of the harsher thermal and radiation environment of deep space. Propulsion technologies with improved performance are sought for Trans Lunar Injection (TLI), lunar orbit insertion and maintenance, return-to-Earth and Earth entry and descent mechanisms. Transfer stages that host small spacecraft and can provide support services to the deployed spacecraft are also desired. Innovations are wanted to increase the speed, economy, and reliability of production; modular designs will facilitate reliable assembly and test of singly- or batch-produced small spacecraft, with specifically sought-after technologies including wirelessly interconnected sensors and modules. De-orbit or rapid disposal devices for single spacecraft, and autonomous space traffic management technologies for small spacecraft swarms and constellations are also needed. These include affordable powerful computing hardware and intelligent software tools and infrastructure for autonomous operation of spacecraft or for cooperation of spacecraft groups, minimizing human-in-the-loop bottlenecks, that are applicable to both the space debris management environment, as well as deep space missions.

NASA’s Small Spacecraft Technology Program will consider promising SBIR technologies for spaceflight demonstration missions and seeks partnerships to accelerate spaceflight testing and commercial infusion.

The following references discuss some of NASA's small spacecraft technology activities:

Small Spacecraft Technology (SST) program: https://www.nasa.gov/directorates/spacetech/small_spacecraft/index.html
Small Satellite Missions: www.nasa.gov/smallsats
Small Spacecraft Virtual Institute: https://www.nasa.gov/smallsat-institute

Another useful reference is the Small Spacecraft Technology State of the Art Report at:

https://www.nasa.gov/sites/default/files/atoms/files/soa_2021.pdf

Lead Center: GRC

Participating Center(s): ARC, GSFC, JPL

Scope Title: End-to-End Deep Space Communications

Scope Description:

Develop enabling communications technologies for small spacecraft beyond Low Earth Orbit (LEO). These technologies will be required by spacecraft to conduct NASA lunar and deep space distributed spacecraft science missions. Innovations in communications technologies for distributed small spacecraft are essential to fulfill the science missions envisioned within the decadal surveys and contribute to the success of human exploration missions. To construct the lunar communications architecture [Ref. 11], it is appropriate to consider a hybrid approach of large and small satellite assets. Primary applications include lunar surface-to-surface data relay, data relay to Earth, and navigational aids to surface and orbiting users. Distributing these capabilities across multiple small satellites may be necessary because of limited size, weight, and power (SWaP), but also to enhance coverage.

Technologies for specific lunar architectures are especially needed. For example, landers near the lunar South Pole may not have—and landers on the far side of the Moon will not have—direct line-of-sight to Earth-based ground stations and will need to send data through a relay satellite (or Gateway) to return data to Earth. Small surface systems (including rovers or astronauts on extravehicular activities (EVAs)) on the Moon will likely not have the necessary system resources to close a direct link to Earth. Human surface operations may require surface-to-surface over-the-horizon communications through an orbital relay. Deployment of sufficient traditional communications assets to maintain persistent global coverage of the lunar surface may be prohibitively expensive. Analogous to emerging LEO communications constellations, small spacecraft can operate as local relays in cislunar space.

Considerations for technology and capability extension to the Martian domain and other deep space applications are also solicited.

Interspacecraft networking is inherent to distributed mission and interoperable communications relay architectures. Enabling networking capabilities in small spacecraft requires low SWaP, low-cost hardware for radio-frequency (RF) and optical crosslinks. While network protocols developed for interoperable communications relays may be interchangeable with those for distributed missions, relay networks may not be scalable to very large-scale sensor webs of small spacecraft. As such, addressing interspacecraft networking gaps may require investment in both hardware crosslinks and networking protocols that scale to hundreds of nodes, and requires robustness for loss of nodes or as new nodes enter the network. Network management technologies may be needed due to the increased operational complexity.

An end-to-end system needs to be considered for the application of small satellites for deep space missions as described in preceding paragraphs. Therefore, enabling technologies also include non-NASA ground services that keep the operations cost commensurate with the lower costs of the small satellites themselves. Automation of the ground services as well as the small satellite constellations are needed.

Communications solutions can operate in optical or various RF bands; however, considerations must be given to bandwidth, public and Government licensing, network and data security, and compatibility with referenced candidate architectures.

Expected TRL or TRL Range at completion of the Project: 2 to 5

Primary Technology Taxonomy:
Level 1: TX 05 Communications, Navigation, and Orbital Debris Tracking and Characterization Systems
Level 2: TX 05.X Other Communications, Navigation, and Orbital Debris Tracking and Characterization Systems

Desired Deliverables of Phase I and Phase II:

Prototype
Hardware
Software

Desired Deliverables Description:

Phase I: Identify and explore options for the deep space small-satellite missions, including ground services. Conduct trade analysis and simulations, define operating concepts, and provide justification for proposed multiple access techniques, frequency bands of operation, command and data handling, and networking solutions. Also identify, evaluate, and develop design for integrated communications payload(s) and one or more constituent technologies that enable distributed spacecraft operations in the relevant space environment beyond LEO. Integrated communications system solutions and constituent component deliverables should offer potential advantages over the state of the art, demonstrate technical feasibility, and show a path toward a hardware/software infusion into practice. Bench-level or laboratory-environment-level demonstrations or simulations are desirable. The Phase I proposal should outline a path that shows how the technology can be developed into space-qualifiable and commercially available small-spacecraft communications payloads through Phase II efforts and beyond.

Phase II: Demonstration of communication technology via prototype or high-fidelity emulation. The relevant deep space environment parameters should be simulated as much as possible.

State of the Art and Critical Gaps:

Small-spacecraft missions beyond Earth require compact, low-power, high-bandwidth radios for use on the Moon, Mars, the rest of the inner planets, around asteroids or other small bodies, and at other deep space destinations. The current state of the art is the Iris radio (0.5U, 1.2 kg, and 35 W) [Ref. 12] that has been operationally used at Mars, and there is no known affordable, readily available competitor. Future missions require systems that are lower SWaP-C, can operate in multiple bands (S, X, Ka-band, and optical), and can reach uplink and downlink speeds in excess of 20 Mbps. Spectral, modulation, information layer, and protocol compatibility with current technologies (Space Communications and Navigation (SCaN)); licensing and spectrum approval; and planned Government or commercial deep space communication architecture must all be considered.

Communications among spacecraft in a distributed spacecraft mission (DSM) configuration and between the DSM configuration and the Earth become more challenging beyond LEO distances. Collaborative configurations of widely distributed (tens to hundreds of kilometers apart) small spacecraft (180 kg or less) will operate far into the near-Earth region of space and beyond into deep space, further stressing the already limited communications capabilities of small spacecraft. Alternative operational approaches with associated enabling hardware and/or software will be needed with the following:

Uplinks (Earth-to-space) and downlinks (space-to-Earth): Alternatives for coordinated command and control of the DSM configuration and individual small spacecraft from Earth as well as return of science and telemetry data to Earth. Each spacecraft cannot rely on its own dedicated Earth link, consuming valuable ground infrastructure and operators.
Integrated communications payload: Hardware and software designs for the common and unique capabilities of each small spacecraft in the DSM configuration. Spacecraft communication SWaP-C should be reduced by at least 25% from a non-DSM spacecraft.
Small-spacecraft antennas: Development of antennas optimized for either intersatellite or uplink/downlink communications are sought across a broad range of technologies including but not limited to deployable parabolic or planar arrays, active electronically steered arrays, novel antenna steering/positioning subsystems, and others suitable for use in high data rate transmission among small spacecraft over large distances. SWaP-C should be reduced from state of the art, such as the recent 6U CubeSat MarCO mission, which used a 0.2 m² X-band reflectarray to achieve 29 dBic gain and 42% efficiency [Refs. 13, 14]. Operations compatible with NASA’s space communications infrastructure [Ref. 9] and Government-exclusive or Government/non-Government-shared frequency spectrum allocations is required [Refs. 6, 7, 8].
Small-spacecraft RF solid-state power amplifiers and RF front ends with smart electronics that increase operational efficiencies.
Compatibility and interoperability with lunar communications and navigation architecture plans [Refs. 1, 2, 3]. Application of the emerging lunar standards includes frequency allocations per link functionality, modulation, coding, and networking protocol standards. Ka-band frequencies and above are highly desired.
Optical end-to-end considerations for Earth links. If a DSM design relies on an optical link to Earth, the needed ground infrastructure should be considered.

Relevance / Science Traceability:

Several missions are being planned to conduct investigations/observations in the cislunar region and beyond. For example: Commercial Lunar Payload Services (CLPS); Cislunar Autonomous Positioning System Technology Operations and Navigation Experiment (CAPSTONE); human exploration (Artemis) landing site and resource surveys; and communications and navigation infrastructure, including LunaNet, Mars communications relay, etc. Commercial and NASA small spacecraft, lunar surface assets, and manned vehicles in cislunar space and beyond will multiply within the decade. All these missions will depend on small-spacecraft communications relays, time reference transmissions, and navigation capabilities.

References:

International Communication System Interoperability Standard (ICSIS):

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