NASA SBIR 2021 Program Solicitations

Digital Transformation is the strategic transformation of an organization's processes and capabilities, driven and enabled by rapidly advancing and converging digital technologies, to dramatically enhance the organization's performance and efficiency. These advancing digital technologies include cloud computing, data analytics, artificial intelligence, blockchain, mobile access, Internet of Things, agile software development and processes, social media, and others. Their convergence is producing major transformations across industries — media and entertainment, retail, advertising, software, publishing, health care, travel, transportation, etc. Through digital transformation, organizations seek to gain or retain their competitive edge by becoming more aware of and responsive to both customer and employee interests, more agile in testing and implementing new approaches, and more innovative and prescient in pioneering the next wave of products and services. Central to the success digital transformation is the pervasive (and often transparent) gathering of data about everything that impacts success--the organization's processes, activities, competencies, products and services, customers, partners, industry, and so on. Organizations can mine this massive, complex, and often unstructured data to develop accurate insights into how to improve organizational performance and efficiency. An organization may also use this data to train machine learning algorithms to automate processes, provide recommendations, or enhance customer experiences. The digital technologies listed above are essential to generate, collect, transform, mine, analyze, and utilize this data across the enterprise. NASA is undertaking a digital transformation journey to enhance mission success and impact. NASA intends to leverage digital transformation to: 

• Boost innovation and creation of new knowledge.
• Reduce cost and increase the effectiveness and efficiency of processes for everything from human resources to science and engineering.
• Reduce the time to develop and mature new technologies.
• Facilitate efficient design and development of advanced aerospace vehicles.
• Ensure that increasingly complex missions are both cost-efficient and safe.
• Achieve data-driven insights and decisions.
• Increase autonomy in aerospace vehicles and ground facilities.
• Engage an enthusiastic and talented workforce.
• Maintain worldwide leadership in aerospace. 

Through this topic, NASA is seeking to help explore and develop technologies that may be critical to the Agency's successful digital transformation. Specific innovations being sought in this solicitation are: 

• Blockchain for aerospace applications, including its use in distributed space missions and in model-based systems engineering.
• Intelligent digital assistants that reduce the cognitive workload of NASA personnel, from scientists and engineers to business and administrative staff. 

Details about these applications of digital transformation technologies are in the respective subtopic descriptions.

Lead Center: ARC

Participating Centers: None

Scope Title:

Model-Based Systems Engineering for Distributed Development

Scope Description:

• Tools or systems that allow models to be shared across development environments and trace the resulting system model back to contributions from multiple partners while maintaining information security [International Traffic in Arms Regulations (ITAR), Export Administration Regulations (EAR), sensitive but unclassified (SBU)] and protecting intellectual property.
• Modeling visualization environments that facilitate user interaction from multiple stakeholders perspectives with varying expertise in MBSE.
• Executable requirements specification with precise semantic links disparate development sources.

Expected TRL or TRL Range at completion of the Project: 4 to 6
Primary Technology Taxonomy:
    Level 1: TX 11 Software, Modeling, Simulation, and Information Processing
    Level 2: TX 11.2 Modeling
Desired Deliverables of Phase I and Phase II:

• Prototype
• Software

Desired Deliverables Description:

Phase I: Prototype software to demonstrate proof of concept.

Phase II: Functional software that is ready to be tested in an operational environment by end users.

State of the Art and Critical Gaps:Relevance / Science Traceability:

This subtopic would be of relevance to all Human Exploration and Operations Mission Directorate (HEOMD) missions, but of particular interest will be Gateway and Artemis development. Those systems have already adopted the use of MBSE tools and tools sought to help reduce potential system integration bottlenecks. Over the next 3 to 5 years, there will be considerable opportunity for small business contributions to be matured and integrated into the support infrastructure as Gateway evolves from concept to development program. Longer term plans for human exploration, including a sustained lunar presence and manned Mars missions, would benefit from disruptive innovations that improve the entire project life-cycle including mission design, acquisition, development, and deployment.

References:

General references:

• NASA's assessment of the state of the art for MBSE: https://www.nasa.gov/nesc/articles/se-mbse-state-of-the-discipline
• SysML: http://www.omgsysml.org

Areas of specific interest with references:

• Ensuring information exchange of digital artifacts are transferable and up to date among multiple stakeholders.
  •  Digital Engineering Information Exchange Working Group (DEIX WG)
  •  CCSDS Electronic Data Sheets
• Computational tools to augment human decision making and reasoning on complex systems with large amounts of data from disparate sources
  • Augmented Intelligence for Systems Engineering challenge team (AI-SECT)
  • Chami, M., Zoghbi, C., Bruel, JM. "A First Step towards AI for MBSE: Generating a Part of SysML Models from Text Using AI", (2019).
•  Web-based interfaces including CRUD (create, read, update, delete) operations and digital review sign-offs for models particularly for reviews required in NPR 7123.1 System Engineering Requirements. 
  • Open-MBEE: https://openmbee.org
  • OSLC: https://open-services.net/
• Formal requirements specification and test case generation with traceability to a single source of truth.
  • ReqIF: https://www.omg.org/reqif/
  • FMI: https://fmi-standard.org
  • OpenCAESAR: https://opencaesar.github.io
  • Digital Thread for Smart Manufacturing

The Life Support and Habitation Systems Focus Area seeks key capabilities and technology needs encompassing a diverse set of engineering and scientific disciplines, all which provide technology solutions that enable extended human presence away from Earth, in deep space and on planetary surfaces, such as Moon and Mars. The focus is on those mission systems and elements that directly support astronaut crews, such as Environmental Control and Life Support Systems (ECLSS), Extravehicular Activity (EVA) Systems and Radiation Protection, as well as systems engineering approaches that enable vehicle and system integration.

Environmental Control and Life Support Systems encompass process technologies, equipment 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 resource recovery. 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. Technologies are of interest that enable long-duration, safe and sustainable deep-space human exploration. Special emphasis is placed on developing technologies that will fill existing gaps, 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. Because spacecraft may not be tended by crew for long periods, systems must be operable after long periods of dormancy or absence of crew.

As we consider human missions beyond earth, new technologies must be compatible with attributes of the environments we encounter, including partial gravity, atmospheric pressure and composition, space radiation, and presence of planetary dust.  Portable Life Support System (PLSS) components that require space vacuum, may not operate in the weak carbon dioxide atmosphere on Mars. For astronauts to walk once again on a distant planetary surface, an effective boot must be incorporated into the design of the exploration space suit's pressure garment. Outside of the protection of the Earth's magnetosphere, radiation in deep space will be a challenge. Electronic systems, including processors for high performance computing and power converters, for avionics within spacecraft cabins and space suits, will need to be radiation hardened or otherwise tolerant to the radiation environment. There is a wealth of commercial off-the-shelf (COTS) hardware that could potentially be used, but only if tested for tolerance to these environments

The current collaborative environment between government, commercial and international sectors will result in the distributed development of human spacecraft elements and systems for human missions of the future, such as Gateway. Their integration may benefit from advances in model based systems engineering approaches.
 
Please refer to the description and references of each subtopic for further detail to guide development of proposals.

Lead Center: JPL

Participating Centers: GRC, JSC, KSC, MSFC

Scope Title:

Spacecraft Microbial Monitoring for Long Duration Human Missions

Scope Description:

With the advent of molecular methods, emphasis is now being placed on nucleic acids to rapidly detect microorganisms. However, the sensitivity of current gene-based microbial detection systems is low (~100 gene copies per reaction), requires elaborate sample process steps, involves destructive analyses, and requires fluids to be transferred and detection systems are relatively large size. Recent advancements in the metabolomics field have potential to substitute (or augment) current gene-based microbial detection technologies that are multistepped, destructive, and labor intensive (e.g., significant crew time). NASA is soliciting nongene-based microbial detection technologies and systems that target microbial metabolites and that quantify the microbial burden of surfaces, air, and water inside for long-duration deep-space habitats.

 

Potable water:

A simple integrated, microbial sensor system that enables sample collection, processing, and detection of microbes or microbial activity of the crew potable water supply is sought. A system that is fully-automated and can be in-line in an Environmental Control and Life Support System- (ECLSS-) like water system is preferred.

 

Habitat surfaces:

Future crewed habitats in cislunar space will be crew-tended and thus unoccupied for many months at a time. When crew reoccupies the habitat they will want to quickly, efficiently, and accurately assess the microbial status of the habitat surfaces. A microbial assessment/monitoring system or hand-held device that requires little to no consumables is sought.

 

Airborne contamination:

Future human spacecraft, such as Gateway and Mars vehicles, may be required to be dormant while crew is absent from the vehicle, for periods that could last from 1 to 3 years. Before crews can return, these environments must be verified prior to crew return. These novel methods have the potential to enable remote autonomous microbial monitoring that does not require manual sample collection, preparation, or processing.

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 06.4 Environmental Monitoring, Safety, and Emergency Response
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Phase I deliverables: Reports demonstrating proof of concept, 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 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 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 a NASA facility.

State of the Art and Critical Gaps:

The state of the art on the International Space Station (ISS) for microbial monitoring is culturing and counting, as well as grab samples that are returned to Earth. NASA has invested in DNA-based polymerase chain reaction (PCR) systems, partially robotic in some cases, to eliminate the need for on-orbit culturing. However, a fully automated system is still not ready and there is still a gap for a low- or no-crew time detection system.

Relevance / Science Traceability:

 

References:

1. A list of targeted contaminants for environmental monitoring can be found at "Spacecraft Water Exposure Guidelines for Selected Waterborne Contaminants" located at: https://www.nasa.gov/feature/exposure-guidelines-smacs-swegs
2. Advanced Exploration Systems Program, Life Support Systems Project: https://www.nasa.gov/content/life-support-systems
3. NASA Environmental Control and Life Support Technology Development and Maturation for Exploration: 2018 to 2019 Overview", 49th International Conference on Environmental Systems, ICES-2019-297   https://ttu-ir.tdl.org/bitstream/handle/2346/84496/ICES-2019-297.pdf
4. National Aeronautics and Space Administration, 2020 NASA Technology Taxonomy, https://www.nasa.gov/offices/oct/taxonomy/index.html
5. NASA Standard 3001 - Requirements:  https://www.nasa.gov/hhp/standards

Lead Center: ARC

Participating Centers: JSC, MSFC

Scope Title:

Additive Manufacturing for Adsorbent Bed Fabrication

Scope Description:

Current state-of-the-art (SOA) Air Revitalization System (ARS) contaminant-removal systems utilize packed beds. Packed beds have high pressure drop, large void volumes, poor heat management, and poor mechanical stability. Some alternate sorbent technologies (e.g., structural sorbent and monolith) have been proposed previously, but they are at a low TRL and require additional research and development to prove the concepts and resolve scale-up issues. Using robocasting techniques, a type of 3D paste printing, sorbent pastes are used to print sorbent beds with custom flow paths and rod size. With this approach, sorbent beds can be designed and fabricated with controlled pressure drop, tailored flow path, minimized void spaces, good heat management, high mechanical and chemical stability, and optimized structures with high mass transfer. In addition, having the ability to formulate one’s own sorbent paste materials allows variability in binders and co-binder selections for optimal contaminant removal and thermal performance. Previous studies have been completed for a variety of sorbent pastes (activated carbon [Ref. 1], zeolite 13X [Ref. 2], 5A, 4A, polymer, amine functionalized zeolite [Ref. 3], etc.). However, these works did not focus on optimizing the printed structure for cyclic operation and addressing scale-up issues.

 

NASA aims to use the 3D-printed sorbent beds as drop-in replacements for packed sorbent beds such as those found in the Carbon Dioxide Removal Assembly (CDRA) on the International Space Station (ISS). Using robocasting techniques to print scale-up sorbent beds is also at a low TRL and requires additional development. However, it is the preferred technique over other options (e.g., structured sorbents) because, if successful, the resulting technology will yield equivalent system mass reduction due to better thermal and fluid management and mass transfer properties. Technology solutions could include, but not be limited to, SOA solid sorbent materials such as zeolite 13X, zeolite 5A, silica gel, metal-organic-frameworks (MOFs), and activated carbon. All proposed technologies should address issues related to scale-up, paste formulation, printability, mechanical and hydrothermal stability, system design, and heaters integration. The components used in the paste formulation must abide by spacecraft chemical safety standards. This subtopic is open for novel ideas that address any of the numerous technical challenges listed below for the design and fabrication of printed sorbent beds for humidity and/or CO₂ removal. This subtopic does not seek new sorbent chemistries, instead, zeolite paste formulation and paste printing are desired.

• Innovative concepts on how to make silica gel paste for use in removing water from air, either in a cabin humidity control system or as part of a CO₂ removal process requiring desiccation.
• Choosing the correct paste formulation for optimal and mechanical stability.
• Designing the lattice structures to minimize pressure drop, provide large surface area for mass transfer, and prevent channeling.
• Designing a heater system for thermal regeneration of the sorbent that would minimize contact resistance between heater and sorbent and minimize mass while providing a uniform temperature throughout the bed.  Heaters could be commercial-off-the-shelf (COTS) types (e.g. cartridge or Kapton® heaters) or they could be 3D printed.

NASA is especially interested in technologies that can be incorporated into closed-loop life-support systems. Three life-support functions of particular Interest are CO₂ removal, cabin humidity control, and trace contaminant control, as solid sorbents are particularly suited to these applications. Technologies targeting other NASA life-support functions are also of interest.

 

Proposals targeting CO₂ removal applications should consider the following:

• Improvements in sorbent CO₂ capacity and selectivity leading to smaller, more efficient components, lower energy consumption, and operation at lower CO₂ partial pressures are highly desirable.
• Increases in the robustness of sorbent materials to mechanical stresses and temperature and humidity changes/cycling.
• Full-scale systems must achieve the following performance targets:
  • CO₂ removal rate of 4.16 kg/day (a 4-crew load).
  • System must  maintain an environment with 3.0 mmHg ppCO₂ for cabin applications (based on the daily average ppCO₂).
  • System size ≤0.3 m³ for a 4-crew system.
  • Average system power  ≤500 W of power for a 4-crew system.
  • System mass of ≤450 kg for the 4-person load.
  • System must effectively separate out water vapor from cabin air (less than 100 ppm water vapor in the CO₂ product is desired).

System must effectively separate out oxygen and nitrogen from cabin air (less than 1% O₂ and 2% N₂ by volume in the CO₂ product is desired).

Expected TRL or TRL Range at completion of the Project: 1 to 3
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:

• Analysis
• Prototype
• Hardware
• Research

Desired Deliverables Description:

Phase I deliverables: Detailed sorbent paste formulation and analysis, proof-of-concept test data, and predicted performance (mass, volume, and thermal performance) for contaminant removal (e.g., carbon dioxide, water, or trace contaminants). Deliverables should clearly describe and predict performance over the SOA with an estimated scaled-up design for a 4-person crew.

 

Phase II deliverables: Delivery of technologically mature components/subsystems that demonstrate functional performance with appropriate interfaces. Prototypes should be at least at a 4-crew-member scale.

State of the Art and Critical Gaps:

Current and future human exploration missions require an optimized ARS that can reduce the system mass, volume, and power, and increase reliability. The SOA systems (CDRA, the Carbon Dioxide Reduction Assembly (CRA), and the Trace Contaminant Control System (TCCS) are adsorbent-based or catalyst-based and their performances are limited because they use COTS sorbent materials. COTS sorbent pellets/beads have fixed performance parameters (e.g., mass transfer capability), which limit the ability to tailor the sorbents to meet specific needs. Spacecraft system design requirements differ from those used in industry. For example, one industrial application focuses on removing carbon dioxide at a relatively high concentration (12% from flue gas), whereas CDRA focuses on removing carbon dioxide at low partial pressure (3 mmHg). Therefore, having the ability to tailor a sorbent to NASA objectives would lead to more efficient adsorbent systems not just for the ARS but also other life support systems that utilize sorbents (e.g., the multifiltration beds). In addition, often times COTS sorbents are sold in bulk (impractical for NASA-scale systems) and become obsolete when manufacturers cease production. Instead of having to reevaluate and redesign systems for new COTS materials to address obsolescence, NASA can use a well-characterized 3D-printed sorbent formulation to remake or even to improve SOA systems. Here, having control over the formulation of these materials could mean continuity in the use of the materials as well and an ability to optimize and tailor the materials for spacecraft use. In addition, as new materials are available for use, (e.g., MOFs), these materials can be adapted using the same 3D-printed design. That is, once the lattice and heater designs are completed, the backbone may be used for other sorbent materials. Moreover, the 3D printing can be done commercially once an acceptable paste formulation has been established. Sorbent paste printing techniques need additional technology investment to reach a level of maturity necessary for consideration for use in a flight Environmental Control and Life Support System (ECLSS). This approach offers high returns and is a paradigm shift from the SOA, as it offers the ability to control flow paths, thermal management, and mass transfer properties.

Relevance / Science Traceability:

This technology could be a drop-in replacement for the current CO₂ adsorption beds and can be proven on the ISS with potential for application in long-duration human exploration missions, including Gateway, Lunar surface, and Mars, including surface and transit. It is imperative that CO₂ be removed to support human life during space missions. This subtopic is supported by the Advanced Exploration Systems (AES) Program in an effort to improve the SOA ARS in the ECLSS.

References:

1. Wójtowicz, Marek A., Joseph E. Cosgrove, Michael A. Serio, Andrew E. Carlson, and Cinda Chullen. "Monolithic Trace-Contaminant Sorbents Fabricated from 3D-printed Polymer Precursors." (2019). https://ntrs.nasa.gov/citations/20190028890
2. Thakkar, Harshul, Stephen Eastman, Amit Hajari, Ali A. Rownaghi, James C. Knox, and Fateme Rezaei. "3D-printed zeolite monoliths for CO2 removal from enclosed environments." ACS applied materials & interfaces 8, no. 41 (2016): 27753-27761.
3. Lawson, Shane, Connor Griffin, Kambria Rapp, Ali A. Rownaghi, and Fateme Rezaei. "Amine-functionalized MIL-101 monoliths for CO2 removal from enclosed environments." Energy & Fuels 33, no. 3 (2019): 2399-2407.

Lead Center: JSC

Participating Centers: GRC

Scope Title:

Flame-Retardant Textiles for Crew Clothing and for Use in Spacecraft Cabins

Scope Description:

There is a textile technology gap for apparel fabrics for lunar and planetary human exploration.  While there are industrial fabrics that are flame retardant in oxygen-enriched atmospheres up to 100% at ambient pressure, there is no apparel or furnishing fabric that is flame retardant in enriched atmosphere of 36% oxygen at a pressure of 8.2 psi (56.5 kPa).  The challenge for developing next-to-the-skin flame-retardant fabrics comes from the many other requirements these fabrics must satisfy.  They must be comfortable.  This means they must have high drape, be soft to the touch, and have no inherent unpleasant smell.  In addition, they cannot be toxic through the skin or outgas toxic chemicals.  These fabrics must be washable and durable over a period of up to three years of repeated use.  In other words, these fabrics must have physical and mechanical properties (no static cling, color fastness, tensile strength and elongation dry and wet, tear resistance, bending stiffness, torsional stiffness, abrasion resistance, etc.) that make them suitable for use in T-shirts and pants to be worn in an atmosphere containing 36% oxygen.  NASA needs such new fabrics to send astronauts to the Moon in order to later establish a sustainable human presence beyond low Earth orbit (LEO) or on the Moon, and in preparation for a future trip to Mars.

 

The gap in textile technology that affects IVA results from the need to protect astronauts inside space vehicles and space habitats with atmosphere of 34 ± 2% oxygen at a pressure of 8.2 psi (56.5 kPa).  During the period the astronauts reside in the Lunar Lander, they will need fire protection provided by their clothing as they will not continuously wear their space suits during the entire period the lander is on the Moon.

Expected TRL or TRL Range at completion of the Project: 1 to 3
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

Desired Deliverables Description:

In Phase I, the deliverable should be a report demonstrating the feasibility to produce new flame-retardant, nontoxic apparel fibers and/or finishing treatments on existing fibers that do not support combustion in an atmosphere of 36% oxygen at a pressure of 8.2 psi. The chemical process for developing a synthetic fiber or a finishing treatment, including any test results, should be fully described to understand any toxicity issue related to processing. Furthermore, the researchers should describe the rheological, physical, and mechanical properties of the new fiber or finishing treatment and explain how these properties will make these fibers suitable for apparel applications.

 

In Phase II, the deliverable should be a fiber that can withstand the production processes used in the textile industry.  The researchers should therefore process the new fiber and experiment with different processing conditions to determine which conditions will lead to consistent results that will enable scaling-up production.  In other words, the researchers must demonstrate that they can make fine yarns that will not break or produce excessive lint when woven into fabrics. It is highly desirable that samples of fabrics be developed and evaluated. 

State of the Art and Critical Gaps:

The state of the art in flame-retardant apparel fibers and fabrics for use next to the skin is mostly represented by meta-aramids, modacrylic, and flame-retardant (FR) fibers (FR rayon, FR wool, etc.).  These fibers will not support combustion in air, but they burn in an atmosphere of 36% oxygen.  

 

The critical gap is the absence of an inherently strong, flame-retardant (in 36% oxygen), nontoxic, and comfortable fiber to use for next-to-the-skin clothing.  

Relevance / Science Traceability:

This work will benefit several space programs, namely the lunar Human Landing System (HLS), Orion, Gateway, and Artemis, enabling the astronauts to function in habitats, pressurized rovers, and other space vehicles with enriched oxygen atmospheres and to shorten prebreathe times prior to extravehicular activities (EVAs).

References:

NASA imagery collection of crew clothing in the Skylab Project. Nonflammable clothing development program, Richard Johnston and Matthew I. Radnofsky, Fire Technology 4, 88-102 (1968) https://airandspace.si.edu/collection-objects/jacket-skylab-2-kerwin/nasm_A19772817000

Lead Center: JSC

Participating Centers: None

Scope Title:

Advancements in Feedwater Supply Assembly Technology

Scope Description:

The current technology for the Feedwater Supply Assembly (FSA) has many challenges to overcome including material durability and water capacity. Therefore, new innovative ideas and solutions are sought. The FSA will be integrated into the Exploration Extravehicular Mobility Unit (xEMU) Portable Life Support System (PLSS) and contained in the suit hatch compartment. The hatch volume is not a uniform shape and the current design uses cylindrical bladders which are not capable of optimizing water volume quantities. Additionally, many challenges exist in the material currently used for the FSA bladders. This material is known for its ability to maintain cleanliness and sterility; however, when made into these particular bladders, material failure and leakages are common at low cycle counts when tested as a pressurized system. NASA has plans to go to the Moon and as the mission extends further out of low Earth orbit, durability and extensibility will become some of the most important requirements as well.
 

The FSA shall be a sterile compliant bladder, capable of storing ultrapure feedwater with a relatively high-cycle life when pressurized. In order for the thermal control loop to operate properly, a water source is needed. A volumetrically adaptable, sterile, and durable feedwater bladder is essential. The suit pressure acts on this bladder and as water evaporates, the bladder resupplies the loop. The bladder must be clean and not leak particulates or polymer chains into the water over long periods of quiescence. The maximum design pressure (MDP) for the system will be 35 psid with a nominal operating pressure of 15 psid. These bladders will be reused in a fill-drain-refill = 1 cycle environment. The current cycle life requirement is 696 cycles per bladder. Additional requirements are captured in the reference located at the following link:  https://ntrs.nasa.gov/search.jsp?R=20190033446. Having a bladder with these qualities not only buys down the safety risk of rupture, it promotes reliability at higher pressures and provides an avenue to extend Extravehicular Activity (EVA) length.
 

This subtopic is relevant to the xEMU, International Space Station (ISS), as well as commercial space companies. The goal is to have proposed solutions to be designed, built, integrated, and tested at the Johnson Space Center and integrated into the xEMU. These solutions have the potential for a direct infusion path as the PLSS 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 a concept design for infusion into the Exploration Portable Life Support System (xPLSS). 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 xPLSS 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, which need to be fulfilled to meet the new exploration requirements. The FSA is at a stall in technology development and requires new innovative ideas. This solicitation is an attempt to seek new technologies for the FSA. NASA has plans to go to the Moon and as the mission extends further out of low Earth orbit, durability and extensibility will become some of the most important requirements.

Relevance / Science Traceability:

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

References:

Feedwater Supply Assembly Requirements are located at the following links:

1. Feedwater Supply Assembly (FSA 431) requirements are located at the following link:  https://ntrs.nasa.gov/search.jsp?R=20190033446
2. Auxiliary Feedwater Supply Assembly (FSA 531) requirements are located at the following link:  https://ntrs.nasa.gov/search.jsp?R=20190033446

Note to vendor: The following two drawings referenced in the requirements shall be provided if vendor is selected for award.

1. Feedwater Supply Assembly (FSA 431) Drawing SLN 13102397
2. Auxiliary Feedwater Supply Assembly (FSA 531) Drawing SLN 13102398

 

Scope Title:

Advanced Pressure Garment Bladder Materials

Scope Description:

The current pressure garment bladder in the legacy space suit is a urethane-coated Oxford-weave nylon. This bladder material serves as the gas bladder of the space suit and, along with the restraint material, comprises the pressure garment bladder/restraint assembly which is sized and patterned to accommodate both anatomical movement and a range of sizing. The bladder is patterned using heat sealing or radio-frequency (RF) welding techniques. While this material has been acceptable for many years, there are known deficits. The urethane coating has high tack and can result in excessive friction against the skin. Embossing or flocking of the bladder, while not significantly increasing weight, may be viable solutions to this issue, although there may be others.
 

In addition, the current bladder needs to be manually wiped with biocide after each Extravehicular Activity (EVA) to prevent microbial growth. This contributes to crew overhead time and may be challenging with advanced suit architectures on the Moon which inhibit routine access to all bladder locations. An antimicrobial treatment or coating on the air-tight side of the pressure bladder will improve long-term performance of the Pressure Garment System (PGS) and reduce crew time and consumables.
 

Lastly, while the bladder material is sufficiently strong to contain the pressurization loads of the suit in the event that the restraint layer experiences catastrophic failure, it is not impervious to damage itself through puncture from a sharp edge/corner or from an incoming micrometeorite, impacting mission success and/or crew safety. As such, a self-healing bladder could mitigate this risk and provide a more robust bladder/restraint system in the next-generation suit assembly.
 

In addition to one or more of the aforementioned design goals, a successful solution should also meet all of the following requirements:

1.       The bladder material is capable of being bonded together into gore or convolute patterns without the use of an adhesive;
2.       The bladder material bonded seams shall have a bond strength of at least 85 lb/in;
3.       The bladder material shall not leak more than 3.9×10^-8 lbm/hr-in² of oxygen at 4.3 psid.

This subtopic is relevant to the Exploration Extravehicular Mobility Unit (xEMU), International Space Station (ISS), as well as commercial space companies. The goal is to have proposed solutions to be designed, built, integrated, and tested at the Johnson Space Center (JSC) 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 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 specific to Pressure Garment Bladder is desired.

State of the Art and Critical Gaps:

As the design for the new xEMU is developed, there are obvious gaps in technologies, which need to be fulfilled to meet the new exploration requirements. This solicitation is an attempt to seek new technologies for the Pressure Garment Bladder.  NASA has plans to go to the Moon and as the mission extends further out of low Earth orbit, durability and extensibility will become some of the most important requirements.

Relevance / Science Traceability:

This may be relevant to the xEMU, ISS, as well as commercial space companies.  As a new xEMU PGS is being designed, built, and tested at JSC, solutions will have a direct infusion path as the xEMU is matured to meet the design and performance goals.

References:

Note to vendor:

Sample drawings of patterned gore and/or convolute bladder assemblies shall be provided if vendor is selected for award.

 

NASA is seeking technological innovations that will accelerate development and adoption of advanced manufacturing technologies supporting a wide range of NASA Missions. NASA has an immediate need for more affordable and more capable materials and processes across its unique missions, systems, and platforms. Cutting-edge manufacturing technologies offer the ability to dramatically increase performance and reduce the cost of NASA's programs. This topic is focused on technologies for both the ground-based advancements and in-space manufacturing capabilities required for sustainable, long-duration space missions to destinations such as Mars. The terrestrial subtopic areas concentration is on research and development of advanced metallic materials and processes and additive manufacturing technologies for their potential to increase the capability and affordability of engines, vehicles, space systems, instruments and science payloads by offering significant improvements over traditional manufacturing methods. Technologies should facilitate innovative physical manufacturing processes combined with the digital twin modeling and simulation approach that integrates modern design and manufacturing. The in-space manufacturing subtopic areas which focus on the ability to manufacture parts in space rather than launch them from Earth represents a fundamental paradigm shift in the orbital supply chain model for human spaceflight.  In-space manufacturing capabilities will decrease overall launch mass, while increasing crew safety and mission success by providing on-demand manufacturing capability to address known and unknown operational scenarios.  In addition, advances in lighter-weight metals processing (on ground and in-space) will enable the delivery of higher-mass payloads to Mars and beyond.  In order to achieve necessary reliabilities, in-situ process assessment and feedback control is urgently needed. Research should be conducted to demonstrate technical feasibility and prototype hardware development during Phase I and show a path toward Phase II hardware and software demonstration and delivering an engineering development unit for NASA testing at the completion of the Phase II that could be turned into a proof-of-concept system for flight demonstration.

Lead Center: MSFC

Participating Centers: GSFC, LaRC

Scope Title:

Development of Advanced Joining Technologies for On-Orbit 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).

 

This subtopic seeks innovative engineering solutions to robotically weld materials, both fully autonomous and semiautonomous, 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.

 

Phase I is a feasibility study and laboratory proof of concept of a robotic welding process and system for on-orbit manufacturing 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 and removal capabilities adds value. 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 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 post-process inspection, in situ monitoring, or robotic arms for manipulation of structures to be welded may also be included in the Phase I effort.

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 12.4 Manufacturing
Desired Deliverables of Phase I and Phase II:

• Prototype
• Analysis
• Hardware

Desired Deliverables Description:

Phase I requires laboratory demonstration/proof of concept that (a) the system enables high-value applications of repair and assembly and (b) the system shows potential for being operated remotely with very little intervention/setup. A preliminary design and concept of operations are also deliverables under Phase I.

 

Phase II includes finalization of the design and demonstration of a ground-based prototype system.

 

Phase III would seek to evolve the technology toward a flight demonstration, either via a system mounted externally on the ISS, Exploration Crew Module (ECM), Gateway, OSAM-1, OSAM-2, lunar lander payload, or as a free-flyer.

State of the Art and Critical Gaps:

Unpressurized environment in-space manufacturing has primarily focused on fabrication of structures in the space environment. Welding is an essential supporting technology to these capabilities. Research on welding tapered off to some extent following the cancellation of the In-Space Welding Experiment (ISWE) for the space shuttle. With the emergence of the OSAM initiative, a renewed interest and focus on manufacturing structures in the space environment as an enhancing capability for long-duration missions, and as a way to remove design constraints imposed by payload fairings and launch loads, additional work on development of an in-space welding capability should be a priority. In-space welding represents an essential complementary technology to in-space fabrication techniques.

Relevance / Science Traceability:

The research requested through this solicitation is relevant to NASA programs, including (but not limited to) the following: ISS, Exploration Crew Module (ECM), Gateway, Lunar Base Camp, OSAM-1 and OSAM-2, in-Space Assembled Telescope (iSAT) and SmallSat.

References:

G. L. Workman and W. F. Kaukler, “Laser Welding in Space,” 1989.

D. Tamir et al., "In-Space Welding: Visions and Realities," 1993.

B. E. Paton and V. F. Lapchinskiĭ, "Welding in Space and Related Technologies," Cambridge International Science Publishing, 1997.

I. D. Boyd, R. S. Buenconsejo, D. Piskorz, B. Lal, K. W. Crane, and E. De La Rosa Blanco, "On-Orbit Manufacturing and Assembly of Spacecraft: Opportunities and Challenges," 2017.

S. Carioscia, B. A. Corbin, and B. Lan, "Roundtable Proceedings: Ways Forward for On-Orbit Servicing, Assembly, and Manufacturing (OSAM) of Spacecraft," 2018.

Scope Title:

Development of Large-Scale Additive Manufacturing Processes for On-Orbit 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.

 

The ability to additively manufacture large-scale structures in-space is an enabling capability needed to fully realize the game-changing impacts of OSAM. Current state-of-the-art (SOA) on-orbit manufacturing systems are constrained to a build volume similar to terrestrial additive manufacturing processes, and others are focused on linear beam and truss designs. 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. Large-scale, free-form additive manufacturing capabilities can potentially eliminate constraints on the system imposed by launch, enabling the construction of larger, more complex, and more optimized structures.

 

This subtopic seeks innovative engineering solutions to fabricate and/or repair large structures, using fully autonomous or semi-autonomous systems, in the unpressurized space environment. Current SOA terrestrial large-scale additive manufacturing processes such as wire-fed directed energy deposition, pellet-fed extruder systems, and additive friction stir deposition should be modified with an effort to reduce the footprint, mass, and power requirements for on-orbit applications.  

 

Phase I is a feasibility study and laboratory proof of concept of a robotic large-scale additive manufacturing process and system for unpressurized in-space manufacturing applications. Targeted applications for this technology include fabrication of truss structures, build-up of structural material for retrofitting spent tanks to habitat modules, and/or solar array back planes. Additional targeted applications include the repair of structures such as spacecrafts and/or payloads damaged during the ascent stage, habitat modules with micrometeoroid impact, and out-of-service components due to unforeseen circumstances and/or scheduled repairs. The Phase I effort should provide a laboratory demonstration of the manufacturing process and its applicability to aerospace-grade metallic materials, focusing on structures that represent the priority in-space manufacturing applications identified above. Work under Phase I will inform preliminary design of a robotic additive manufacturing process and a concept of operations for how the system would be deployed and operate in the space environment. The Phase I effort should also provide an assessment of the proposed process operational capabilities, volume, and power budget. A preliminary design and concept of operations are also deliverables under Phase I. Concepts for ancillary technologies such as post-process inspection, in situ monitoring, or robotic arms for manipulation of structures to be fabricated may also be included in the Phase I effort.

 

Phase I requires a demonstration/proof of concept that (a) the system enables high-value applications of in-space fabrication of large-scale structures and (b) the system shows potential for being operated remotely with very little intervention/setup. Phase II includes finalization of the design and demonstration of a ground-based prototype system. Phase III would seek to evolve the technology toward a flight demonstration, either via a system mounted externally on the ISS, Gateway, OSAM-1, OSAM-2, a lunar lander payload, or as a free-flyer. 

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 12.4 Manufacturing
Desired Deliverables of Phase I and Phase II:

• Prototype
• Analysis
• Hardware

Desired Deliverables Description:

Phase I requires a demonstration/proof of concept that (a) the system enables high-value applications of in-space fabrication of large-scale structures and (b) the system shows potential for being operated remotely with very little intervention/setup. The Phase I effort should provide a laboratory demonstration of the manufacturing process and its applicability to aerospace-grade metallic materials, focusing on structures that represent the priority in-space manufacturing applications identified above. The Phase I effort should also provide an assessment of the proposed process operational capabilities, volume, and power budget. A preliminary design of a robotic additive manufacturing process and a concept of operations for how the system would be deployed and operate in the space environment are also deliverables under Phase I. Concepts for ancillary technologies such as post-process inspection, in situ monitoring, or robotic arms for manipulation of structures to be fabricated may also be included in the Phase I effort.

 

Phase II includes finalization of the design and demonstration of a ground-based prototype system including autonomous capability.

 

Phase III would seek to evolve the technology toward a flight demonstration, either via a system mounted externally on the ISS, Gateway, OSAM-1, OSAM-2, a lunar lander payload, or as a free-flyer. 

State of the Art and Critical Gaps:

Unpressurized structure in-space manufacturing has primarily focused on fabrication of 3D-printed truss structures and beams. The OSAM-1 and OSAM-2, funded by the STMD (Space Technology Mission Directorate) Technology Demonstration Mission Program, are planning the demonstration of 3D-printed truss structures and beams. The technology advancement to multiple degrees of freedom, large-scale fabrication of structures is a priority for on-orbit manufacturing.

Relevance / Science Traceability:

The research requested through this solicitation is relevant to NASA programs, including (but not limited to) the following: ISS, Exploration Crew Module, Gateway, Lunar Base Camp, OSAM-1, OSAM-2 and in-Space Assembled Telescope (iSAT).

References:

R. G. Clinton, Jr., “NASA’s In Space Manufacturing Initiatives: Conquering the Challenges of In-Space Manufacturing,” 2017. [Online]. Available: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170011108.pdf [Accessed: 10-Oct-2019].

I. D. Boyd, R. S. Buenconsejo, D. Piskorz, B. Lal, K. W. Crane, and E. De La Rosa Blanco, "On-Orbit Manufacturing and Assembly of Spacecraft: Opportunities and Challenges," 2017.

S. Carioscia, B. A. Corbin, and B. Lan, "Roundtable Proceedings: Ways Forward for On-Orbit Servicing, Assembly, and Manufacturing (OSAM) of Spacecraft," 2018.

Scope Title:

Development of Metal Recycling Processes for On-Orbit Manufacturing

Scope Description:

Deep space missions will require a shift in the logistics paradigm to enable reuse and recycling of materials. Recycling is a significantly enhancing capability for space missions and would enable what would otherwise be nuisance material (spent components, waste items) to be utilized as feedstock for further manufacturing. This subtopic seeks innovative engineering solutions to facilitate recycling of metals commonly used in space systems in either an intravehicular (IVA) or extravehicular (EVA) environment. In an IVA use scenario, technologies might be used inside a habitat to process spent components by breaking down the structure, generating chips, and consolidating the chips into feedstock that can be used to generate new components through various in-space manufacturing processes. In an EVA use scenario, recycling technologies might be used to take metal material scavenged from large structures (such as space habitats, satellites, and spent upper stages) and reprocess it into material feedstock for on-orbit servicing, assembly, manufacturing, and repair applications. High-value materials for metal recycling include those commonly used in large-scale space structures and components of space systems: aluminum alloys, stainless steel, and titanium.

 

Current state of the art (SOA) includes metal recycling technologies commonly applied in industry (e.g., shredding, melting, solidification), but these must be modified to fit the physical footprint, power, and mass requirements for on-orbit applications. For an IVA environment, the system for initial demonstration would be constrained to an EXPRESS* rack, occupying some portion of its 0.45-m³ volume. An EXPRESS rack is designed to support up to eight individual payloads, each occupying one EXPRESS locker (each locker has an internal volume of approximately 0.057 m³, with the total rack of eight lockers providing in excess of 0.45 m³ of payload volume). The width of a single locker is 0.44 m, the height is 0.25 m, and the depth is 0.52 m. The full rack can accommodate approximately 260 kg. In an EVA environment, the metal recycling capability would likely operate as a free-flyer platform. 

 

*EXPRESS is an acronym for "EXpedite the PRocessing of Experiments to the Space Station."

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 12.1 Materials
Desired Deliverables of Phase I and Phase II:

• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Phase I is a feasibility study and proof of concept of a system for in-space recycling of one or more metal materials. Targeted applications for the technology include processing of spent components into feedstock for in-space manufacturing and/or processing of large-scale structures (such as spent upper stages) into material forms for further use. While the focus should be on material processing (which may include shredding, melting, solidifying, and other processes necessary to obtain a final feedstock form), other ancillary techniques such as sorting, purification, and material delivery/transport may also be considered as part of an overall concept of operations. How the system will be deployed and operated in either an IVA or EVA environment should also be considered. The Phase I effort should also include a laboratory demonstration of the core recycling technique with an aerospace-grade metallic material and some characterization of properties post-recycling. Relevant metrics such as power consumption, system footprint, and mass of system should also be reported, with an emphasis on scalability and adaptation to the relevant space environment. This work will collectively inform preliminary design of a metal recycling capability for in-space use. An additional constraint is that the system must be operated remotely with very little intervention/setup. Even in an IVA environment with crew, the availability of crew time to tend or service a recycling system will be very limited. An EVA environment would require fully remote operation.   

 

Phase II would include finalization of the design and demonstration of a ground-based prototype system.  

 

Phase III would seek to evolve the technology to a flight demonstration for ISS (internal or external payload) or as a free-flyer. 

State of the Art and Critical Gaps:

The current SOA is terrestrial systems for metal recycling, but these processes must be adapted to operate in the relevant space environment and comply with the system constraints of their intended use (for example, in an IVA environment as a payload). The capability to break down spent components and larger structures and repurpose them into a useful product is needed on long-duration space missions where logistics are constrained. Recycling of polymer materials on the International Space Station (ISS) into manufacturing feedstock (for 3D printing) has been previously demonstrated, but there are no current recycling capabilities for metals on the ISS. For external applications of metal recycling, NASA has funded some Next Space Technologies for Exploration Partnerships (NextSTEP) work previously on repurposing of spent rocket stages left in-orbit. Recycling, reuse, and repurposing of metals is seen as a critical gap for on-orbit servicing, assembly, and manufacturing (OSAM) and shifting the logistics paradigm from pre-positioning of spare parts to point-of-use manufacturing. 

Relevance / Science Traceability:

The research requested through this solicitation is relevant to NASA programs, including (but not limited to) the following: ISS, Exploration Crew Module, Gateway, Lunar Base Camp, OSAM-1, OSAM-2 and in-Space Assembled Telescope (iSAT).

References:

T. Prater et al., "In-Space Manufacturing at NASA Marshall Space Flight Center: A Portfolio of Fabrication and Recycling Technology Development for the International Space Station," 2018.

Lead Center: LaRC

Participating Centers: MSFC

Scope Title:

Autonomous Modular Assembly Technology for On-Orbit Servicing, Assembly, and Manufacturing (OSAM)

Scope Description:

Modeling and simulation of structures and assembly agents is necessary for verifying autonomous agent algorithms and behavior used for structures that cannot be assembled on the ground. Accurate sensing of complex and uncertain environments is necessary to provide autonomous agents with situational awareness to accomplish assembly tasks. Validation of the autonomous system behavior and in-space assembled structure accuracy in situ will require in-space metrology capabilities.

 

The scope of this subtopic includes modular hardware and software systems:

• Element 1: Algorithms and software for sensing, planning, and control of both autonomous robots and mission/task management agents.
• Element 2: Novel hardware designs (modular robots and structures).
• Element 3: Hardware and software for global (worksite-scale) metrology systems for accurately sensing agent and structure pose within an on-orbit or lunar assembly worksite.
• Element 4: Novel approaches to dynamics-based mathematical modeling for complex rigid-body connections and independent verification and validation for dynamics-based rigid multi-body mathematical models.

A solution for autonomous modular technologies for OSAM will be built on the following foundational areas:

• Heterogeneous multi-agent planning and control: Algorithms for collaboration on shared tasks for assembly of large modular space structures; task allocation among multiple agents; trajectory planning through the worksite and real-time updating of tasks and trajectories to respond to unplanned scenarios; robust and adaptive control for guaranteed performance or graceful degradation of performance for robotic manipulators and/or novel assembly agents; and teaming of humans and machines for planning, validation, and post-assembly analysis.
• Strategies and solutions for error detection and correction during the assembly process: Perception systems and/or classification algorithms independent from the assembly agent for verifying assembly steps and characterizing assembly errors. Fault/anomaly detection, diagnosis, and response to restore nominal operations or derive an acceptable alternative goal.
• Metrology systems: Global metrology systems or sensing tools that can map a worksite to facilitate agent and structure assembly path-planning for real-time task management and situational awareness and facilitate verification and validation of assembly tasks. A scalable system that can accurately measure structures at an in-space (orbital or surface) worksite with a focus on minimal supporting infrastructure is desired. Concepts with potential for integration and repurposing after construction are favored.
• Modular structures, systems, and tools: Deployables that are rigidizable by an accompanying in situ system (i.e., trusses or functional modules), can be serviced (due to modularity), are capable of moving along truss structures of variable geometries, and/or can interface with agents or be stored/stowed at a worksite where the agent mostly acts as a driver for a mobility system. Of particular interest are approaches to efficiently connect truss modules together—hardware concepts that support the interconnection of modules in the 100- to 5,000-kg range using some form of space robotics. The objective is to minimize the parasitic mass of the completed spacecraft from the modularity features that are required for intermodule assembly. Features can be added and removed to reduce this parasitic mass. Proposals are preferred that include features to connect both electrical (power and data) and structural features, noting that the connections can occur sequentially. Joining strategies that support fluid connections are of interest but are not necessary to be responsive to this subtopic area. The structural connection should occur at a minimum of three discrete locations, fixing the rigid-body motion of the two modules in all six degrees of freedom while isolating (minimizing) forces resulting from thermal-induced strain between the modules consistent with a low Earth orbit (LEO) orbit. The three (or more) connections do not have to occur simultaneously.
• Modeling and simulation: Novel approaches to dynamics-based mathematical modeling for complex rigid-body connections with nonlinear effects (for example, slider, ball, or slot connections) and independent verification and validation for dynamics-based rigid multibody mathematical models. Of particular interest are accurate dynamics-based models for joining of modules on-orbit or in planetary environments.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 10 Autonomous Systems
    Level 2: TX 10.X Other Autonomous Systems
Desired Deliverables of Phase I and Phase II:

• Hardware
• Software

Desired Deliverables Description:

Phases I and II should both result in the following deliverables: 

• Software implementations of designed algorithms.
• Documentation verifying the efficacy of the designed algorithms.
• Physical prototypes of designed hardware.
• Design documentation for any designed hardware.

Level of detail in the documentation should be commensurate with project phase.

State of the Art and Critical Gaps:

As humans venture into deeper space, communication latency will increase to the point that autonomous operations are crucial. Current technologies for autonomous robots are low Technology Readiness Level (TRL), application specific, and fragile with respect to environmental uncertainties. To enable OSAM, these technologies must be made more resilient. Many interesting ideas exist in academia but have yet to be made into a viable product.

 

Existing interfaces for modular trusses are purely structural. A critical gap is the development of interfaces that can exchange power, data, and other services over the interface.

Relevance / Science Traceability:

Achieving a robust and resilient autonomous solution for OSAM requires the intersection of many disciplines, including mechanical and electrical systems, robotics, dynamics modeling, control theory, and computer science. NASA goals that would directly benefit from this work are future lunar exploration missions, including sustained human presence on the Moon and persistent space platforms.

References:

NASA in-Space Assembled Telescope (iSAT) Study: https://exoplanets.nasa.gov/exep/technology/in-space- assembly/iSAT_study/

NASA Raven: https://www.nasa.gov/mission_pages/station/research/experiments/explorer/Investigation.html?#id=1734

NASA Robotic Refueling Mission (RRM3): https://sspd.gsfc.nasa.gov/RRM3.html

NASA Restore-L: https://sspd.gsfc.nasa.gov/restore-L.html

NASA Dragonfly:  https://www.nasa.gov/mission_pages/tdm/irma/nasas-dragonfly-project-demonstrates-robotic- satellite-assembly-critical-to-future-space.html

Autonomous Systems NASA Capability Overview: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20180007804.pdf

Lead Center: GSFC

Participating Centers: LaRC, MSFC

Scope Title:

Adjustable End-of-Robot-Arm Force-Torque Sensor Technique

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 end-of-arm force-torque sensing technique that is adjustable is needed to enable a variety of OSAM mission-specific tasks to be performed. A ground-based demonstration of the technique is of interest.

 

Current state of the art in end-of-arm force-torque sensors is that they are not placed close enough to the sensing point to minimize the load distal of the sensor, do not work in both unrestricted 1g and 0g environments, and cannot simultaneously handle large (6× to 10×) full-range overload for launch loads while providing absolute accuracy stability for at least 30 min before zeroing. A sensing technique is needed that can provide the following:

 

• Range of 0 to 200-500 N, resolution of 2 N, and absolute accuracy of ±5 N for surface contact measurements.
• Range of ±120 N, resolution of 0.1 N to 0.2 N, and absolute accuracy of ±2 N for servicing tasks.
• Range of 0 to 20 N, resolution of 0.02 N, and absolute accuracy of ±0.1 N for payload determination measurements.

Expected TRL or TRL Range at completion of the Project: 2 to 4
Primary Technology Taxonomy:
    Level 1: TX 04 Robotics Systems
    Level 2: TX 04.1 Sensing and Perception
Desired Deliverables of Phase I and Phase II:

• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Phase I deliverables include:

• Background research and feasibility studies.
• Modeling to demonstrate feasibility.
• Conceptual design, trade studies, and description of proposed solution.
• Demonstrations of subsystems or key technologies.
• Pathfinder technology demonstrations.
• Brassboard force-torque sensor.

Phase II deliverables include:

• Demonstration using the brassboard force-torque sensor.
• Environmental testing of key components.

State of the Art and Critical Gaps:

Current state of the art in end-of-arm force-torque sensors is that they are not placed close enough to the sensing point to minimize the load distal of the sensor, do not work in both unrestricted 1g and 0g environments, and cannot simultaneously handle large (6× to 10×) full-range overload for launch loads while providing absolute accuracy stability for at least 30 min before zeroing.

Relevance / Science Traceability:

An adjustable force-torque sensor is relevant to dexterous robotic missions such as OSAM-1, OSAM-2, International Space Station (ISS) robotics, robotic sample return missions, Gateway, and in-Space Assembled Telescope.

References:

On-Orbit Satellite Servicing Study Project Report, October 2010, https://sspd.gsfc.nasa.gov/images/NASA_Satellite%20Servicing_Project_Report_0511.pdf

Scope Title:

Centralized Robot Actuator Servo Controller Board for MUSTANG (Modular Unified Space Technology Avionics for Next Generation) Form Factor

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. A robot actuator servo control electronic board that fits in the Modular Unified Space Technology Avionics for Next Generation (MUSTANG) form factor is needed to enable a variety of OSAM and robotic sample return missions. A ground-based demonstration of the control electronics is of interest.

 

Current state of the art for robot actuator servo control electronic boards is that they do not read all of the sensors necessary for closed-loop control, are too big (>30 by 30 cm), and weigh too much (>2 kg). An actuator servo control electronic board is needed that fits the MUSTANG form factor, reads all of the necessary sensors (Hall effect sensor, resolver), and weighs less than 2 kg.

Expected TRL or TRL Range at completion of the Project: 2 to 4
Primary Technology Taxonomy:
    Level 1: TX 04 Robotics Systems
    Level 2: TX 04.3 Manipulation
Desired Deliverables of Phase I and Phase II:

• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Deliverables include a ground-based demonstration of a robot actuator servo control electronic board in the MUSTANG form factor commanding a robot actuator.

Phase I deliverables include:

• Background research and feasibility studies.
• Modeling to demonstrate feasibility.
• Conceptual design, trade studies, and description of proposed solution.
• Demonstrations of subsystems or key technologies.
• Pathfinder technology demonstrations.
• Brassboard robot actuator servo control electronic board.

Phase II deliverables include:

• Demonstration using the brassboard robot actuator servo control electronic board.
• Environmental testing of key components.

State of the Art and Critical Gaps:

Existing centralized robot actuator servo control electronics do not fit the MUSTANG form factor or require additional resources other than those provided by the rest of the MUSTANG avionics platform.

Relevance / Science Traceability:

Robot actuator servo control electronics are relevant to dexterous robotic missions such as OSAM-1, OSAM-2, International Space Station (ISS) robotics, robotic sample return missions, Gateway, and in-Space Assembled Telescope.

References:

On-Orbit Satellite Servicing Study Project Report, October 2010, https://sspd.gsfc.nasa.gov/images/NASA_Satellite%20Servicing_Project_Report_0511.pdf

MUSTANG (Modular Unified Space Technology Avionics for Next Generation): https://ntrs.nasa.gov/citations/20190028692

As NASA strives to explore deeper into space than ever before, lightweight structures and advanced materials have been identified as a critical need. The Lightweight Materials, Structures, Advanced Assembly and Construction focus area seeks innovative technologies and systems that will reduce mass, improve performance, lower cost, be more resilient and extend the life of structural systems. Reliability will become an enabling consideration for deep space travel where frequent and rapid supply and resupply capabilities are not possible.  

Improvement in all of these areas is critical to future missions. Applications include structures and materials for launch, in-space and surface systems, deployable and assembled systems, integrated structural health monitoring (SHM) and technologies to accelerate structural certification. Since this focus area covers a broad area of interests, specific topics and subtopics are chosen to enhance and or fill gaps in the space and exploration technology development programs as well as to complement other mission directorate structures and materials needs. 

Specific interests include but are not limited to: 

• Improved performance and cost from advances in composite, metallic and ceramic material systems as well as nanomaterial and nanostructures.
• Improved performance and mass reduction in innovative lightweight structural systems, extreme environments structures and multifunctional/multipurpose materials and structures.
• Improved cost, launch mass, system resiliency and extended life time by advancing technologies to enable large structures that can be deployed, assembled/constructed, reconfigured and serviced in-space or on planetary surfaces.
• Improved life and risk mitigation to damage of structural systems by advancing technologies that enhance nondestructive evaluation and structural health monitoring.
• Improved approaches that provide the development of extreme reliability technologies. 

The specific needs and metrics for this year's focus technology needs are requested in detail in the topic and subtopic descriptions.

Lead Center: LaRC

Participating Centers: GRC

Scope Title:

Lunar Surface Solar Array Structures

Scope Description: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 Technology Readiness Level (TRL). TRLs at the end of Phase II of 4 or higher are desired.

State of the Art and Critical Gaps:

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 10-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 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 40 kW of power. The 10-kW-class solar array structures are also applicable for Science Mission Directorate (SMD) ConOps (concept of operations) on the Moon to recharge batteries on a Mars Science Laboratory- (MSL-) class rover.

References:

1. 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/
2. Fincannon, J., “Characterization of Lunar Polar Illumination From a Power System Perspective,” NASA TM-2008-215186, May 2008, https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080045536.pdf.
3. 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.
4. McEachen, M. et al., “Compact Telescoping Array: Advancement from Concept to Reality,” AIAA Paper 2018-1945, January 2018, https://doi.org/10.2514/6.2018-1945.

Lead Center: LaRC

Participating Centers: MSFC

Scope Title:

Real-Time Defect Detection, Identification, and Correction in Wire-Feed and Fused-Filament Additive Manufacturing

Scope Description:Expected TRL or TRL Range at completion of the Project: 2 to 3
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:

• Hardware
• Software
• Prototype
• Research
• Analysis

Desired Deliverables Description:

Phase I: Concept studies documenting the feasibility of incorporating sensor data feedback and appropriate software tools and computation resources to be used to detect defects during fabrication of parts with complex geometries, evaluating the potential impact of the defects on the performance of the parts and the correction of those defects.

 

The proposed solution must include all of the following: (1) defect sensing and detection, (2) assessment of the impact of the defect on part performance, and (3) corrective actions other than scrapping of the build. Proposals that do not clearly include these three elements will be considered out of scope.

 

Phase II: Scale demonstration of a printer with closed-loop control that incorporates defect detection, identification, and correction during fabrication. The complexity of defects that are detected and corrected, as well as the size of the parts, should demonstrate the challenges that would come up in full-scale use of the control processes. Printed part sizes should be at least 10 cm per side for cubes, with detectable defects down to the mm scale or smaller. The defects should have a demonstrable effect on the part performance, such as a decrease in mechanical properties, that is then corrected for by the process.

State of the Art and Critical Gaps:

AM is seeing rapidly expanding applications in many areas, including in aerospace. Despite this growth in AM, filling its full potential has always been limited by quality-control issues and certification of the manufactured parts, as each component that is built is unique. Some work has begun to add defect detection and correction to powder-based manufacturing processes, such as direct metal laser sintering (DMLS) and wire-feed AM. There has, however, not been the requisite advance in ensuring that defect detection and identification is coupled with the real-time correction of those defects and ensuring final performance of the manufactured part in a particular application.

 

Gap: Real-time defect detection, identification, and correction in AM processes that would ensure the performance of the as-printed parts without relying on postproduction inspection processes, with parts built with acceptable form, fit, and function the first time, with sufficient documentation to allow direct entry into service, has not been demonstrated.

Relevance / Science Traceability:

This topic fits under STMD (Space Technology Mission Directorate). It supports Advanced Manufacturing of Lightweight Structures. Enhancing quality control in AM opens up its use in many industrial applications as well as its use by NASA. In particular, in-space use of AM in future Gateway, lunar, and Mars exploration missions will require that parts that are produced are ready for use as-produced, because there will be limitations in availability of material for reprinting as well as limitations on crew time and equipment for postprinting inspection.

References:

1. https://www.ge.com/additive/blog/new-manufacturing-milestone-30000-additive-fuel-nozzles [GE Additive news release: “New manufacturing milestone: 30,000 additive fuel nozzles”] (last visited on 09/28/2020)
2. https://newatlas.com/superdraco-test/32292/ [New Atlas: “SpaceX completes qualification test of 3D-printed SuperDraco thruster”] (last visited on 09/28/2020)
3. https://3dprintingindustry.com/news/rocket-lab-celebrates-100th-3d-print-of-its-rutherford-engine-158380/ [3D Print Industry: "Rocket Lab Celebrates 100th Rutherford Engine Build"] (last visited on 09/28/2020)
4. https://www.nasa.gov/sites/default/files/atoms/files/msfcstd3716baseline.pdf [MSFC Technical Standard EM20: "Standard for Additively Manufactured Spaceflight Hardware by Laser Powder Bed Fusion in Metals," MSFC-STD-3716] (last visited on 09/28/2020)
5. https://www.thefabricator.com/additivereport/blog/wire-feed-3d-printing-grows-in-popularity [the Additive Report: "Wire-feed 3D printing grows in popularity"] (last visited on 09/28/2020)
6. https://www.ibm.com/blogs/internet-of-things/iot-3d-printing-quality-manufacturing/ [IBM Internet of Things blog: “Why quality is the obstacle to mass adoption of 3D printing”] (last visited on 09/28/2020)
7. https://www.engineering.com/AdvancedManufacturing/ArticleID/19416/The-Importance-of-Closed-Loop-Control-in- Directed-Energy-Deposition-Additive-Manufacturing.aspx [Isaac Maw engineering.com: “The Importance of Closed-Loop Control in Directed Energy Deposition Additive Manufacturing”] (last visited on 09/28/2020)
8. https://www.mdpi.com/2076-3417/9/4/787 [Shassere et al.: "Correlation of Microstructure and Mechanical Properties of Metal Big Area Additive Manufacturing," Applied Sciences, 9, 2019 (4) 787] (last visited on 09/28/2020)

Lead Center: LaRC

Participating Centers: ARC, GSFC

Scope Title:

Nondestructive Evaluation (NDE) Sensors, Modeling, and Analysis

Scope Description:

Additionally, techniques for quantitative data analysis of sensor data are desired. It is also considered highly desirable to develop tools for automating detection of material foreign object debris (FOD) and/or defects and evaluation of bondline and in-depth integrity for lightweight rigid and/or flexible ablative materials. 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 targeting 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, but techniques addressing weld inspection will also be considered. Most of the aerospace components will be metallic in nature, and critical flaws are on the range of 1 mm or smaller and can be volumetric or fracturelike in nature.

Structural Health Monitoring (SHM)

Future manned space missions will require spacecraft and launch vehicles that are capable of monitoring the structural health of the vehicle and diagnosing and reporting any degradation in vehicle capability. This subtopic seeks new and innovative technologies in SHM and integrated vehicle health management (IVHM) systems and analysis tools.

Techniques sought include modular/low mass-volume systems; low-power, low-maintenance systems; and systems that reduce or eliminate wiring, as well as standalone smart-sensor systems that provide processed data as close to the sensor as practical and systems that are flexible in their applicability. Examples of possible systems include surface-acoustic-wave- (SAW-) based sensors, passive wireless sensor tags, flexible sensors for highly curved surfaces, and direct-write film sensors. Damage detection modes include leak detection, ammonia detection, micrometeoroid impact, and others. Reduction in the complexity of standard wires and connectors and enabling sensing functions in locations not normally accessible with previous technologies is also desirable. Proposed techniques should be capable of long-term service with little or no intervention. Sensor systems should be capable of identifying material state awareness and distinguishing aging-related phenomena and damage-related conditions. It is considered advantageous that these systems perform characterization of age-related degradation in complex composite and metallic materials. Measurement techniques and analysis methods related to quantifying material thermal properties, elastic properties, density, microcrack formation, fiber buckling and breakage, etc., in complex composite material systems and in adhesively bonded/built-up and/or polymer-matrix composite sandwich structures are of particular interest. Some consideration will be given to the IVHM/SHM ability to survive in on-orbit and deep space conditions, allow for additions or changes in instrumentation late in the design/development process, and enable relocation or upgrade on orbit. The system should allow NASA to gain insight into performance and safety of NASA vehicles as well as commercial launchers, vehicles, and payloads supporting NASA missions.

Inclusion of a plan for detailed technical operation and deployment is highly favored.

NDE Modeling

Technologies sought under this SBIR include near real-time realistic NDE and SHM simulations and automated data reduction/analysis methods for large datasets. Simulation techniques will seek to expand NASA’s use of physics-based models to predict inspection coverage for complex aerospace components and structures and to utilize inverse methods for improved defect characterization. Analysis techniques should include optimized automated reduction of NDE/SHM data for enhanced interpretation appropriate for detection/characterization of critical flaws in space-flight structures and components, and may involve methods such as machine learning and domain transformation. NASA's interest area is lightweight structural materials for spaceflight, such as composites and thin metals. Future purposes will include application to long-duration space vehicles as well as validation of SHM systems.

Techniques sought include advanced material-energy interaction (i.e., NDE) simulations for high-strength lightweight material systems and include energy interaction with realistic damage in complex 3D component geometries (such as bonded/built-up structures). Primary material systems can include metals, but it is highly desirable to target composite structures. NDE/SHM techniques for simulation can include ultrasonic, laser, microwave, terahertz, infrared, x-ray, x-ray computed tomography, fiber optic, backscatter x-ray, and eddy current. It is assumed that any data analysis methods will be focused on NDE techniques with high-resolution, high-volume data. Modeling efforts should be physics-based, and it is desired that they can account for material aging characteristics and induced damage, such as micrometeoroid impact. Examples of damage states of interest include delamination, microcracking, porosity, and fiber breakage. Techniques sought for data reduction/interpretation will yield automated and accurate results to improve quantitative data interpretation to reduce large amounts of NDE/SHM data into a meaningful characterization of the structure. It is advantageous to use coprocessor/accelerator-based hardware (e.g., field-programmable gate arrays (FPGAs) and GPUs) for simulation and data reduction. Combined simulation and data reduction/interpretation techniques should demonstrate ability to guide the development of optimized NDE/SHM techniques, lead to improved inspection coverage predictions, and yield quantitative data interpretation for damage characterization.

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:

NDE tools for flight still do not have sufficient resolution to provide representation on the order of finite-element models (FEMs) allowing for a digital twin. Also, as NDE tools grow and sensors get faster, larger volumes of data are created and thus new computational tools are required. At the same time, low-cost emerging computational hardware, such as GPUs, is enabling the growing use of advanced physics-based models for improved NDE inspection and for advanced data analysis methods such as machine learning. Development of new techniques are enabling Orion to meet its 100% inspected mission directive. In addition, as NASA strives to go deeper and longer, new tools need to be developed in order to support long-duration spaceflight.

Relevance / Science Traceability:References:

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, Cincinatti, 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 Centers: MSFC

Scope Title:

Manufacturability Assessment as a Design Constraint for Advanced Tailorable Composites

Scope Description:

While use of composite material systems in space vehicle structures is increasing, their broader implementation is still facing multiple challenges. One of these challenges, especially relevant in the context of advanced tailorable composites, is a manufacturability assessment that could be leveraged as a design constraint early in the development. Absence of a reliable and comprehensive manufacturability assessment often restricts exploitation of the full potential of the composite material system and promotes conservative fabrication approaches that can negate potential mass benefits of composites. Suboptimal designs such as quasi-isotropic lamination stacking sequence treat the material as “black aluminum,” failing to permit tailored load paths possible with composites. Consequently, a material system that can enable large reductions in system mass ends up being used in the design process in a way that results in structural components that are much heavier than necessary.

 

This solicitation seeks to advance the analytical predictive manufacturability assessment capabilities for layered composites. Such an assessment is required as a constraint in the structural design and optimization process. Layered pre-impregnated composite materials reinforced with either continuous or short fibers, and with a broad spectrum of ply thicknesses, are in the scope of this solicitation. For the continuous-fiber composites, the typical fiber areal weights range from approximately 30 grams per square meter (gsm) to just under 200 gsm. For short-fiber applications, the ultrathin plies in the single-digit gsm range are also of interest. Since a broad variety of composite material systems is within the scope of this solicitation, the scope of related manufacturability constraints can be tailored to the proposed material system. Consequently, the examples of manufacturability considerations provided in this solicitation are not all-inclusive, and parameters not listed below can be offered up for investigation when a compelling rationale is presented. The list of potential implementations of advanced tailorable composites is also provided to aid in the determination of manufacturing constraints of primary importance. The overarching approach should be based on recognizing that the more aggressive the tailoring of a composite material system, the greater the risk of developing defects in the manufacturing process. While avoiding any defects can be one countermeasure, quantifying the effects of defects can provide additional insights. Specifically, if the effect of defect can be quantified and it can be demonstrated that the benefits gained from the associated design tailoring can overcome the adverse effect of a small defect, then a rationale for the more aggressive tailoring can be justified. Examples of defects include, but are not limited to, tow gaps, overlaps, foldovers, or poor compaction/separation from the underlying plies. In cases where a placement imperfection can be subjected to restorative actions, analysis of how successful such actions can be is also of value.

 

The manufacturability assessment shall be based on analytical tools validated with experimentation/actual manufacturability trials. Examples of the performance-based manufacturing constraints for the continuous-fiber composites include, but are not limited to, determination of the minimum allowable steering radius, relative proximity of ply terminations such that no significant adverse interaction occurs, and minimizing thermal warping in the cure process when nonsymmetric lamination stacking sequence is present. For example, one of the important examples of the manufacturability assessment for short-fiber composites is determination of an initial preform thickness distribution that would produce a desired thickness distribution after forming that involves preform expansion and/or turning or steering (i.e., forming similar to deep-stamping or closed-cross-section forming with a clamshell tool and an expandable bladder). Similarly, estimation of the short-fiber orientation distortions in the forming process and predictive countermeasures to this behavior are also sought.

 

Subscale manufacturing demonstrations representative of the challenges that can be encountered in the full-scale applications are sought as the manufacturability predictive analysis validation cases. The full-scale structures of interest that can be scaled down for the manufacturing demonstrations include cryogenic tanks, crew modules, pressurized habitats, and other dry and unpressurized primary structural components or portions of thereof. Discrete features such as hatches, access and windows cutouts, and hard point attachments can be considered as manufacturability challenges to be encountered in the full-scale applications in addition to the general acreage considerations. Furthermore, manufacturability assessment for short-fiber composites is expanded to allow smaller size demonstrations, for example, those representative of hard point attachment brackets, hinges, clevises, etc.

Expected TRL or TRL Range at completion of the Project: 5 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
• Software

Desired Deliverables Description:

In Phase I, a comprehensive identification of manufacturability constraints shall be conducted for a selected material system and application. Analytical and experimental results shall be obtained for at least one selected aspect of the manufacturability assessment.

In Phase II, the analytical tool capabilities and related manufacturability trials shall be expanded to a more comprehensive and exhaustive list of manufacturability constraints identified in Phase I. A manufacturing demonstration article and the computational tool available for release shall be delivered at the conclusion of Phase II. 

State of the Art and Critical Gaps:

Present composite designs are typically limited to straight fiber arrangements and lamination stacking sequences resulting in quasi-isotropic material properties. Apart from the lack of structural design and optimization tools, highly tailored fabrication of composites presents manufacturing challenges and is prone to introducing manufacturing defects. Manufacturability assessment is often performed via the trial-and-error approach, which is costly, time consuming, and very limited in scope. A validated computational manufacturability assessment tool would enable more rapid and comprehensive manufacturability assessment that can inform the design effort in its early stages.

Relevance / Science Traceability:

 

• Space Technology Mission Directorate (STMD), Artemis/HLS (human landing systems) programs, developers of air-launched systems, e.g., Generation Orbit Launch Services.
• Aeronautics Research Mission Directorate (ARMD), next-generation airframe technology beyond "tube and wing" configurations, e.g., hybrid/blended wing body.
• Developers of the in situ nondestructive evaluation inspection capabilities within both space and aeronautics programs.

References:

1. Brooks, T. R., Martins, J. R., On Manufacturing Constraints for Tow-Steered Composite Design Optimization, Composite Structures, Vol. 204, pp. 548-559, November 2018.
2. Grenoble, R. Nguyen, T., McKenney, M., Przekop, A., Juarez, P., Gregory, E., Jegley, D., Oremont, L., Fabrication of a Composite Tow-Steered Structure for Air-Launch Vehicle Applications, AIAA-2018-1376, AIAA SciTech Forum, Kissimmee, FL, 2018.
3. Wu, K., Tatting, B., Smith, B., et al., Design and Manufacturing of Tow-Steered Composite Shells Using Fiber Placement, AIAA-2009-2700, Structures, Structural Dynamics, and Materials Conference, Palm Springs, CA, 2009.
4. Antunes, N., Dardis, J., Grandine, T., Farmer, B., Hahn, G., Design Optimization of Short Fiber Composite Parts, AIAA-2020-0163, AIAA SciTech Forum, Orlando, FL, 2020.

NASA is interested in technologies for advanced in-space propulsion systems to reduce travel time, increase payload mass, reduce acquisition costs, reduce operational costs, and enable new science capabilities for exploration and science spacecraft.  The future will require demanding propulsive performance and flexibility for more ambitious missions requiring high duty cycles, more challenging environmental conditions, and extended operation.  This focus area seeks innovations for NASA propulsion systems in chemical, electric, nuclear thermal and advanced propulsion systems related to human exploration and science missions.  Propulsion technologies will focus on a number of mission applications including ascent, descent, orbit transfer, rendezvous, station keeping, and proximity operations.

Lead Center: GRC

Participating Centers: JSC, MSFC

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 (ISRU). Anticipated outcome of Phase I proposals is expected to deliver proof of the proposed concept with some sort of basic testing or physical demonstration. Proposals shall include plans for a prototype and demonstration in a defined relevant environment (with relevant fluids) at the conclusion of Phase II.

• Integrated refrigeration cycles for a combination of hydrogen and oxygen liquefaction on the lunar surface. Cycles should be initially sized for at least 11.7 metric tons per year (3.3 kg/hr of oxygen and 0.4 kg/hr of hydrogen). It is desired to minimize the mass and power of the system. Proposals should compare total input power and mass to liquefaction of fluids separately. The main contaminate is water; while the final contamination level is not known, some sensitivity should be explored in the 10s of ppm range in each stream. For Phase I, the main product should be cycle analysis and configuration, including the key sensitivities of the cycle. Phase II should include some level of buildup and test/demonstration of system.
• 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. The condensation subgrid model should be validated against experimental data (with a target accuracy of 25%), with emphasis on cryogenic fluid-based condensation data. The subgrid model and implementation scheme should be a deliverable. Phase I should be focused on simplified geometries (vertical plates/walls), while Phase II should be focused on complicated geometries (full propellant cylindrical).
• Integrated cryogenic propellant gas generation system for lander vehicles and supporting architecture: Design and develop concepts to enable integrated cryogenic propellant gas generation for lander vehicle reactor coolant system (RCS) gas accumulators. Proposers shall consider vehicle designs that use either liquid hydrogen/liquid oxygen or liquid methane/liquid oxygen main propellant combinations. Designs shall be capable of outputting a minimum 3,000-psia storage press at 300-K storage temperature and meeting the following minimum mass gasification rates: 0.1 g/sec hydrogen, 0.3 g/sec methane, and/or 0.5 g/sec oxygen. The gas generation system shall demonstrate novel integration into alternate vehicle heat sources such as thermal control systems, active CFM cooling systems, fuel cells, internal combustion (I/C) engines, electrical power systems, pumps, etc. Proposed gas generation system shall not couple to vehicle main engines or RCS thruster during firing operations. Proposers should consider integration into vehicle system architectures, mass efficiency, and minimization of propellant waste. Phase I effort should include vehicle integration concept design, design of autogenous pressurization hardware, and test demonstration of autogenous pressurization hardware using liquid cryogens. Phase II should focus on system refinement and a scale test demonstration using liquid propellants.
• Develop cryogenic mass flow meters applicable to liquid oxygen and methane, having a volumetric flow measurement capacity of 1 to 20 L/min (fluid line size of approximately ½ in.), of rugged design that is able to withstand launch-load vibrations (e.g., 20g rms), with remote powered electronics (not attached to the flowmeter), able to function accurately in microgravity and vacuum environment, and having measurement error less than +/- 0.5% of the mass flow rate reading. Ability to measure bidirectional flow, compatibility with liquid hydrogen, and ability to measure mass flow rate during two-phase flows is also desired. Designs that can tolerate gas flow without damage to the flowmeter are also desired. Goal is proof of concept end of Phase I, working flowmeter  end of Phase II.

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

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 it is hydrogen/oxygen or methane/oxygen systems, including chemical propulsion and nuclear thermal propulsion. Several recent Phase IIs have resulted from CFM subtopics, most notably for advanced insulation, cryocoolers, and liquid acquisition devices.

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, the latter two of which are 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:

1. Johnson, et al. "Comparison of oxygen liquefaction methods for use on the Martian surface." Cryogenics 90, 60-69, 2018.
2. Green, R. and Kleinhenz, J. "In-Situ Resource Utilization (ISRU) Living off the Land on the Moon and Mars." American Chemical Society National Meeting & Exposition; March 31, 2019 - April 04, 2019; Orlando, FL; United States.
3. Stochl, R., et al. "Autogenous pressurization of cryogenic vessels using submerged vapor injection." NASA-TM-104516, 1991.

Lead Center: MSFC

Participating Centers: GRC, SSC

Scope Title:

Reactor and Fuel System for Nuclear Thermal Propulsion (NTP) and Nuclear Electric Propulsion (NEP)

Scope Description:Expected TRL or TRL Range at completion of the Project: 2 to 5
Primary Technology Taxonomy:
    Level 1: TX 01 Propulsion Systems
    Level 2: TX 01.4 Advanced Propulsion
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware
• Research

Desired Deliverables Description:

Desired deliverables for this technology would include research that can be conducted to determine technical feasibility of the proposed concept during Phase I and show a path toward a Phase II hardware demonstration. Testing the technology in a simulated (as close as possible) NTP environment as part of Phase II is preferred. Delivery of a prototype test unit at the completion of Phase II allows for followup testing by NASA.

Phase I Deliverables: Feasibility analysis and/or small-scale experiments proving the proposed technology to develop a given product (Technology Readiness Level (TRL) 2 to 3). The final report includes a Phase II plan to raise the TRL. The Phase II plan includes a verification matrix of measurements to be performed at the end of Phase II, along with specific quantitative pass-fail ranges for each quantity listed.

Phase II Deliverables: A full report of component and/or breadboard validation measurements, including populated verification matrix from Phase I (TRL 3 to 5). Also delivered is a prototype of the proposed technology for NASA to do further testing if Phase II results show promise for NTP application. Opportunities and plans should also be identified and summarized for potential commercialization of the proposed technology. Unique government facilities can be used as part of Phase II.

State of the Art and Critical Gaps:

The state of the art is reactor fuel developed for the Rover/NERVA program in the 1960s and early 1970s. The fuel was carbon based and had what is known as "midband" corrosion, which affected the fuel endurance. Switching over to cermet (metal and ceramics) or advance carbide fuels shows promise but has fabrication challenges. Limited property data for most materials at ultrahigh temperatures considered makes the material performance analysis to meet the engine operating requirements riskier.

Focus is on a range of modern technologies associated with NTP using solid-core nuclear fission reactors and technologies needed to ground test the engine system and components. The engines are pump fed ~25,000 lbf with an Isp goal of 900 sec (using hydrogen) and are used individually or in clusters for the spacecraft's primary propulsion system. The NTP can have multiple startups (>5) with cumulative run time >200 min in a single mission, which can be no more than 2-yr round trip, according to a recent NASA study. The Rover/NERVA program ground tested a variety of engine sizes for a variety of burn durations and startups.

Relevance / Science Traceability:

By closing these ultrahigh-temperature data gaps, the Space Nuclear Propulsion (SNP) project intends to infuse the results into design considerations/optimizations for risk reduction. In addition to directly benefiting SNP by closing the current material data gaps, the technology improvements in high-temperature materials would also benefit the following:

• Department of Defense (DOD) Defense Advanced Research Projects Agency (DARPA) NTP program.
• Wing leading-edge systems, due to their use of refractory alloy base structures to 2,000 °C.
• Fission surface power, due to the use of materials in long-term high-temperature environments.
• Refractory reaction control systems (RCSs) that reach up to 2,000 °C temperatures.
• Refractory rocket nozzles for upper stages and landers that reach ~2,200+ °C.

STMD (Space Technology Mission Directorate) is supporting the SNP project.  

 Future mission applications:

• Human missions to Mars.
• Science missions to the outer planets.
• Planetary defense.

Some technologies may have applications for fission surface power systems.

 

 

References:

Solid-core NTP has been identified as an advanced propulsion concept that could provide the fastest trip times for human missions to Mars over a variety of mission years. NTP had major technical work done between 1955 and 1973 as part of the Rover and Nuclear Engine for Rocket Vehicle Application (NERVA) programs. A few other NTP programs followed, including the Space Nuclear Thermal Propulsion (SNTP) program in the early 1990s. The NTP concept is like a liquid chemical propulsion system, except instead of combustion in the thrust chamber, a monopropellant is heated with a fission reactor (heat exchanger) in the thrust chamber and exposes the engine components and surrounding structures to a radiation environment.

1. Durham, F. P., “Nuclear Engine Definition Study Preliminary Report”, LA-5044-MS Vol I, Informal Los Alamos Report, 1972.
2. Altseimer, J. H., Mader, G. F., Stewart, J. J., “Operating Characteristics and Requirements for the NERVA Flight Engine,” Paper 70-676, AIAA 6th Propulsion Joint Specialist Conference, June 1970.
3. Angelo, J. A., Buden, D., “Space Nuclear Power”, OrbitBook Company 1985.
4. Koenig, D. R., “Experience Gained from the Space Nuclear Rocket Program (Rover)”, LA-10062-H, Los Alamos Document, 1986.
5. Walton, J. T., “An Overview of Tested and Analyzed NTP Concepts”, AIAA-91-3503.
6. Finseth, J. L., “Overview of Rover Engine Tests- Final Report”, NAS 8-37814, 1991.
7. Bhattacharyya, S.K. et al, “Space Exploration Initiative Fuels, Materials and Related Technologies – Nuclear Propulsion Technology Panel Final Report”, NASA Technical Memorandum 105706, September 1993.
8. Haslett, R. A., “Space Nuclear Thermal Propulsion Program Final Report”, PL-TR-95-1064, Phillips Laboratory Air Force Report, 1995.

Lead Center: GRC

Participating Centers: None

Scope Title:

Structurally Robust Magnetic Circuit Materials for Hall-Effect Thrusters

Scope Description:

Electric propulsion for space applications has demonstrated tremendous benefit to a variety of NASA, military, and commercial missions. Critical NASA electric propulsion needs have been identified in the scope areas detailed below. Proposals outside the described scope shall not be considered. Proposers are expected to show an understanding of the current state of the art (SOA) and quantitatively (not just qualitatively) describe anticipated improvements over relevant SOA materials, processes, and technologies that substantiate NASA investment.
 

To shape the magnetic fields needed for operations, Hall-effect thrusters utilize a magnetic circuit that also forms the thruster structure. The magnetic circuit components direct magnetic flux (typically produced by electromagnetic coils) and may experience operational temperatures in excess of 500 °C due to coil self-heating and the close proximity of plasma-wetted surfaces. Both low-carbon magnetic iron and cobalt-iron (Co-Fe) soft ferromagnetic alloys have been traditionally used in the role; low-carbon magnetic iron is typically cheaper with larger billet size availability, whereas Co-Fe soft ferromagnetic alloys are attractive due to high magnetic saturation and Curie temperature properties. As Hall-effect thrusters become larger to support future high-power applications, thruster components also experience and must survive increased inertial launch loads. To address this issue, prospective magnetic circuit materials are desired with improved structural strength compared to SOA options while retaining comparable or better magnetic and thermal properties. Prospective materials capable of being produced in machinable, large-diameter (i.e., >400 mm) solid billets—or that can be additively manufactured to achieve comparable sizes—are of particular interest. This solicitation seeks such prospective magnetic circuit materials suitable for Hall-effect thruster applications with the following properties:

• Mechanical: Meets or exceeds yield stress properties in Table X2.4 of ASTM Standard A801-14.
• Magnetic: Meets or exceeds properties in Appendix X1 of ASTM Standard A848-17.
• Thermal: Meets or exceeds Curie temperature of 770 °C.

Expected TRL or TRL Range at completion of the Project: 2 to 4
Primary Technology Taxonomy:
    Level 1: TX 01 Propulsion Systems
    Level 2: TX 01.2 Electric Space Propulsion
Desired Deliverables of Phase I and Phase II:

• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Phase I:

1. Virtual kickoff meeting with the NASA Technical Monitor and potential stakeholders within the first month of the period of performance. 
2. A final report containing test data characterizing key material properties as well as an assessment of material size scalability for future production.
3. Material samples that can be utilized for independent verification of claimed improvements over SOA materials.

Phase II:

1. Kickoff meeting with NASA Contracting Officer Representative (COR) and potential stakeholders within the first month of the period of performance.
2. A final report with test data either characterizing key material properties for produced large billets or demonstrating the functionality of one or more thruster components integrated with operating thruster hardware (in which partnering with electric propulsion developers may be necessary).

State of the Art and Critical Gaps:

SOA magnetic circuit materials used for Hall-effect thrusters are typically in two families: low-carbon magnetic iron or cobalt-iron (Co-Fe) soft ferromagnetic alloys (e.g., Hiperco®). While Co-Fe alloys are frequently preferred because of their magnetic and thermal properties, their available billet sizes do not readily accommodate larger thruster components needed for future high-power (i.e., >50 kW) electric propulsion applications. Low-carbon magnetic iron does come in large billet sizes, but past NASA high-power thruster development efforts (e.g., NASA-457Mv2 thruster) have identified potential risks regarding the survivability of components when subjected to launch loads. A magnetic circuit material that retains or exceeds the magnetic and thermal properties of SOA options while providing improved structural strength and scalability to large billet sizes is highly desirable to mitigate the risk.

Relevance / Science Traceability:

Both NASA's Science Mission Directorate (SMD) and Human Exploration and Operations Mission Directorate (HEOMD) need spacecraft with demanding propulsive performance and greater flexibility for more ambitious missions requiring high duty cycles and extended operations under challenging environmental conditions. Planetary spacecraft need the ability to rendezvous with, orbit, and conduct in situ exploration of planets, moons, and other small bodies (i.e., comets, asteroids, near-Earth objects, etc.) in the solar system; mission priorities are outlined in the decadal surveys for each of the SMD divisions (https://science.nasa.gov/about-us/science-strategy/decadal-surveys). For HEOMD, higher-power electric propulsion is a key element in supporting sustained human exploration of cislunar space.

This subtopic seeks innovations to meet future SMD and HEOMD propulsion requirements in electric propulsion systems related to such missions. The roadmap for such in-space propulsion technologies is covered under the 2020 NASA Technology Taxonomy, with archival information contained in the 2015 NASA Technology Roadmap TA-2 (In-Space Propulsion Technologies).

References:

• D.M. Goebel and I. Katz, “Chapter 7: Hall Thrusters,” Fundamentals of Electric Propulsion: Ion and Hall Thrusters, https://descanso.jpl.nasa.gov/SciTechBook/SciTechBook.html
• ASTM Standard A801-14, “Standard Specification for Wrought Iron-Cobalt High Magnetic Saturation Alloys (UNS R30005 and K92650).”
• ASTM Standard A848-17, “Standard Specification for Low-Carbon Magnetic Iron.”
• D.F. Susan, et al., “Equal Channel Angular Extrusion for Bulk Processing of Fe-Co-2V Soft Magnetic Alloys, Part I: Processing and Mechanical Properties,” Journal of Materials Research, 33.15 (2018): 2168-2175.
• A.B. Kustas, et al., “Equal Channel Angular Extrusion for Bulk Processing of Fe-Co-2V Soft Magnetic Alloys, Part II: Texture Analysis and Magnetic Properties,” Journal of Materials Research, 33.15 (2018): 2176-2188.
• Z. Turgut, et al., “High Strength Bulk Fe-Co Alloys Produced by Powder Metallurgy,” Journal of Applied Physics, 103.7 (2008): 07E724.
• M.S. Masteller, J.W. Bowman, and L. Li, “High Temperature Aging Behavior of High Strength 49% Co-1.9% V-0.3% Nb-Fe Soft Magnetic Alloy,” IEEE Transactions on Magnetics, 32.5 (1996): 4839-4841.
• Decadal surveys for each of the SMD divisions, https://science.nasa.gov/about-us/science-strategy/decadal-surveys
• 2020 NASA Technology Taxonomy, https://www.nasa.gov/offices/oct/taxonomy/index.html
• 2015 NASA Technology Roadmap TA-2 (In-Space Propulsion Technologies), https://www.nasa.gov/sites/default/files/atoms/files/2015_nasa_technology_roadmaps_ta_2_in-space_propulsion_final.pdf

Scope Title:

High-Efficiency, Long-Life Hollow Cathodes

Scope Description:

Electric propulsion for space applications has demonstrated tremendous benefit to a variety of NASA, military, and commercial missions. Critical NASA electric propulsion needs have been identified in the scope areas detailed below. Proposals outside the described scope shall not be considered. Proposers are expected to show an understanding of the current state of the art (SOA) and quantitatively (not just qualitatively) describe anticipated improvements over relevant SOA materials, processes, and technologies that substantiate NASA investment.

 

Hollow cathodes in electric propulsion systems are utilized for generating discharge plasma and effecting plume neutralization in gridded-ion and Hall-effect thrusters. In SOA hollow cathodes, operating temperatures can range from 1,000 to 1,700 °C, and the cathode assembly may need to survive in excess of 10,000 operational hours and 10,000 thermal on-off cycles without failure. Critical NASA needs for hollow cathodes are:

 

1. High-current hollow cathodes with reduced power consumption. While SOA hollow cathodes can provide up to 25-A direct current necessary for electric propulsion applications, future interest in 100-kW electric propulsion systems will require a substantial increase in cathode current output to the range of 100 to 200 ADC. Scaling of current cathode architectures using various emitter technologies have achieved cathodes operating at >100-ADC emission current; however, these results typically require substantial increases in electrical power needed to drive plasma generation in the cathode and/or in an associated heating element for impregnate-based emission sources. Size increases for emitter and cathode, including heating elements, can also be significant to maintain the necessary thermal conditions for stable cathode life; the resultant larger sized cathodes can stress heater elements and limit their cyclic life—a concern facing cathodes utilizing LaB6 emitters. This solicitation seeks stable-performance, long-life cathode architectures that reduce power consumption (i.e., improve electrical efficiency) for >100-ADC emission current via improved heater design and operation, emitter material selection and configuration, lower plasma generation costs, reduced cathode thermal losses via conduction or radiation, etc.
2. Reduced-flow hollow cathodes in Hall-effect thrusters. Hollow cathodes used in Hall-effect thrusters are frequently operated with a fixed flow fraction relative to the anode flow; this approach is commonly utilized to reduce the cost and complexity of the propellant feed system. To promote efficient discharge plasma generation, these cathodes are typically operated with a higher than necessary propellant flow, which reduces specific impulse and may have negative impacts on cathode lifetime due to pressure-driven emission behavior. Past efforts to bifurcate the cathode flow between the cathode and an external (i.e., keeper or downstream region) contribution have demonstrated some success in providing stable and efficient cathode operation while reducing the total cathode (i.e., non-anode) flow fraction to less than 7% to 10% of the anode flow rate typically used in thruster operations. Being able to sustain thruster operations at such low total cathode flow fractions can result in significant propellant savings, particularly for high-power Hall-effect thrusters. This solicitation seeks readily adaptable methods to reduce cathode propellant flow needs (i.e., improve propellant utilization) without adversely affecting cathode and Hall-effect thruster stability and life.

Expected TRL or TRL Range at completion of the Project: 2 to 5
Primary Technology Taxonomy:
    Level 1: TX 01 Propulsion Systems
    Level 2: TX 01.2 Electric Space Propulsion
Desired Deliverables of Phase I and Phase II:

• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Phase I:

1. Kickoff meeting with the NASA Technical Monitor and potential stakeholders within the first month of the period of performance.
2. A final report containing quantitative analysis, modeling, or proof-of-concept test data addressing key risk factors associated with the technical approach and comparisons to SOA cathodes.
3. A cathode subsystem design that is compatible with high-power Hall-effect thruster concepts.

Phase II:

1. Kickoff meeting with the NASA Contracting Officer Representative (COR) and potential stakeholders within the first month of the period of performance.
2. A final report with test data supporting cathode performance, stability, and lifetime claims.
3. Cathode assembly hardware that can be utilized for independent verification of claimed improvements over SOA cathode assemblies.

State of the Art and Critical Gaps:

Future interest in 100-kW electric propulsion systems will require cathode current outputs in the range of 100 to 200 ADC. Experience to date with scaling current cathode architectures has resulted in cathodes that consume several kilowatts of power during operations. Such cathodes pose significant thermal management challenges for the thruster and concerns about the cathode's cyclic lifetime. Alternative cathode architectures that can significantly reduce power consumption are highly desirable to reduce risk for high-power electric propulsion applications.

Typical Hall-effect thrusters utilize a cathode flow fraction between 7% and 10% of the anode flow, with past studies of 50-kW-class thrusters at times requiring >10% cathode flow fraction to promote thruster stability at certain throttle points. For high-power electric propulsion systems utilizing Hall-effect thrusters, reducing cathode propellant flow needs can result in significant propellant savings on the order of hundreds of kilograms for typical NASA mission lifetimes. Past efforts to bifurcate the cathode flow between the cathode and an external (i.e., keeper or downstream region) contribution have demonstrated some success in providing stable and efficient cathode and thruster operations while achieving <7% total cathode flow fraction. Approaches for reducing cathode flow needs that can be readily adapted to SOA thruster architectures are highly desirable to improve system efficiency and lifetime.

Relevance / Science Traceability:

Both NASA's Science Mission Directorate (SMD) and Human Exploration and Operations Mission Directorate (HEOMD) need spacecraft with demanding propulsive performance and greater flexibility for more ambitious missions requiring high duty cycles and extended operations under challenging environmental conditions. Planetary spacecraft need the ability to rendezvous with, orbit, and conduct in situ exploration of planets, moons, and other small bodies (i.e., comets, asteroids, near-Earth objects, etc.) in the solar system; mission priorities are outlined in the decadal surveys for each of the SMD divisions (https://science.nasa.gov/about-us/science-strategy/decadal-surveys). For HEOMD, higher-power electric propulsion is a key element in supporting sustained human exploration of cislunar space.

 

This subtopic seeks innovations to meet future SMD and HEOMD propulsion requirements in electric propulsion systems related to such missions. The roadmap for such in-space propulsion technologies is covered under the 2020 NASA Technology Taxonomy, with archival information contained in the 2015 NASA Technology Roadmap TA-2 (In-Space Propulsion Technologies).

References:

• V.J. Friedly and P.J. Wilbur, “High Current Hollow Cathode Phenomena,” AIAA 90-2587.
• M.A. Mantenieks and R.M. Myers, “Preliminary Test Results of a Hollow Cathode MPD Thruster,” IEPC 91-076.
• D.M. Goebel and E. Chu, “High Current Lanthanum Hexaboride Hollow Cathodes for High Power Hall Thrusters,” IEPC-2011-053.
• H. Kamhawi and J. Van Noord, “Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters,” AIAA-2012-4080.
• M.L. Plasek, et al., “Experimental Investigation of a Large-Diameter Cathode,” AIAA-2014-3825.
• D.M. Goebel, K.K. Jameson, and R.R. Hofer, “Hall Thruster Cathode Flow Impact on Coupling Voltage and Cathode Life,” Journal of Propulsion and Power, Vol. 28, No. 2, March-April 2012.
• S.J. Hall, et al., “Operation of a High-Power Nested Hall Thruster with Reduced Cathode Flow Fraction,” Journal of Propulsion and Power, July 2020.
• Decadal surveys for each of the SMD divisions, https://science.nasa.gov/about-us/science-strategy/decadal-surveys
• 2020 NASA Technology Taxonomy, https://www.nasa.gov/offices/oct/taxonomy/index.html
• 2015 NASA Technology Roadmap TA-2 (In-Space Propulsion Technologies), https://www.nasa.gov/sites/default/files/atoms/files/2015_nasa_technology_roadmaps_ta_2_in-space_propulsion_final.pdf

NASA's Human Research Program (HRP) investigates and mitigates the highest risks to astronaut health and performance for exploration missions. The goal of the HRP is to enable space exploration beyond low earth orbit by reducing the risks to human health and performance through a focused program of basic, applied and operational research leading to the development and delivery of:

• Human health, performance, and habitability standards.
• Countermeasures and other risk mitigation solutions.
• Advanced habitability and medical support technologies.

HRP has developed an Integrated Research Plan (IRP) to describe the requirements and notional approach to understanding and reducing the human health and performance risks. The IRP describes the Program's research activities that are intended to address the needs of human space exploration and serve HRP customers. The Human Research Roadmap (http://humanresearchroadmap.nasa.gov) is a web-based version of the IRP that allows users to search HRP risks, gaps, and tasks.

The HRP is organized into several research Elements:

• Human Health Countermeasures.
• Human Factors and Behavioral Performance.
• Exploration Medical Capability.
• Space Radiation.

Each of the HRP Elements address a subset of the risks. A fifth Element, ISS Medical Projects (ISSMP), is responsible for the implementation of the research on various space and ground analog platforms. HRP subtopics are aligned with the Elements and solicit technologies identified in their respective research plans. 

Lead Center: JSC

Participating Centers: None

Scope Title:

Radioprotectors and Mitigators of Space Radiation-Induced Health Risks

Scope Description:

Space radiation is a significant obstacle when sending humans on long-duration missions beyond low Earth orbit. Although various forms for radiation exist in space, astronauts during Lunar or Mars missions will be exposed constantly to galactic cosmic radiation (GCR), which consists of high-energy particles ranging from protons to extremely heavy ions. Astronaut health risks from space radiation exposure are categorized into cancer, late and early central nervous systems (CNS) effects, and degenerative risks, which include cardiovascular diseases (CVD) and premature aging. With the current gender and age-specific exposure limits for cancer risks, few female astronauts will be able to fly long-duration missions without countermeasures.

This subtopic solicits proposals to develop biological countermeasures that mitigate one or several of the radiation risks associated with space travel. Compounds that target common pathways (e.g., inflammation) across aging, cancer, cardiovascular disease, and neurodegeneration would be preferred. Most of the countermeasure developments in the medical arena have focused on mitigating the effects of X- or gamma rays. The proposed project should focus on repurposing of technology and compounds for high-energy charged-particle applications. Compounds that are under current development or have been proven effective for other applications are both suitable for this subtopic.

Expected TRL or TRL Range at completion of the Project: 5 to 8
Primary Technology Taxonomy:
    Level 1: TX 06 Human Health, Life Support, and Habitation Systems
    Level 2: TX 06.5 Radiation
Desired Deliverables of Phase I and Phase II:

• Analysis

Desired Deliverables Description: State of the Art and Critical Gaps:

Exposure of crew members to space radiation during Lunar and Mars missions can potentially impact the success of the missions and cause long-term diseases. Space radiation risks include cancer, late and early CNS effects, CVD, and accelerated aging. Abiding by the current exposure limits for cancer risks, few female astronauts will be able to fly long-duration missions. Mitigation of space radiation risks can be achieved with physical (shielding) and biomedical means. This subtopic addresses development of drugs that mitigate one or several of the identified space radiation risks. Development of countermeasures for adverse health effects from radiation exposure is also actively supported by the Department of Defense (DOD), Department of Homeland Security (DHS), and the National Institute of Health (NIH). However, some of the radioprotectors used in radiotherapy might have toxic levels that are unacceptable for astronauts. Some of the countermeasures developed for DOD/DHS are aimed at mitigating acute radiation syndromes, but not cancer risks. Furthermore, these radioprotectors are mostly for exposure to X- or gamma rays. This SBIR subtopic solicits specifically proposals to evaluate the radioprotectors that have been proven effective in mitigating biological impacts of X- or gamma rays for space radiation applications.

Relevance / Science Traceability:

This subtopic seeks technology development that benefits the Space Radiation Element of the NASA Human Research Program (HRP). Biomedical countermeasures are needed for all of the space radiation risks.

References:

The following references discuss the different health effects NASA has identified in regard to space radiation exposure:

• Evidence report on central nervous systems effects: https://humanresearchroadmap.nasa.gov/evidence/reports/CNS.pdf.
• Evidence report on degenerative tissue effects: https://humanresearchroadmap.nasa.gov/evidence/reports/Degen.pdf.
• Evidence report on carcinogenesis: https://humanresearchroadmap.nasa.gov/evidence/reports/Cancer.pdf.

Lead Center: JSC

Participating Centers: None

Scope Title:

Portable Spatial Disorientation Simulator - Trainer

Scope Description:

Astronauts are at risk of spatial disorientation due to vestibular alterations during and following g-level transitions, such as landing on Earth. This disorientation has previously been simulated using a bilateral bipolar Galvanic vestibular stimulation (GVS) delivered in a suprathreshold range (2 to 5 mA) over the mastoid processes independent of head orientation. NASA needs a portable GVS-based system that can be coupled to head orientation and movements to enhance the simulation of the g-transition induced spatial disorientation effect astronauts experience.

This system will be used for astronaut crewmembers to simulate performing landing and recovery type tasks while experiencing head-tilt contingent vertigo due to vestibular alterations. This simulator will also be used by recovery operations personnel to validate nominal and contingency procedures with a simulated deconditioned crewmember. Finally, this disorientation simulator will be used experimentally to develop sensorimotor standards related to fitness to perform critical mission tasks.

 

The requirements include:

• Phase 1A head-worn inertial measurement unit (IMU) sensor that can measure natural head rotation (position and velocity) and linear acceleration in all three planes.
• A GVS that is head-coupled and proportional to head tilt orientation as well as pitch and roll velocity, with the ability to adjust the algorithms to alter the IMU sensor combinations that drive the GVS signal.
• The system should also allow a user-adjustable manual gain to allow for individual sensitivity, with minimal two-fault current limit at 5 mA and emergency on/off switch.
• The system should allow a two-channel multiple electrode configuration that can provide illusory motion in both head roll and pitch axes.
• The system should be self-powered for minimally 1 hr with user switchable rechargeable batteries.
• The system should include nonvolatile memory (onboard data storage) to record IMU sensor data, GVS current delivery, and external trigger and/or manual synch event push-button timing.
• This system should be able to be worn while performing nonsuited crew landing and egress type activities without interfering with other crew-worn equipment.

Expected TRL or TRL Range at completion of the Project: 2 to 6
Primary Technology Taxonomy:
    Level 1: TX 06 Human Health, Life Support, and Habitation Systems
    Level 2: TX 06.6 Human Systems Integration
Desired Deliverables of Phase I and Phase II:

• Prototype

Desired Deliverables Description:

Phase I deliverable is a laboratory version of the disorientation trainer that successfully demonstrates the proof of concept for the requirements listed under scope description have been met.
 

Phase II deliverable is a portable wearable version of the disorientation trainer that can be deployed in field settings.

State of the Art and Critical Gaps:

While there are GVS available, there are no GVS devices on the market that are portable or that can be coupled to head movement. This capability would provide the ability to train astronauts on what to expect with regards to spatial disorientation in a realistic mission simulation.

Relevance / Science Traceability:

This is relevant to Human Exploration and Operations Mission Directorate (HEOMD), because of its applicability in human research and exploration. For example, this technology would assist in the success of the sensorimotor standards project, sponsored by NASA's Human Research Program.

References:

Moore ST, Dilda V, MacDougall HG. Galvanic vestibular stimulation as an analogue of spatial disorientation after spaceflight. Aviat Space Environ Med. 2011;82(5):535-542. (download from https://www.ingentaconnect.com/content/asma/asem/2011/00000082/00000005/art00006)

 

The Science Mission Directorate 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 like: 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 following references discuss some of NASA's science mission and technology needs:

• The Astrophysics Roadmap: http://nasascience.nasa.gov/about-us/science-strategy.
• Astrophysics Decadal Survey - “New Worlds, New Horizons: in Astronomy and Astrophysics”: http://www.nap.edu/catalog.php?record_id=12951 (link is external).
• The Earth Science Decadal Survey: http://books.nap.edu/catalog.php?record_id=11820 (link is external).
• The Heliophysics roadmap: “The Solar and Space Physics of a New Era: Recommended Roadmap for Science and Technology 2009­2030”: http://hpde.gsfc.nasa.gov/2009_Roadmap.pdf.
• The 2011 Planetary Science Decadal Survey was released March 2011. This decadal survey is considering technology needs: https://solarsystem.nasa.gov/2013decadal/.

Lead Center: GSFC

Participating Centers: AFRC, JPL

Scope Title:

Planetary Aerial Vehicles for Venus

Scope Description:

NASA is interested in scientific exploration of Venus using aerial vehicles to perform in situ investigations of its atmosphere, surface, and interior structure. The 2019 Venus Exploration Analysis Group (VEXAG) Strategic Plan identified several key science objectives 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.

 

Venus features a challenging atmospheric environment that significantly impacts the design and operation of aerial vehicles. NASA is currently developing concepts for controlled-variable-altitude balloons with payloads of up to 200 kg operating at an altitude range between 52 and 62 km over a latitude range of 0° to +/-60°. Proposals for the following Venus aerial vehicle components are encouraged: (1) Entry, deployment, inflation technologies for a Venus balloon, (2) Instrument sondes, and (3) Helium transfer pump.

1. The most critical phase of a Venus balloon mission is the transition from atmospheric entry to a free-flying configuration. Concepts for any or all of the critical phases of the transition are desired: deployment of the balloon from the atmospheric entry vehicle, inflation of the balloon, and separation of the balloon from the inflation system and parachute system.
2. Deployment of instrument sondes from the payload could enhance and lengthen the balloon mission operating lifetime by reducing payload mass as lift capability is lost over time. Sondes with a mass up to 5 or 10 kg should be capable of operating for several hours, carry a small science instrument payload, and be able to communicate with the primary balloon mission. The sondes envisioned for this solicitation are categorized into ascending and descending investigations. Ascending science investigations carry small science payloads up to 70 km altitude, and descending science investigations carry a small science payload down to near the surface (i.e., <10 km altitude). Proposals offering both heavier-than-air and lighter-than-air (relative to the Venus atmosphere) solutions are solicited. Furthermore, the sonde concepts may have powered propulsion or unpowered flight. Suggested vehicle types include, but are not limited to:
   • Solar-heated balloons that would operate on the sunlit side. This kind of sonde would be deployed from the payload gondola, auto-inflate in a free fall through the atmosphere, and attain float as the solar balloon heats from the Sun. This could possibly operate either above or below the primary balloon mission altitude range.
   • Probes deployed from the payload gondola that perform stabilized vertical descents, gliding descents, powered ascents, or a combination of both ascents and descents.
3. A controlled-variable-altitude balloon may use a pump to transfer helium from a zero-pressure balloon into a superpressure balloon. Pumping technologies capable of pumping helium with a pressure rise of 50 kPa at 100 liters per minute are desired. Multistage or parallel flow pump solutions are acceptable for consideration. Light weight and high efficiency are important factors in the pump since it must fly with the balloon payload.

It is expected that a Phase 1 effort will consist of a system-level design and a proof-of-concept experiment on one or more key components. Proposers should be familiar with the atmospheric pressure, temperature, solar, infrared (IR) heating, and corrosion aspects of the Venus atmosphere as described in this call. The atmospheric temperature ranges from -30°C at 62 km to 62°C at 52 km, the pressure ranges from about 18 kPa at 62 km to 80 kPa at 52 km (Venus International Reference Atmosphere, VIRA [see Kliore, 1985]), the solar flux can be as high as 2,300 W/m² at 62 km, and the IR heat flux coming up from the lower atmosphere can be as high as 830 W/m² at 52 km [Crisp, 1986]. The sulfuric acid vapor abundance is less than less than 1 ppmv at 52 km and above [Oschlisniok, 2012]. The sulfuric acid liquid aerosols have a concentration between 75% (pH -1.5) and 90% (pH -2.0) [Titov, 2018]. Although the cloud droplets are highly acidic, they are very small, typically in the range of 1 to 10 µm in diameter, and fairly diffuse, with cloud droplet abundance only on the order of 100 droplets/cm³ for the 1-µm-sized particles; and on the order of 10 droplets/cm³ for the larger (r > 3µm) particles. The maximum observed aqueous H₂SO₄ content in the balloon operating environment is on the order of only 30 mg/m³ [Knollenberg, 1980]. Additional information on the Venus atmospheric environment can be found in the References section.

Expected TRL or TRL Range at completion of the Project: 2 to 3
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

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. 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:

Terrestrial-based aerial vehicles, including lighter-than-air and heavier-than-air vehicles, are mature technologies and continue advancing in capability, reliability, and autonomy. However, these need adaptation for operation in the Venus environment.
Several gaps exists in aerial vehicle technology for Venus atmospheric flight:

1. There is a strong need for aerial deployment systems for balloons and their payloads since most balloons are launched from the ground and from the upper atmosphere. Methods for deployment may leverage techniques for Mars entry vehicle systems that deploy from an aeroshell and eventually separate from a parachute. However, a balloon inflation inserted into the middle of this sequence is a complicating element, and preventing damage to the balloon is paramount.
2. Small instrument sondes or vehicles for expanding the exploration range and mission duration have not been sufficiently developed for a Venus mission to be included as part of future mission proposals. Novel vehicles for conducting science that can be deployed from the balloon payload could play an important role in meeting these objectives. The guidance, stabilization, and control of sondes has been identified as a need for collecting images of the surface during a deep atmospheric descent.
3. Altitude variation of a balloon requires changing the density of the lifting gas. There are no commercially available pumps in the market today that have the pressure rise and flow rate capabilities needed for a Venus balloon. Most pumps or compressors are not built to be lightweight, which is of critical importance on a balloon mission.

Relevance / Science Traceability:

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.

 

JPL's Solar System Mission Formulation Office and the NASA Science Mission Directorate'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 as identified in this call.

References:

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

Scope Title:

Improved Downlink Satellite Communications for Balloons

Scope Description:

Improved downlink bit rates and innovative solutions using satellite relay communications from balloon payloads are needed. Long-duration balloon flights currently utilize satellite communications systems to relay science and operations data from the balloon to ground-based control centers. The current maximum downlink bit rate is 150 kbps, operating continuously during the balloon flight. Future requirements are for bit rates of 1 Mbps or more. Improvements in bit rate performance, reduction in size and mass of existing systems, or reductions in cost of high-bit-rate systems are needed. Tracking and Data Relay Satellite System (TDRSS) and Iridium satellite communications are currently used for balloon payload applications. A commercial S-band TDRSS transceiver and a mechanically steered 18 dBi gain antenna provide 150 kbps continuous downlink. TDRSS K-band transceivers are available but are currently cost prohibitive. Open port Iridium service is also in use, but the operational cost is high per byte transferred.

Expected TRL or TRL Range at completion of the Project: 1 to 6
Primary Technology Taxonomy:
    Level 1: TX 05 Communications, Navigation, and Orbital Debris Tracking and Characterization Systems
    Level 2: TX 05.5 Revolutionary Communications Technologies
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Phase I: Desired deliverables include: (1) results of analysis or simulation or (2) test results of actual prototype hardware and/or software.

Phase II: Deliverables could include a prototype that could be test flown on a balloon mission.

State of the Art and Critical Gaps:

Current commercially available satellite relays systems that could be used for balloon flight are either too costly or do not provide the needed downlink data rates. Tracking and Data Relay Satellite System (TDRSS) and Iridium satellite communications are currently used for balloon payload applications. A commercial S-band TDRSS transceiver and a mechanically steered 18-dBi-gain antenna provide 150 kbps continuous downlink. TDRSS K-band transceivers are available but are currently cost prohibitive. Open port Iridium service is also in use, but the operational cost is high per byte transferred.

Relevance / Science Traceability:

Science Mission Directorate (SMD) - NASA HQ (Astrophysics Division) enables multiple Research Opportunities in Space and Earth Science (ROSES) opportunities, Small Explorer (SMEX) Announcement of Opportunity (AO) (Astrophysics), Astrophysics Mission of Opportunity, and Hands-On Project Experience (HOPE) (annually).

Improvements to satellite communications for research balloons would enable greater and better data collection, possibly extended flight duration, and other such potential benefits.

References:

• NASA's SuperTIGER Balloon Flies Again to Study Heavy Cosmic Particles: https://sites.wff.nasa.gov/code820/
• GUSTO (Galactic/Extragalactic ULDB Spectroscopic Terahertz Observatory) mission is a planned high-altitude balloon mission that will carry an infrared telescope to measure emissions from the interstellar medium. The mission is being developed by NASA's Explorers Program - GUSTO, University of Arizona (Prof. Chris Walker).
• Scientific balloon information: https://sites.wff.nasa.gov/code820/technology_capabilities.html
• 2020 NASA Technology Taxonomy: https://www.nasa.gov/offices/oct/taxonomy/index.html

Scope Title:

Steerable Recovery/Parachute System

Scope Description:

NASA is 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 be 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: 1 to 6
Primary Technology Taxonomy:
    Level 1: TX 17 Guidance, Navigation, and Control (GN&C)
    Level 2: TX 17.X Other Guidance, Navigation, and Control
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype

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 includes an engineering development unit and testing with a report of the results.

State of the Art and Critical Gaps:

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 (from this point on “science gondola”) is separated from the balloon and falls to Earth on a parachute following the wind currents at the time of release and lands on cardboard crush pads.  In most cases this allows recovery of the science gondola, though 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 in water or in areas that require special equipment or are difficult for recovery (e.g., 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 entail primarily steerability, but this also results frequently in a safety analysis, which will allow more “green lights” for launch than would otherwise be the case.

Relevance / Science Traceability:

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, UAV, and aircraft programs.

References:

https://patents.google.com/patent/EP1463663A4/en

https://www.airforce-technology.com/features/featurejpads-circumventing-gps-for-next-gen-precision-airdrops-4872436/

Scope Title:

Relative Wind Speed Sensor for Scientific Balloons

Scope Description:

A trajectory control system (TCS) for high-altitude scientific ballooning has been a long-term goal of NASA’s Balloon Program Office (BPO). One milestone in the critical path of TCS development is the ability to measure the speed of the winds seen by the gondola during a balloon mission. In addition, NASA has identified wind-speed measurements from a balloon explorer under the TX10.1.2 of the 2020 NASA Technology Taxonomy (see References below). Currently, the BPO has no method of measuring relative winds (wind speed relative to the gondola) in situ above ~15 km in altitude for terrestrial applications. Although several methods of wind speed measurement exist for a variety of applications, there is effort required to port those technologies for the conventional balloon float environment. The goal of this technology development is to develop a sensor to meet the following specifications:

1. Measure relative wind in three axes (u, v, and w).
2. Operate at 4.4 mbar (~36.5 km altitude) or lower pressure.
3. Operate in air temperature from -70 to +65 °C.
4. Accuracy of 10 cm/s or better.
5. Resolution of 5 cm/s or better.
6. Sample rate of 1 Hz or faster.
7. Power consumption of 30 W or less at steady state.
8. Mass of 20 kg or less.
9. Withstand shocks of 10g or greater.

Expected TRL or TRL Range at completion of the Project: 2 to 6
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.3 In-Situ Instruments/Sensor
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware
• Software

Desired Deliverables Description: State of the Art and Critical Gaps:Relevance / Science Traceability:

A relative wind sensor for balloon missions would benefit the Science Mission Directorate (SMD)/Astrophysics mission by furthering the state of the art in sensor technology. In addition, the development of a relative wind sensor is a key milestone in the path towards a trajectory control system for high-altitude balloons. Specifically, NASA’s Super Pressure Balloon (SPB) would benefit from trajectory control while pursuing 100-day flights in the Southern Hemisphere.

References:

Scientific balloon information: https://sites.wff.nasa.gov/code820/technology_capabilities.html

2020 NASA Technology Taxonomy: https://www.nasa.gov/offices/oct/taxonomy/index.html

Lead Center: GSFC

Participating Centers: JPL, LaRC, MSFC

Scope Title:

Command, Data Handling, and Electronics

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 2021 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 subtopic 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 digital-to-analog 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. 
• Radiation-hardened ASIC devices to enable direct capture of analog waveforms.  
• Multiple-output point-of-load power regulator: This module, preferably implemented utilizing one or more controller ASICs, will source a minimum of three settable output voltages when provided with an input voltage between +5 and +12 V. Output voltages shall be independently settable to any voltage between 3.3 and 0.9 V with efficiency of at least 95%. Regulation, noise filtering, and other operational specifications should be commensurate with industry standards for space-based systems. Output current in the 10 A range to handle field-programmable gate array (FPGA) core requirements. The module should provide standard spacecraft power supply features, including overvoltage protection, fault tolerance, load monitoring, sequencing, synchronization, and soft start and should allow control and status monitoring by a remote power system controller. Using fewer external components is also highly desirable. There is also interest in a capability to provide data over power line communication to the converter for control and monitoring functions. The offeror should determine radiation-tolerance levels achievable utilizing commercially available processes and indicate, in the proposal, the radiation-tolerance goals.
• High-density high-reliability interconnections: A high-reliability connector or interconnect mechanism that can operate in space environments (vacuum, vibration) and deliver hundreds of signal/power connections while using as little physical board area as possible is desired. The connector wiring and cabling in addition to the connector shape and size should be considered in providing a complete system that further reduces the size and weight of the harnessing. The design should handle everything from carrying power to high-speed (10+ Gbps) impedance-controlled connections. The design should be scalable in different sizes to accommodate fewer connections and save board space. Low insertion force is desirable. Right angle and stacking design options should be considered.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.3 In-Situ Instruments/Sensor
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware
• Software

Desired Deliverables Description:

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

Desired Phase II deliverables:

• For mixed-signal structured ASIC platforms—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.
• For the direct analog waveform capture ASIC—should include a prototype ASIC device implemented on a test board and demonstration of the wave form capture capabilities of the device.
• For the multiple output point of load switcher—a prototype multi-output point of load regulator. The regulator should be integrated onto a test board and be performance tested under varying resistive, capacitive, and transient load conditions.
• For the high density high-reliability interconnect—prototypes of the connection system (different size, orientations, wiring, etc.). The connector should be integrated onto a test board where its performance (speed, cross talk, etc.) can be verified. 

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.
 

There are multiple output point-of-load converters available from commercial companies. The existing commercial parts require many external components, eliminating their space savings. Commercial parts are not built on radiation-tolerant processes.
 

Current connectors and interconnect harnessing are too large, especially for small satellites and CubeSats. As the size of the printed circuit boards has shrunk, the percent of board space being used by the input/output (I/O) connectors has become unacceptable. The connectors are taking away from circuitry and sensors that could be providing additional functionality and science products. High-density commercial connectors tend to be lacking in their general ruggedness, outgassing, and ability to prevent intermittent connections in high-vibration environments like orbital launches.

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 planets 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.
 

Multi-output point-of-load converters and high-density high-reliability interconnects are relevant to miniaturizing electronics. Miniaturized flight electronics allows one to fit more functionality into less volume, allowing smaller spacecraft to perform science that was previously done by larger satellites. These missions include interplanetary CubeSats and SmallSats, outer planets instruments, and heliophysics missions.

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 Centers: GRC, GSFC

Scope Title:

Critical Technologies for Sample-Return Missions

Scope Description:

This Subtopic focuses on robotic sample-return (SR) missions that require landing on large bodies (e.g., Luna, Mars Sample Return (MSR)), 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. The challenges associated with these SR missions may be grouped into four categories: (1) Mass-efficient spacecraft architectures (e.g., efficient propulsion or materials that significantly reduce the mass of the launch payload required), (2) Sample handling (e.g., subsurface acquisition mechanisms), (3) Sample integrity (e.g., surviving reentry), and (4) Planetary protection/contamination control (PP/CC) (e.g., preventing leakage into the orbital sample (OS) canister). This Subtopic seeks potential solutions to areas (1), (3), and (4), considering it best that technologies associated with (2), sample handling, be directed to Subtopic S4.02. The intent is to have this Subtopic S4.03 manage only those technologies in areas (1), (3), and (4) that are specifically related to SR missions; technology solutions related to other classes of missions should instead be directed to Subtopics S4.04 (Extreme Environments) and S4.05 (Contamination Control and Planetary Protection).

 

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 require three heavy launch vehicle launches with rough spacecraft mass of 5,000 kg each in order 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 lander mass, for example, would yield many times more savings in the launch payload mass, enhancing the feasibility of these missions.

 

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 capsule 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 analyses, 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 that are expected to return samples in December 2020 from asteroid Ryugu and in September 2023 from asteroid Bennu, respectively. The first segment of NASA's MSR mission is the sample collection rover Perseverance, launch of which took place in July 2020. The MSR sample retrieval segment (lander, fetch rover, Mars Ascent Vehicle) is scheduled to begin Phase A development in October 2020 for a 2026 launch. The third MSR segment will be ESA's Earth return vehicle (ERV).

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 acquisition, sample integrity, and planetary protection.

This solicitation seeks proposals that have the potential to increase the Technology Readiness Level from 3 or 4 to 6 within 5 years, and 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 ten "Priority Questions" drawn from three Crosscutting Science Themes, with "Missions in the Recommended Plan that Address Them". SR missions are shown to address eight out of the ten 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://www.jpl.nasa.gov/missions/mars-sample-return-msr/

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

Lead Center: JPL

Participating Centers: 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:

1. Very low temperature environments (e.g., temperatures at the surface of Titan and of other Ocean Worlds as low as -180 °C; and in permanently shadowed craters on the Moon).
2. Combination of low-temperature and radiation environments (e.g., surface conditions at Europa of -180 °C with very high radiation).
3. Very high temperature, high pressure, and chemically corrosive environments (e.g., Venus surface conditions, having very high pressure and 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-arc-second/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/ultra-low-power, wide-operating-temperature, low-noise mixed-signal electronics for space-borne 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).

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.

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 include proof-of-concept working prototypes that demonstrate the innovations defined in the proposal and enable direct operation in extreme environments. 

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.

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 (electronics, 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 Centers: None

Scope Title:

CC and PP Implementation and Verification

Scope Description: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: proof-of-concept study for the approach to include data validation and modeling. 

 

Phase II: detailed analysis and prototype for testing.

 

Areas to consider for deliverables are, 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 leverages the technologies resulting from the 1960s to 1970s Viking spacecraft assembly and test era. The predominant means to control biological contamination on spacecraft surfaces is using some combination of heat microbial reduction processing and solvent cleaning (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, the hardware then is protected in a cleanroom environment (ISO 8 or better) using protective coverings when hardware is not being assembled or tested. Biological cleanliness is then verified through the NASA standard assay, which is a culture-based method. 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. 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 developed approaches for compliance, environmental assessments are implemented to understand recontamination potential for cleanroom surfaces and air. Although the NASA standard assay is performed on the cleanroom surfaces, DNA-based methodologies have been adopted to include 16S and 18S ribosomal-ribonucleic-acid- (rRNA-) targeted sequencing, while metagenomic approaches are currently undergoing development. Thus, 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 low biomass and compounds (e.g., viable organisms, DNA) but also compliant with cleanroom and electrostatic discharge limits, quantification of the widest spectrum of viable organisms, enhanced microbial reduction/sterilization modalities that are compatible with flight materials, and a ground- and flight-based recontamination 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 3.0×10¹⁵ g/cm²/s with massspectrometry, under flight conditions (low and high operating temperatures) and with combined exposure to natural environment (high-energy radiation, ultraviolet radiation, atomic oxygen exposure).
• Particulate transport modeling and analysis for continuum, rarefied, and molecular flow environments with electrostatic, vibro-acoustic, particle detachment and attachment capabilities.
• Modeling and analysis of molecular return flux using direct simulation Monte Carlo (DSMC) and the Bhatnagar–Gross–Krook (BGK) formulations.

Relevance / Science Traceability:

Protection requirements has emerged in recent years with increased interest in investigating bodies with the potential for life detection such as Europa, Enceladus, Mars, etc. and the potential for sample return from such bodies. The development of such technologies would enable missions to be able to be responsive to PP requirements as they would be able to assess viable organisms and establish microbial reduction technologies to achieve acceptable microbial bioburden levels for sensitive life detection instruments to prevent inadvertent “false positives,” to ensure compliance sample return planetary protection and science requirements, and to provide a means to comply with probabilistic-based planetary protection requirements for biologically sensitive missions (e.g., outer planets and sample return). 

References:

Planetary  Protection, https://planetaryprotection.nasa.gov/

 

Handbook for the Microbial Examination of Space Hardware, https://explorers.larc.nasa.gov/2019APSMEX/SMEX/pdf_files/NASA-HDBK-6022b.pdf

 

Lead Center: JPL

Participating Centers: GSFC

Scope Title:

High-Performance Space Computing Technology

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

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 in 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 provide the performance, the power-to-performance ratio, or the 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 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.

In-Situ Resource Utilization (ISRU) involves any hardware or operation that harnesses and utilizes ‘in-situ' resources (natural and discarded) to create products and services for robotic and human exploration.  ISRU encompasses a broad range of systems, and is typically divided into six focus areas:  Resource Assessment, Resource Acquisition, Resource Processing/Consumable Production, In Situ Manufacturing, In Situ Construction, and In-Situ Energy.  ISRU products and services can be used to reduce Earth launch mass or lander mass by not bringing everything from Earth, reduce risks to the crew and/or mission by reducing logistics, increasing shielding, and providing increased self-sufficiency, or reduce costs by needing less launch vehicles to complete the mission and/or through the reuse of hardware and lander/space transportation vehicles.  Since ISRU can be performed wherever resources may exist, ISRU technologies and systems may need to operate in a variety of environments and gravities, and may need to consider a wide variety of potential resource physical and mineral characteristics. This year's solicitation will focus on critical technologies needed in the areas of Resource Acquisition and Consumable Production for the Moon and Mars.

Lead Center: JSC

Participating Centers: GRC, JPL, KSC, MSFC

Scope Title:

Solar Concentrator Technologies for Oxygen Extraction and In Situ Construction

Scope Description:

Solar concentrators have been used to successfully demonstrate multiple in situ resource utilization (ISRU) technologies, including hydrogen and carbothermal reduction, sintering of regolith to produce launch/landing pads, and production of blocks for construction. Terrestrial state-of-the-art solar concentrators are heavy, not designed for easy packaging/shipping and assembly/installation, and can be maintained and cleaned on a periodic basis to maintain performance. For ISRU space applications, NASA is interested in solar concentrators that are able to be packaged into small volumes, are lightweight, easily deployed and set up, can autonomously track the Sun, and can perform self-cleaning operations to remove accumulated dust. Materials, components, and systems that would be necessary for the proposed technology must be able to operate on the lunar surface in temperatures of up to 110 ^oC (230 ^oF) during sunlit periods and as as low as -170 ^oC (-274 ^oF) during periods of darkness. 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 regolith mining operations will be tens of meters away from the solar concentrators, but that regolith processing systems and solar concentrators will be co-located on a single lander. Phase I efforts can be demonstrated at any scale; Phase II efforts must be scalable up to 11.1 kW of delivered solar energy, assuming an incoming solar flux of ~1,350 W/m² while also considering volumetric constraints for launch and landing. Each of the following specific areas of technology interest may be developed as a standalone technology.

• Lightweight mirrors/lenses: Proposals must clearly state the estimated W/kg for the proposed technology. Phase II deliverables must be deployed and supported in Earth 1g (without wind loads) but should include design recommendations for mass reductions for lunar gravity (1/6g) deployment. Proposals should address the following attributes: high reflectivity, low coefficient of thermal expansion, strength, mass, reliability, and cost.
• Efficient transmission of energy for oxygen/metal extraction: While the solar concentrator will need to move to track the Sun, reactors requiring direct thermal energy for oxygen extraction will be in a fixed position and orientation. Concentrated sunlight must be directed to a single or multiple spots to effectively heat or melt the regolith. Proposals must define the expected transition losses from collection to delivery and should capture any assumptions made regarding the distance from collection to delivery.
• Sintering end effector: Solar concentrators have been used to demonstrate the fabrication of 3D printed components using regolith as the only feedstock. Proposals responding to this specific technology area must produce and maintain a focal point temperature between 1,000 and 1,100 ^oC for the purpose of sintering lunar regolith. Proposals should assume that the focal point can move along the regolith at a speed between 1 and 10 mm/sec.

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:

• Prototype
• Analysis
• Hardware

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. See Scope Description for additional information on Phase I and Phase II deliverables. 

State of the Art and Critical Gaps:

The 2011 paper "Thermal Energy for Lunar In Situ Resource Utilization: Technical Challenges and Technology Opportunities" [Ref. 1] summarized the work performed in this area and recommends future efforts focus on lightweight mirrors (possibly using composite materials) and dust mitigation techniques (dust mitigation is addressed in another subtopic).

The last solar concentrator system developed for ISRU had an overall efficiency of ~33%. The performance of the system is captured in the 2011 paper "Solar Thermal System for Lunar ISRU Applications: Development and Field Operation at Mauna Kea, HI" [Ref. 6].

Relevance / Science Traceability:

NASA Strategic Knowledge Gap (SKG) 1-F, "Determine the likely efficiency of ISRU processes using lunar simulants in relevant environments," as well as NASA SKG 1-G, "Measure the actual efficiency of ISRU processes in the lunar environment," are both important for the development of future ISRU systems. There are multiple ISRU processes that involve the use of solar concentrators, and determining their efficiency through technology development efforts may address NASA SKGs. 

References:

1. Gordon, P. E., Colozza, A. J., Hepp, A. F., Heller, R. S., Gustafson, R., Stern, T., & Nakamura, T. (2011). Thermal energy for lunar in situ resource utilization: Technical challenges and technology opportunities. https://ntrs.nasa.gov/citations/20110023752
2. Gustafson, R., White, B., Fidler, M., & Muscatello, A. (2010). Demonstrating the solar carbothermal reduction of lunar regolith to produce oxygen. In 48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition (p. 1163).
3. Mueller, R. P., Sibille, L., Hintze, P. E., Lippitt, T. C., Mantovani, J. G., Nugent, M. W., & Townsend, I. I. (2014). Additive construction using basalt regolith fines. In Earth and Space 2014 (pp. 394-403). https://ntrs.nasa.gov/citations/20150000305
4. Muscatello, A., & Gustafson, R. B. (2010). The 2010 Field Demonstration of the Solar Carbothermal Reduction of Regolith to Produce Oxygen. https://ntrs.nasa.gov/citations/20110006938
5. Muscatello, T. (2017). Oxygen Extraction from Minerals. https://ntrs.nasa.gov/citations/20170001458
6. Nakamura, T., & Smith, B. (2011, January). Solar thermal system for lunar ISRU applications: Development and field operation at Mauna Kea, HI. In 49th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition (p. 433).

Scope Title:

Novel Oxygen Extraction Concepts

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 silicates using various techniques. The target production rates are 1,000 kg of O₂ per year for a lunar pilot plant, and 10,000 kg of O₂ per year for a lunar full-scale plant. Each of the following specific areas of technology interest may be proposed as individual efforts to support existing oxygen extraction development projects.

• Contaminant Removal: Proposed concepts should be capable of removing 0.36 g of HCl, 0.68 g of HF, and 0.1 g of H₂S per kg of processed regolith from a mixed gas stream of CO, CO₂, and H₂ in a way that minimizes the use of consumables. Phase I efforts should provide an estimated mass/power as a function of contaminant quantities. Phase II efforts should demonstrate the technology using actual gases.
• Regolith Inlet/Outlet Valves: Proposed concepts should be capable of passing abrasive granular material through the valve for at least 1,000 cycles and should be actuated with a type of motor that has flight heritage (e.g., brushless direct current (BLDC) motors or stepper motors). Phase I efforts should provide an estimated mass and power for the concept through analysis and/or demonstration. Phase II efforts should demonstrate the technology using lunar regolith simulant and collect data to predict leak rates for up to 10,000 cycles.
• Contamination-Tolerant Vacuum Pump: Some in situ resource utilization (ISRU) processes may require a pressurized volume to be evacuated in order to prevent the loss of products and consumables to the vacuum of space when regolith either enters or exits the volume. The pump may be exposed to corrosive substances such as HCl, HF, and H₂S. Proposed concepts should be capable of evacuating a volume of 50 L with an initial pressure of 5 psia down to a pressure of <5 torr at the pump inlet in <2 min while compressing the gases to 1 atm at the pump outlet. Phase I efforts should provide an estimated mass, power, and life for the concept. Phase II efforts should demonstrate the technology using actual gases.

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:

See Scope Description for definitions of Phase I and II deliverables for each technology.

State of the Art and Critical Gaps:

The carbothermal reduction process was demonstrated at a relevant scale using an automated reactor in 2010. Multiple efforts are underway to bring carbothermal reduction technology to TRL6. Other techniques that use ionic liquids, molten salts, and molten regolith electrolysis 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:

The Space Technology Mission Directorate (STMD) 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. A lunar water prospecting mission is required to properly assess the utilization potential of water on the lunar surface. Until water prospecting data becomes available, NASA recognizes the need to make progress on the technology needed to extract oxygen from dry lunar regolith. 

References:

1. Fox, E. T. (2019). Ionic Liquid and In Situ Resource Utilization. https://ntrs.nasa.gov/citations/20190027398
2. 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.
3. 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
4. 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).
5. Muscatello, A., & Gustafson, R. B. (2010). The 2010 Field Demonstration of the Solar Carbothermal Reduction of Regolith to Produce Oxygen. https://ntrs.nasa.gov/citations/20110006938
6. Muscatello, T. (2017). Oxygen Extraction from Minerals. https://ntrs.nasa.gov/citations/20170001458
7. Paley, M. S., Karr, L. J., & Curreri, P. (2009). Oxygen Production from Lunar Regolith using Ionic Liquids. https://ntrs.nasa.gov/citations/20090017882
8. 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. One challenge with extracting the water is that desorption and sublimation can occur at temperatures as low as 150 K. The inverse challenge exists with water collection. Unless the water vapor is under pressure, extremely cold temperatures will be necessary to capture it. NASA is seeking methods to acquire lunar water ice from PSRs. Proposals must describe a method for extracting and/or collecting lunar water ice that exists at temperatures between 40 and 100 K and 10-9 torr vacuum.

• Phase I demonstrations can be at any scale, but eventually the technology must be able to demonstrate an average rate of 2.78 kg H₂O/hr (15 metric tons of water in 225 days).
• Phase II demonstrations can be subscale, but must define the number of subscale units necessary to achieve an average extraction rate of 2.78 kg H₂O/hr.
• Proposals should state expected energy requirements (both electrical and thermal).
• Proposers should assume a mobile platform is considered to be available, but should not be necessary for technology demonstration.
• Proposers should state their assumptions about water ice concentration.
• Proposals should describe a tolerance for a trace amount of organics or volatiles that may accumulate on collection surfaces.
• Proposers should estimate Wh/kg H₂0 for concepts and/or provide a plan to determine that value as part of the effort.
• Proposers should address the ability of a concept to be able to operate for at least 1 year, with a goal of 5 years without substantial maintenance.

Estimates for mass and volume of the final expected hardware should be specified.

In addition, each of the following specific areas of technology interest may be proposed to support existing efforts related to lunar ice mining.

• Regolith/Ice Excavation: Proposed concepts should be able to excavate frozen regolith simulant with a water ice content of at least 5% by mass while minimizing a temperature increase in the excavated material. Phase I efforts should provide an estimated mass/power for the excavation concept as well as an estimate for any temperature increase in the frozen regolith caused by the excavation technique. Phase II efforts should demonstrate the technique with lunar simulant at a target production rate of 0.28 kg H₂0/hr and collect data to predict the estimated wear over time.
• Regolith/Ice Crushing: Proposed concepts should be able to crush frozen regolith simulant with a water ice content of at least 5% by mass while minimizing a temperature increase in the excavated material. Phase I efforts should provide an estimated mass/power for the crusher concept as well as an estimate for any temperature increase in the frozen regolith caused by the crushing technique. Phase II efforts should demonstrate the technique with lunar simulant mixed with ice having an initial unconfined compressive strength of 10 MPa at a target production rate of 0.28 kg H₂0/hr and collect data to predict the estimated wear over time. 
• Subsurface Volatile Extraction: Proposed concepts should be able to release volatiles at a depth of 50 cm below the surface with a water ice content of at least 5% by mass. Phase I efforts should provide an estimated mass/power for the concept. Phase II efforts should demonstrate the technique with lunar simulant at a target production rate of 0.28 kg H₂0/hr and collect data to predict the estimated wear over time if applicable.

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:

See Scope Description for definitions of Phase I and II deliverables for each technology.

State of the Art and Critical Gaps:

Scoops and bucket-wheel excavators have been demonstrated for the collection of unconsolidated material but may not be effective at excavating consolidated regolith-ice composites. The Planetary Volatiles Extractor (PVEx) developed by Honeybee Robotics is the state of the art for heated core drills, but life testing is required to determine the rate of wear due to repeated excavation. Multiple groups have investigated the use of thermal mining methods to separate water from regolith, but the depth of water removed is relatively shallow. Very little work has been performed on the ability to capture water in a lunar environment after it has been released from the surface.

Relevance / Science Traceability:

The current NASA Administrator has referenced water ice as one of the reasons we have chosen the lunar poles as the location to establish a sustained human presence. 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:

1. 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.
2. 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.
3. 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.
4. Andreas, E. L. (2007). New estimates for the sublimation rate for ice on the Moon. Icarus, 186(1), 24-30.

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 some technologies have application to commercial space capabilities.

Robust, efficient, and predictable EDL systems fulfill the critical function of delivering payloads to planetary surfaces through challenging environments, within mass and cost constraints. Future NASA missions will require new technologies to break through historical constraints on delivered mass, or to go to entirely new planets and moons. Even where heritage systems exist, no two planetary missions are exactly “build-to-print,” so there are frequently issues of environmental uncertainty, risk posture, and resource constraints that can be dramatically improved with investments in EDL technologies. New capabilities and improved knowledge are both important facets of this focus area.

Because this topic covers a wide area of interests, subtopics are chosen to enhance and or fill gaps in the existing technology development programs. Future subtopics will support one or more of four broad capability areas, which represent NASA's goals with respect to planetary Entry, Descent and Landing:

• High Mass to Mars Surface
• Precision Landing and Hazard Avoidance
• Planetary Probes and Earth Return Vehicles
• EDL Data Return and Model Improvement

A cross-cutting set of disciplines and technologies will help mature these four capability areas, to enable more efficient, reliable exploration missions. These more specific topics and subtopics may include, but are not limited to:

• Thermal Protection System materials, modeling, and instrumentation
• Deployable and inflatable decelerators (hypersonic and supersonic)
• Guidance, Navigation, and Control sensors and algorithms
• Aerodynamics and Aerothermodynamics advances, including modeling and testing
• Precision Landing and Hazard Avoidance sensors
• Multifunctional materials and structures

This year the Entry, Descent and Landing focus area is seeking innovative technology for:

• Deployable Decelerator Technologies
• EDL Sensors, including those embedded in thermal protection systems and those used for proximity operations and landing
• Hot Structure Technology for Atmospheric Entry Vehicles
• Lander Systems Technology

The specific needs and metrics of each of these specific technology developments are described in the subtopic descriptions.

Lead Center: MSFC

Participating Centers: AFRC, JSC, LaRC

Scope Title:

Hot Structures Technology for Aerospace Vehicles

Scope Description:

This subtopic deals with the development of reusable nonmetallic hot structure technology for structural components exposed to extreme heating environments on aerospace vehicles. Desired hot structure systems encompass multifunctional structures that can reduce or eliminate the need for active cooling or 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), improved aerodynamics for aeroshell components, improved structural efficiency, and increased ability for nondestructive inspections. Hot structure is an enabling technology for reusability between missions or mission phases, such as 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. Examples of prior flight-proven hot structures include: (a) the nozzle extension for the Centaur RL10B-2 upper-stage rocket engine, and (b) 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, and lightweight fiber-reinforced hot structure technology applicable to aerospace vehicles and components exposed to extreme temperatures. At a minimum, materials developed under this subtopic should be capable of operating at a temperature of at least 1,371 °C (2,500 °F)—higher temperatures are of even greater interest, such as up to 2,204+ °C (4,000+ °F). 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. The material systems of interest for use in developing hot structure technology include advanced 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, 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. Focus areas should address one or more of the following:

• 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.

• 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 protection that enhance durability are also of interest: matrix inhibition, oxidation resistant matrices, exterior environmental coatings, etc.

• 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, 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 demonstrations. Phase I feasibility studies should also address cost and the risks associated with the hot structures technology.

 

In addition to the final report, delivery of 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—for example: (a) coupons appropriate for thermal and/or mechanical material property tests, or (b) arc-jet test specimens. Plans for potential 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 test articles should be a part of the anticipated Phase II effort. Depending upon the emphasis of the Phase II work, such test articles may include subscale nozzle-extension test articles or arc-jet test specimens/hot structure components. 

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 reuse. Critical gaps or technology needs include:(a) increasing operating temperatures to 1,649 to 2,204+ °C (3,000 to 4,000+ °F); (b) increasing resistance to environmental attack (primarily oxidation); (c) increasing manufacturing technology capabilities to improve reliability, repeatability, and quality control; (d) increasing durability/toughness and interlaminar mechanical properties (or introducing 3D architectures); (e) decreasing cost, and (f) decreasing overall manufacturing time required.

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 to various space destinations, such as the lunar South Pole, is also of great interest.
 

The Advanced Exploration Systems (AES) Program would be ideal for further funding a prototype hot structure system and technology demonstration effort. Commercial Space programs, such as Commercial Orbital Transportation Services (COTS), Commercial Lunar Payload Services (CLPS), 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 include the Space Launch System (SLS) and the Human Landing System (HLS) for propulsion applications.
 

Potential NASA users of this technology exist for a variety of propulsion systems, 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.

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. Navy, the U.S. Army, the Missile Defense Agency (MDA) and the Defense Advanced Research Projects Agency (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 Hot 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 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 noted, both foreign and domestic.

Lead Center: JSC

Participating Centers: ARC, GSFC, JPL, LaRC

Scope Title:

High-Accuracy, Lightweight, Low-Power Fiber-Optic or Recession Sensing System for Thermal Protection Systems and Low-Cost Data Acquisition System

Scope Description:

Current NASA state-of-the-art entry, descent, and landing (EDL) instrumentation and associated data acquisition are very expensive to design and incorporate on planetary missions because they must meet functional and performance requirements during and after exposure to loads and environments associated with spaceflight and atmospheric entry.

 

Commercial fiber-optic systems offer an alternative to traditional sensors that could result in a lower overall cost and weight reduction while actually increasing the number of measurements. Fiber-optic systems are also immune to electromagnetic interference (EMI), which reduces design and qualification efforts. This would be highly beneficial to future planetary missions requiring thermal protection systems (TPSs). In addition, as NASA looks to the future of science missions to the outer planets, extreme entry environments will require the new, 3D-woven Heatshield for Extreme Entry Environment Technology (HEEET) TPS recently matured within the Agency. Gathering flight performance data on this new material will be key—particularly the measurement of recession, which was so very important on the Galileo probe mission to Jupiter. Minimizing the sensor intrusion of the outer mold line is critical in this case because the extreme environment dictates that the TPS be as aerothermally monolithic as possible. In applications to planetary entry vehicles greater than about 1 m in diameter, however, the HEEET TPS is expected to contain seams that might be used for accommodating instrumentation.

 

Recession measurements in carbon fiber/phenolic TPSs such as Phenolic Impregnated Carbon Ablator (PICA) and AVCOAT are also of interest. When ablation is not severe and/or rapid, accurate measurements have proven difficult with the historic Galileo-type sensor, which was based on the differential resistance resulting from sensor materials that have charred.

 

To be considered against NASA state-of-the-art TPS sensing systems for future flight missions, fiber-optic systems must be competitive in sensing capability (measurement type, accuracy, and quantity) and associated data acquisition system mass, size, power, and cost. Therefore, NASA is looking for a fiber-optic system that can meet the following requirements:

• TPS Temperature
  • Measurement Range: -200 to 1,250 ºC (up to 2,000 ºC is preferred).
  • Accuracy: +/-5 ºC desired.
• Surface Pressure
  • Measurement Range: 0 to 15 psi.
  • Accuracy: < +/-0.5%.

 

Destinations such as Mars, Venus, and Titan pose many challenges for EDL data acquisition systems, including radiation, g-loading, and volume constraints. Recent notable examples of such systems are the Mars Entry, Descent, and Landing Instrument (MEDLI) sensor suite, which successfully acquired EDL data in 2012, and the upcoming MEDLI2 system, which will gather data during EDL at Mars in February of 2021. The NASA MEDLI and MEDLI2 data systems are 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. In an effort to bring EDL instrumentation to all missions, NASA is seeking a low-cost, robust, high-accuracy data acquisition system that can meet the following requirements:

• Performs instrument signal conditioning and analog-to-digital conversion, and includes a spacecraft bus serial interface.
• Weight: 5 kg or less.
• Size: Modularity encouraged; maximum module size 10 cm³; 4 modules maximum.
• Power: 16 W or less.
• Measurement Resolution: 12-bit or higher.
• Accuracy: +/-0.5% of full-scale output.
• Acquisition Rate per Measurement: 8 Hz or higher.
• 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).
• Operating Temperature Capability: -40 °C to 85 °C.
• Cost: Fully qualified target of ~$1M (recurring).
• Sensor Compatibility.
  • Minimum 15 thermocouples with at least 2 Type R and 8 Type K.
  • Minimum 8 pressure transducers (120 or 350 ohm bridge).

 

For recession measurements acquired in extreme entry environments requiring 3D woven TPSs, NASA is seeking novel concepts that fit into the sensor/electronics architecture described above and meet the following requirements:

• Up to 5,000 W/cm² total heat flux (convective plus radiative).
• Up to 5 atmospheres of pressure on the vehicle surface.
• Minimum recession measurement accuracy within +/-1 mm.

 

For recession measurements in moderate entry environments requiring carbon fiber/phenolic TPSs, NASA is seeking novel concepts that fit into the sensor/electronics architecture described above, and meet the following requirements:

• 150 to 2,000 W/cm² total heat flux (convective plus radiative).
• Up to 1 atmosphere of pressure on the vehicle surface.
• Minimum recession measurement accuracy within +/-1 mm.

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

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.

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 TPS performance could lead to reduced design margins, enabling a greater payload mass-fraction and smaller landing ellipses. Improved real-time measurement knowledge during entry could also minimize the landing dispersions for placing advanced payloads onto the surface of atmospheric and airless bodies. 

 

NASA science missions are frequently proposed that include high-speed Earth return (New Frontiers, Discovery, and Mars Sample Return) and Venus and Mars entry. Capsules used for these missions must withstand both convective and radiative aeroheating, and NASA now requires EDL instrumentation for these missions. Current radiative measurement techniques (radiometers) provide only an integrated heating over a limited wavelength range; past interpretation of such flight data [Ref. 3, 4] shows the need for spectrally resolved measurements from spectrometers. For Earth and Venus, the radiative component may be the dominant source of heating, and emission comes from the vacuum ultraviolet (VUV), which NASA currently has no capability to measure. For Mars and Venus, the aftbody radiation is dominated by midwave infrared (MWIR). Again, NASA does not have a method to measure MWIR radiation in flight; the current radiometers integrate across several band systems. Miniaturized spectrometers that can measure in VUV and MWIR would have immediate application to Science Mission Directorate (SMD) planetary missions. Such spectrometers may also inform what ablation species are emitted from the heat shield and backshell during entry. 

References:

1. M. Munk, A. Little, C. Kuhl, D. Bose, and J. Santos, "The Mars Science Laboratory (MSL) Entry, Descent and Landing Instrumentation (MEDLI) Hardware," Proc. AAS/AIAA Space Flight Mechanics Conference, AAS 13-310, Kauai, HI, 2013.
2. F. Milos, "Galileo Probe Heat Shield Ablation Experiment," Journal of Spacecraft and Rockets, Vol. 34, No. 6, Nov-Dec 1997.
3. B. A. Cruden and C. O. Johnston, "Characterization of EFT-1 Radiative Heating and Radiometer Data," 46th AIAA Thermophysics Conference, Washington, D.C., June 2016.
4. A. Brandis, C. O. Johnston, B. A. Cruden, D. Prabhu, and D. Bose, "Uncertainty Analysis and Validation of Radiation Measurements for Earth Reentry," Journal of Thermophysics and Heat Transfer, Vol. 29, No. 2, 2015, pp. 209-221.

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) Guidance, Navigation, and Control (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 either provide terrain mapping (range point cloud) capability or direct velocity measurement. The continued maturation of these technologies is targeting (1) further size, mass, and power reductions of components; (2) multicomponent integration; and (3) multimodal operation (i.e., combing mapping and velocimetry functions). 

 

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:

• Dense focal plane arrays for simultaneous ranging and Doppler velocimetry, plus associated signal processing approaches including photonic integrated circuits (PICs), with the following characteristics:
  • Simultaneous measurements from each pixel or from subsets of pixels.
  • Functionality (when integrated into a lidar system) that would operate up to 8 km range.
  • Functionality (when integrated into a lidar system) for measuring velocity from 0 m/sec along the line of sight (LOS) up to 200 m/sec or greater.
  • PICs approaches that integrate multiple components into a single device or provide a single component in a miniaturized robust package (e.g., master laser, modulator, and detectors).
  • Ability to reject false locks on dust plumes due to exhaust.
  • Implementation for low power, mass, and size.
  • Optical losses comparable with fiber-optic or bulk optical components.
• High-speed (5 MHz) wavelength tuning laser modules with low power driving electronics, which have random wavelength access or predefined wavelength-lookup-table tuning, and meet the following requirements:
  • Semiconductor laser module.
  • Tuning range: 1550 nm with tuning range of C-band.
  • Tuning speed: 5 MHz, and less than 100 ns settling time.
  • Wavelength grid: 10,000 evenly distributed over the whole tuning range.
  • Tuning fashion: Random wavelength grid access, or sequential predefined wavelength lookup table tuning.
  • High wavelength and power repeatability.
  • Low temperature or environmental dependency.

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:

The EDL GN&C and sensors community has been developing for more than a decade 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 near-term science missions, such as Mars 2020, are starting to infuse some of the PL&HA capabilities.

References:

1. A. Martin et al., "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.
2. C.V. Poulton, A. Yaacobi, D.B. Cole, M.J. Byrd, M. Raval, D. Vermeulen, and M.R. Watts, "Coherent solid-state LIDAR with silicon photonic optical phased arrays," Opt. Lett. 42, 4091-4094 (2017).
3. F. Amzajerdian, G.D. Hines, D.F. Pierrottet, B.W. Barnes, L.B. Petway, and J.M. Carson, “Demonstration of coherent Doppler lidar for navigation in GPS-denied environments,” Proc. SPIE 10191, Laser Radar Technology and Applications XXII, 1019102 (2017).

Lead Center: LaRC

Participating Centers: ARC

Scope Title:

Entry and Descent System Technologies

Scope Description:

Background: 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. They 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 also enables delivery of very large (20 metric tons or more) usable payload, which will likely be needed to support human exploration. The technology also allows for reduced cost access to space by enabling the recovery of launch vehicle assets. This subtopic area solicits innovative technology solutions applicable to deployable entry concepts. Specific technology development areas include:

1. Advancements in textile manufacturing technologies that can be used to simplify production, reduce the mass, or reduce the stowed volume of mechanically deployed structures, inflatable structures, or their flexible thermal protection system. Thermal protection concepts can also lead to improvements in thermal management efficiency of radiant and conductive heat transport at elevated temperatures (exceeding 1,200 °C). Concepts can be either passive or active dissipation approaches. For smaller scale inflatable systems for small-spacecraft/satellite applications, less than 5 m in diameter, thin-ply or thin-film manufacturing approaches that can be used to reduce the minimum design gauge are of particular interest for inflatable structures. 
2. High-temperature-capable structural elements to support mechanically deployable decelerators that surpass the performance capability of metallic ribs, joints, and struts. Anticipated systems would include composite elements or hybrid approaches that combine metallic structures with high-temperature-capable interface materials to improve thermal performance. Minimum-mass approaches that address volumetric/packing efficiencies at small-scale (approx. 1 to 2 m) implementations are of interest for small-satellite applications.
3. Development of gas-generator technologies used as inflation systems that result in improved mass efficiency and system complexity over both current pressurized cold gas systems and present state-of-the-art gas generators for inflatable structures. 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 the overall system mass versus amount of gas generated.

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 for Phase I development should be material coupon up to subscale manufacturing demonstration articles that demonstrate proof of concept, and lead to Phase II manufacturing scaleup and testing in relevant environments for applications related to Mars entry, Earth return, launch asset recovery, or the emergent small-scale satellite community.  

State of the Art and Critical Gaps:

The current state of the art for deployable aerodynamic decelerators is limited due to the novelty of this technology. Developing more efficient, lighter, and thinner flexible thermal protection system component materials with higher temperature capability could potentially enable more efficient designs and extend the maximum range of use of the concepts. Novel and innovative high-temperature structural concepts are needed for the mechanically deployed decelerator. Development of gas generator technologies that improve mass efficiency over current pressurized cold gas systems for inflatable structures is needed. 

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. The Human Exploration and Operations Mission Directorate (HEOMD), Space Technology Mission Directorate (STMD), and Science Mission Directorate (SMD) 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.
• Olds, A. D., et al., “IRVE-3 Post-Flight Reconstruction,” AIAA Paper 2013-1390.
• 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.

Lead Center: MSFC

Participating Centers: GRC, LaRC

Scope Title:

Landing Systems Technologies

Scope Description:

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

 

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:

1. Ground test data, test techniques, and diagnostics across physical scales and environments, with particular emphasis on nonintrusive approaches and methodologies.
2. PSI-specific flight instrumentation, with particular emphasis on in situ measurements of particle size and particle velocity during the landing phase.
3. Solutions to alleviate or mitigate the PSI environments experienced by propulsive landers.
4. Validated 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 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 2020, 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.  

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 dataset, 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 post-landing 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 Lander 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. ISRU implications for lunar and martian plume effects. 47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition. 2009.

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. International Space Exploration Coordination Group Assessment of Technology Gaps for Dust Mitigation for the Global Exploration Roadmap. AIAA SPACE 2016. 2016. 5423.

Lead Center: ARC

Participating Centers: LaRC

Scope Title:

Optical and Laser-Spectroscopic Imaging Techniques for High-Enthalpy Arc-Heated Test Facilities

Scope Description:

Arc-heated high-enthalpy test facilities at NASA’s Ames and Langley Research Centers are used for evaluation and certification of high-temperature materials and structures of an entry vehicle’s thermal protection system (TPS). Future exploration missions will utilize new ablative TPS materials that release decomposition products into the gas stream ahead of the vehicle, influencing flow-field behavior. Data and observations from materials testing programs using NASA’s arc jet facilities are critical for validation of high-fidelity modeling and simulation tools used to design and margin TPS specifications for entry vehicles. However, the complex multiphysics processes that manifest as entry aeroheating of ablative TPSs present formidable challenges for model validation. The available diagnostic techniques for arc jet testing provide little direct evidence of the subject aerothermal and thermophysical processes.

 

NASA is seeking advanced and new optical and laser-spectroscopic techniques applied to arc jet testing programs. Experimental methods for arc jet facility characterization strive to quantify thermodynamic and gas dynamic properties of the arc jet stream and serve multiple purposes, such as verification of test conditions (facility operations), validation of arc heater and flow-field simulations, and measurement of incident/boundary conditions for material response simulations. Of equal importance are methods that can detect and identify pyrolysis gases and particles injected into the shocked gas region ahead of TPS material test articles, providing needed insight to the complex interactions of the flow field with material response. Experimental methods that measure recession, temperature, and optical properties of the TPS surface enable characterization of surface thermal response phenomenology.

 

The off-body gas phase diagnostics are to detect and quantify:

• Major species in the arc jet stream (N, O, N_2, NO for air; CO and CO₂ for facilities capable of operating with CO₂ mixtures).
• Ablation species and recombination products in the shock layer (C, CN, CH, H, Ca).
• Spalled particles from test articles penetrating the shock layer.

Also of importance are measurements of velocity and free stream and shock layer temperatures, including vibrational temperature. Planar or line imaging techniques are desired to characterize spatial distributions with 1-mm or smaller resolution at kHz data rates. Burst mode (>100 kHz) imaging approaches that enable correlation of temporal-spatial intermittencies are of particular interest.

 

The requested surface imaging diagnostics are to measure test-article temperature and spectral emissivity, topology, and recession rate. Hyperspectral techniques are preferred if they enable characterization of multiple surface properties simultaneously while discriminating from shock layer radiation. Spatial resolutions and acquisition rates of <1 mm and >30 Hz, respectively, are desired. Adaptation of standoff surface spectroscopy techniques may hold promise for time-resolved species detection and identification.

 

Spallation characterization requires measurements of ejected particle size distributions, 2D and 3D trajectories, and velocity distributions. Techniques for both stagnation and shear testing configurations are desired. Imaging spatial resolution and field of view needs to account for particle trajectories that travel upstream and penetrate shock waves. Methods that provide insight to the chemical composition of spalled particles would be particularly valuable. Anticipated particle size and speeds are 1 to 100 µm and 1 to 200 m/s, respectively.

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 09.1 Aeroassist and Atmospheric Entry
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware

Desired Deliverables Description:

Phase I: Assessment study of potential diagnostic techniques.
Phase II: Prototype instrument demonstration in relevant environment with hardware delivery to NASA.

State of the Art and Critical Gaps:

The requirements for spatially resolved, species-specific measurements of high-temperature-reacting gas properties necessitate the use of optical-spectroscopic methods. The state of the art at NASA’s arc-heated facilities are techniques based on nanosecond-pulsed laser-induced fluorescence (ns-LIF). Pointwise flow property measurement requires approximately 60 to 90 sec per acquisition to recover properties calibrated to engineering units (velocity, temperature, species densities) at one location. The low sensitivity and long standoff distances in arc-heated facilities preclude line or planar imaging. The long acquisition times result in a poor use of expensive testing resources. The ns-LIF approach is not suitable for post-shock and near-surface regions, as the moderate pressures, high temperatures, and strong gas and surface luminosity confound signal interpretation, effectively prohibiting quantitative ablation species detection. Emission spectroscopy techniques for free-stream and shock-layer measurements have been used with success for many years, but these are limited to observation of excited-state populations along integrated lines of sight. Tunable laser absorption spectroscopy can target ground-state or excited-state populations of certain species of interest; however, the low-density/high-temperature test conditions and the short path lengths yield absorbances too low for detection with demonstrated laser absorption techniques.

 

Recent advances in tunable amplified sources pumped by kHz and burst-mode femtosecond-pulse lasers has enabled the development of several nonlinear laser spectroscopy techniques. These new techniques have capabilities that directly address the sensitivity, temporal, and spatial resolution shortcomings identified above.

 

For spallation diagnostics, both active interrogation/response and passive imaging techniques have potential for characterizing size distributions, trajectories, and velocity distributions. NASA’s current capability for observing and quantifying spallation is the use of high-speed gated video, which is limited to situations where particle luminosity can be distinguished from background sources.

Relevance / Science Traceability:

Several potential future missions, outlined in decadal surveys, crewed exploration mission studies, and other supporting analyses, have ED/EDL (entry and descent or entry, descent, and landing) architectures: Mars Sample Return, high-speed crewed return, high-mass Mars landers, Venus, and gas/ice giant probes. With few exceptions, entry vehicle TPS for these missions will be composed of materials currently under development and without certification heritage in multiple gases. Arc jet testing at conditions relevant for certification will invariably be required for each of these proposed missions. Ground testing at more extreme environments for future missions will challenge existing capabilities. There is a compelling need now to bring research-level diagnostic technologies forward to ensure that facility operations can confidently demonstrate required performance to TPS technology projects.

Conventional instrumentation will continue to be the primary source of facility characterization data. The purposes of the advanced techniques are to provide validating evidence for the conventional instrumentation, reveal error and bias in interpretation of heat flux measurements, and ultimately reduce uncertainty in facility performance data provided to test programs.

NASA planetary exploration programs supporting ED/EDL missions are the intended beneficiaries of this subtopic. The first-line project is the Space Technology Mission Directorate's Entry Systems Modeling Project.

References:

1. Entry Systems Modeling Project: https://gameon.nasa.gov/projects/entry-systems-modeling-esm/
2. O. Chazot, “Experimental Studies on Hypersonic Stagnation Point Chemical Environment,” RTO-EN-AVT-142, Experiment, Modeling and Simulation of Gas-Surface Interactions for Reactive Flows in Hypersonic Flights, pp. 13-1 to 13-32.
3. A. Gülhan, “Heat Flux Measurements in High Enthalpy Flows,” RTO EN-8, Measurement Techniques for High Enthalpy and Plasma Flows, April 2000.
4. J. Grinstead, et al., “Consolidated laser-induced fluorescence diagnostic systems for the NASA Ames arc jet facilities,” AIAA 2016-4159.
5. S. Bailey, et al., “Experimental analysis of spallation particle trajectories in an arc-jet environment,” Experimental Thermal and Fluid Science v. 93, pp. 319-325 (2018).
6. D.A. Codron, “Emission Spectroscopy Characterization of Thermal Protection System Materials in Arc-Heated Flows,” AIAA 2014-2112.
7. J.A. Inman, et al., “Nitric Oxide PLIF Measurements in the Hypersonic Materials Environmental Test System (HYMETS),” AIAA Journal Vol. 51, No. 10, pp. 2365-2379, October 2013.
8. D.J. Bamford, et al., “Characterization of the H1 Arc Heater Facility by Laser-Induced Fluorescence,” AIAA 2013-0035.
9. P. Danehy, et al., “Quantitative Spectral Radiance Measurements in the HYMETS Arc Jet,” AIAA 2012-0856.
10. M. Winter, et al., “Non-Equilibrium Analysis of Emission Spectroscopy Data Taken in the Freestream of the NASA IHF Arc Jet Facility,” AIAA 2015-2963.

Scope Title:

Advanced Instrumentation for NASA's Shock Tube and Ballistic Range Facilities

Scope Description:

NASA Ames Research Center operates two specialized-use impulse test facilities for aerodynamic and aerothermodynamic research investigations that support atmospheric entry systems modeling and validation. The Electric Arc Shock Tube (EAST) facility replicates shocked gas environments encountered by entry vehicles transiting planetary atmospheres at hypersonic velocities. EAST creates these environments through calibrated gas mixtures of target atmospheres (Earth, Venus, Mars, Titan, gas giants, etc.) and prescribed preshock pressures and shock speeds. Radiative heating of both fore- and after-body surfaces of an entry vehicle can influence, and in some cases dominate, the design requirements for thermal protection systems. Spectroscopic instrumentation is used to characterize the absolute radiance and gas kinetics behind a traveling shock wave. The Hypervelocity Free Flight Aerodynamic Facility (HFFAF), an enclosed aeroballistic range, is used for the study of dynamically similar supersonic and hypersonic aerodynamics, transition to turbulence, and laminar and turbulent convective heat transfer. Optical imaging instrumentation is used to characterize aerodynamic forces and moments of scaled models launched through the range. Thermographic and spectral imaging instrumentation is used to characterize spatially resolved heating rates to scaled models.

 

NASA is seeking innovative imaging and spectroscopic measurement techniques for these two facilities. New electro-optic products and methods will enable measurement of quantities unattainable with current capabilities as well as improve current practices.

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 09.1 Aeroassist and Atmospheric Entry
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware

Desired Deliverables Description:

Phase I: Assessment study of potential diagnostic techniques or technology upgrades.
Phase II: Prototype instrument demonstration in relevant environment (preferably with hardware delivery to NASA).

State of the Art and Critical Gaps:

The EAST facility’s instrumentation acquires data for shocked gas phenomenology and facility performance characterization. Measurements of radiance, absorbance, electron density, and temperature are used for validation of comprehensive radiation transport simulations of planetary atmospheres. Those measurements are primarily acquired using calibrated optical-spectroscopic instruments with sufficient temporal and/or spatial resolution to correlate observed magnitudes with localized, spectrally resolved absolute radiant fluxes or columnar property densities (including electron densities). Ancillary instrumentation is used to measure shock arrival times and transient pressures at the tube wall to establish shock speeds adjacent to the science instruments.

 

Measurement techniques that correlate observables to atomic and molecular state populations and radiance magnitudes enable validation of radiance models. Emission spectroscopy techniques, which capture the transient characteristics of excited atomic and molecular state populations, have reached a high degree of maturity and performance.

 

However, post-shock electron and ground-state or other dark-state population dynamics also influence shock radiance. Measurement of these states rely on more complicated absorption, induced fluorescence, or scattering (spontaneous and coherent) techniques. The lack of light sources and/or detectors with suitable spectral and temporal characteristics, or the challenges of implementation in impulse facilities, have limited the opportunities for such measurements. Techniques that enable measurement of these states would greatly expand opportunities for radiation transport model validation, particularly for conditions in which self-absorption would influence emission spectroscopy measurements. Specific quantities of interest are rotational/translation and vibrational temperatures, electron temperatures, and ground-state population densities of N₂, N₂+, N, O, CO, CN, H₂, H.

 

Spatiotemporal resolution is necessary to discern nonequilibrium relaxation processes behind a traveling shock wave and is the key requirement for EAST diagnostics. Field imaging at a single instant in time, or time-resolved imaging of a single point in space, must have resolutions equivalent to >1 mm and >1 µs to capture these relaxation processes.

 

For the HFFAF, shadowgraph and schlieren photography are used to provide time-resolved imagery for aerodynamic force and moment analyses of scaled flight vehicles in free flight. A high-speed shutter (40-ns duration) and a spark-gap light source enable images to be captured without motion blur or image fogging from shocked gas radiance enveloping a test model. The shuttering system relies on Kerr cells filled with benzonitrile and a 35 kV pulse-shaping and switching network. Advances are sought for the eventual replacement of the 32 heritage light source/shutter systems with components that offer equal or greater performance as well as improved safety and reliability.

Relevance / Science Traceability:

Several potential future missions outlined in decadal surveys, crewed exploration mission studies, and other supporting analyses have ED/EDL (entry and descent or entry, descent, and landing) architectures: Mars sample return, high-speed crewed return, high-mass Mars landers, and Venus and gas/ice giant probes. Entry vehicles to these destinations will encounter radiative heating to varying degrees. Radiative heating of a vehicle’s backshell has been recognized as a significant concern, so ensuring a full range of diagnostic techniques for expanding flows has become a high priority for the EDL community.

Characterizing the aerodynamic stability of emerging deployable drag devices for entry vehicles is also of high importance for future high-mass lander missions. The HFFAF will be a key ground-test facility for acquiring crucial free-flight aerodynamic data for study and simulation validation.

NASA planetary exploration programs supporting ED/EDL missions are the intended beneficiaries of this subtopic. Technology development projects supporting these programs are potential beneficiaries of new instrumentation for the EAST and HFFAF.

References:

1. Entry Systems Modeling Project: https://gameon.nasa.gov/projects/entry-systems-modeling-esm/
2. ADEPT (Adaptable Deployable Entry and Placement Technology) Project: https://gcd.larc.nasa.gov/projects-2/deployable-aeroshell-concepts-and-flexible-tps/
3. Many journal papers, conference proceedings, and technical reports describing the NASA Ames EAST and HFFAF test facilities and research are available in the open literature.

From the smallest satellite to the most complicated human rated spacecraft, thermal is seen as an enabling function to a vehicle. Temperatures must be maintained within design limits, whether those be cryogenic systems for science instruments, or comfortable shirt sleeve operations temperatures for crew missions. As missions evolve and waste energy rejection becomes more of a demand, NASA seeks components for both active and passive thermal systems. Such components complete the thermal cycle which includes waste energy acquisition, transport, rejection/storage, and insulation. The ultimate goal for any thermal control hardware is to minimize mass, volume, and power while maintaining the aforementioned temperature limits on a spacecraft.

Lead Center: GSFC

Participating Centers: 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 provides 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

• Successfully develop the formulations of the coating that leads 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 its mechanical durability test under the influence of simulated space and lunar environmental conditions.
• Deliverable of 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-relating project and projects involved with robotic science rovers and landers.

References:

• The following website provides links to some references for dust mitigation coatings such as lotus thermal coatings: https://ntrs.nasa.gov/search.jsp?R=20150020486
• The following website provides links to some 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 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.
• Physics-based analysis or model.
• Proof-of-concept hardware (Phase I).
• Proof-of-concept hardware tested against simulated loads in proposed environments (Phase II).
• Final report.

State of the Art and Critical Gaps:

Specifically, heat pump systems are needed for 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.
• System should be tolerant of being powered down during the lunar night and restarted during the day reliably over multiple diurnals.
• Exported vibrations, if any, should be minimal for compatibility with science instruments.

Novel heat pump systems are desired. Enabling improvements over 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 spacecraft. 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 in order 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.
• 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.

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. In order 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 ICESat (Ice, Cloud, and Land Elevation Satellite), 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 SMD (Science Mission Directorate) 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 nonconsumable/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, etc.

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.
 

While interest in lunar science and the development of abilities to deliver payloads to the lunar surface are 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 2020, 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 Centers: GRC, GSFC, JPL, MSFC

Scope Title:

Spacecraft Thermal Management

Scope Description: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. 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. Additionally, the exploration of embedded thermal control technologies may have a direct impact on the AMPS program in the near term and provide valuable insight into techniques for other embedded cooling applications. These improvements may come through either novel hardware solutions or modernization of software tools used to assess human vehicle interactions. The current state of the art in thermal control results in 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 control (both actively and passively) within the thermal control system becomes more apparent. 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:

• Advanced Modular Power System
• Europa Clipper/Lander
• Lunar Lander
• Long-duration habitats (Moon, Mars, etc.)
• EVA

References:

1. Kandilian, R., Bhandari, P., & Mastropietro, A. J. (2018). Thermal and Flow Analysis of Europa Clipper Thermal Control Valves. NASA Thermal and Fluids Analysis Workshop. TFAWS18-AT-14.
2. Hartenstine, J., Walker, K., and Anderson W. (2012). Passive Control of a Loop Heat Pipe with Thermal Control Valve for Lunar Lander Application. AIAA 2012-354
3. Stephan, R. (2011). Overview of the Altair Lunar Lander Thermal Control System Design and the Impacts of Global Access. AIAA 2011-5001.
4. Ewert, M.K. (1993). Investigation of Lunar Base Thermal Control System Options. SAE Transactions. J. of Aerospace. 102(1). 829-840.
5. Nyberg, K. L., Diller, K. R., & Wissler, E. H. (2001). Model of human/liquid cooling garment interaction for space suit automatic thermal control. Journal of biomechanical engineering. 123(1). 114-120.
6. Pallo, N., et al., (2018). Modular heat sink for chip-scale GaN transistors in multilevel converters. 2018 IEEE Applied Power Electronics Conference and Exposition (APEC). San Antonio, TX. 2798-2805. doi: 10.1109/APEC.2018.8341414.
7. Colozza, A. J., &Gardener, B. G. (2019, May 1). Advanced Modular Power System Electronics Enclosure Thermal Testing. NASA Glenn Research Center. Cleveland, OH. NASA/TM-2019-220011.
8. Izenson, M., Phillips, S., Chepko, A., Quinn, G., Steele, J., & Bue, G. (2016, July). Design of a Lithium Chloride Absorber Radiator for Flight Testing on an Extravehicular Mobility Unit. 46th International Conference on Environmental Systems.
9. Bue, G., Makinen, J., Vogel, M., Tsioulos, G., Honas, M., Dillon, P., & Porwitz, D. (2011, July). Hollow Fiber Flight Prototype Spacesuit Water Membrane Evaporator Design and Testing. 41st International Conference on Environmental Systems.

The concept of distributed spacecraft missions (DSM) involves the use of multiple spacecraft to achieve one or more science mission goals. Small distributed spacecraft acting in cooperation can execute science and exploration missions that would be impossible by traditional large spacecraft operating alone and offer the potential for new concepts in mission design. The goal of this topic is to develop enabling technologies for small spacecraft DSM configurations operating over large distances beyond low Earth orbit (LEO). The term DSM or “swarm” refers to a group of cooperatively distributed spacecraft, scalable up to 100s of spacecraft, in a specific configuration, which has three distinct characteristics. First, as opposed to a constellation, where spacecraft are distributed across multiple orbits, a DSM is comprised of small spacecraft orbiting relatively close to one another, with intersatellite ranges on the order of tens to hundreds of kilometers. Second, the DSM requires inter-spacecraft communications where each spacecraft is capable of sharing data and relative position information so that all swarm members are aware of the overall topology. The swarm topology would be dependent on the spatial and temporal distribution, orbit, ground reference, or other requirements of the science mission. Third, the swarm is commanded from the ground as an entity rather than each spacecraft individually. Thus, the swarm has inherent autonomous capabilities to control individual or complete swarm topology redistribution depending or requirements or in response to commands. Small spacecraft, for the purpose of this solicitation, are defined as those with a mass of 180 kilograms or less and capable of being launched into space as an auxiliary or secondary payload. Small spacecraft are not limited to Earth orbiting satellites but might also include interplanetary spacecraft, planetary re-entry vehicles, and landing craft. Specific innovations being sought in this solicitation will be outlined in the subtopic descriptions. Proposed research may focus on development of new technologies but there is particular interest in technologies that are approaching readiness for spaceflight testing. NASA's Small Spacecraft Technology Program will consider promising SBIR technologies for spaceflight demonstration missions and seek partnerships to accelerate spaceflight testing and commercial infusion. Some of the features that are desirable for small spacecraft technologies across all system areas are the following: 

• Simple design.
• High reliability.
• Tolerant of extreme thermal and/or radiation environments.
• Low cost or short time to develop.
• Low cost to procure flight hardware when technology is mature.
• Small system volume or low mass.
• Low power consumption in operation.
• Suitable for rideshare launch opportunities or storage in habitable volumes (minimum hazards).
• Able to be stored in space for several years prior to use.
• High performance relative to existing system technology. 

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

• www.nasa.gov/smallsats https://www.nasa.gov/smallsat-institute 

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

• http://www.nasa.gov/sites/default/files/atoms/files/small_spacecraft_technology_state_of_the_art_2015_tagged.pdf

Lead Center: GRC

Participating Centers: ARC, GSFC, JPL

Scope Title:

End-to-End Deep Space Communications

Scope Description:

Develop enabling communications technologies for small spacecraft beyond 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 envisioned science missions 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 data relay from lunar surface to surface, 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 architecture 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 of extension of the technologies and capabilities 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-C hardware for radio-frequency (RF) and optical cross links. 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 cross links and networking protocols that scale to hundreds of nodes, and require robustness for loss of nodes or as new nodes enter the network. 

 

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, 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 05.2 Radio Frequency
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 (10s to 100s of km 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 0.2m² 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].
• 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.

Relevance / Science Traceability:References:
Scope Title:

Relative and Absolute Deep Space Navigation

Scope Description:

Develop enabling technologies for beyond low Earth orbit (LEO) relative and/or absolute position knowledge. This situational awareness allows for autonomous control of small spacecraft as well as determining and maintaining position within a swarm or constellation of small spacecraft. In addition, timing distribution solutions for the SmallSats are important. Earth-independent and Global Positioning System- (GPS-) independent navigation and timing are enabling capabilities required by spacecraft to conduct NASA lunar and deep space distributed spacecraft science missions. Innovations in navigation 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.

 

Multiple small spacecraft operating in coordinated orbital geometries or performing relative stationkeeping can further expand human knowledge deeper into the universe by performing coordinated occultation, acting as virtual telescopes, and forming distributed apertures that would be prohibitively complex and expensive to launch into space as monolithic structures. Small-spacecraft formation flight can also enable swarm gravimetry, synchronized observation of transient phenomena, and proximity operations for inspection of other assets. Realizing these capabilities on affordable small spacecraft requires sensors and maneuvering systems that are low in mass, volume, power consumption, and cost.

 

Further expansion of small spacecraft use into deep space requires highly accurate position knowledge and precision timing that does not depend on GPS or other Earth-centric aids. Exploration mission operations that involve multiple-element distributed-mission architectures may involve 30 to 100 spacecraft, and the general expansion of the number of cislunar and deep space missions will stress or exceed current capacity of the Deep Space Network (DSN). Access to DSN ranging may not be available for multiple concurrent missions, may be blocked by terrain for surface operations, or may be limited by the radio capabilities of smaller missions. In concert with other available signals of opportunity and landed beacons, small spacecraft can provide relative ranging or triangulation to aid lunar navigation. Knowledge at the spacecraft of relative (between-spacecraft) situational awareness is needed for real-time stationkeeping/relative position control where required rapid reaction speeds preclude human-in-the-loop operation.

 

Future small-spacecraft missions will need to autonomously determine and transmit relative and absolute position as well as keep and exchange precise timing. These capabilities are required for small spacecraft to act as infrastructure for other missions and for distributed missions composed of small spacecraft beyond Earth. Navigation technologies and techniques may include inertial navigation combined with enhanced visual navigation capabilities (e.g., dual use of star-tracking instruments for relative navigation using surface features or other nearby spacecraft), x-ray emissions (from pulsars), and laser rangefinding to other spacecraft or surface landmarks. For use with small spacecraft, these systems must be compatible with the inherent size, weight, power, and cost (SWaP-C) constraints of the platforms.

Precise timekeeping and timing exchange is not only required for navigation but is fundamental to science data collection. Internetworked small spacecraft can help synchronize timing across multiple mission assets using an external timing source. Improvements in chip-scale atomic clocks that can be carried by the small spacecraft themselves can augment this capability to reduce the accumulation of errors over time or serve as the primary clock when other larger but more accurate reference sources are not available or feasible. The vast majority of current commercial interests and Government missions operate in near-Earth orbits. To date, both NASA and the commercial spaceflight industry have enjoyed strong investment in near-Earth situational awareness made possible by tracking and identification capabilities provided by the Department of Defense. As the number of cislunar missions grows and NASA encourages the development of the lunar service economy, similar investments in situational awareness capabilities in these new orbital regimes will be needed to help support NASA and commercial operations.

 

Primary applications include navigational aids to lunar surface and orbiting users. Distributing these capabilities across multiple SmallSats may be necessary because of limited SWaP, but also to enhance coverage. Technologies for specific lunar architecture are especially needed, but considerations of extension to the martian domain are also solicited. Navigation solutions for deep space distributed spacecraft missions (DSMs) may be addressed via hardware or software solutions or a combination thereof. 

Expected TRL or TRL Range at completion of the Project: 2 to 5
Primary Technology Taxonomy:
    Level 1: TX 17 Guidance, Navigation, and Control (GN&C)
    Level 2: TX 17.2 Navigation Technologies
Desired Deliverables of Phase I and Phase II:

• Prototype
• Software
• Hardware

Desired Deliverables Description:

Phase I: Identify and explore options for the deep space navigation technology, conduct trade analysis and simulations, define operating concepts, and provide justification for proposed techniques, frequency bands of operation, command and data handling, and networking solutions. Also identify, evaluate, and develop design for integrated navigation payload(s) and one or more constituent technologies that enable distributed spacecraft operations in the relevant space environment beyond LEO. Integrated navigation 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 navigation 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:

Science measurements of distributed satellite missions (DSMs) are based on temporal and spatially distributed measurements where position knowledge and control are fundamental to the science interpretation. Current space navigation technologies are not adequate when relative or absolute position knowledge of multiple spacecraft are involved. State of the art (SOA) for attitude is the Jet Propulsion Laboratory's ASTERIA (Arcsecond Space Telescope Enabling Research in Astrophysics) 6U CubeSat demonstrated pointing stability of 0.5 arcsec (0.1 microdeg) rms over 20 min using guide stars. For position knowledge, missions still primarily use ranging transponders relying on a two-way Earth link. Examples of SOA for this ranging are the Iris transponder and the Small Deep Space Transponder (SDST) [Ref. 13].

Global navigation satellite services like the United States' Global Positioning System (GPS) provide very limited services beyond Geostationary Earth Orbit distances, and no practical services in deep space. Autonomous navigation capabilities are fundamental to DSMs to ensure known topography of the configuration at the time of data acquisition. Control of the distributed configuration requires robust absolute and relative position knowledge of each spacecraft within the configuration and the ability to control spacecraft position and movement according to mission needs. Critical areas for advancement are:

• Long-term, high-accuracy attitude determination: In particular, low-SWaP absolute attitude determination using star trackers, etc., to achieve sub-arcsec accuracy.
• Optical navigation: Solutions are sought for visual-based systems that leverage advances in optical sensors (e.g., cameras, star trackers) to observe and track a target spacecraft and perform pose and relative position estimation. Opportunities for innovation include methods that do not require the execution of satellite maneuvers and/or the design of external satellite features that enhance observability. Innovations may be appropriate for only certain regimes, such as near, medium, or far range; however, this context should be described. Solutions for various lunar and deep space mission operations concepts are of interest.
• Other novel navigation methods: Stellar navigation aids, such as navigation via quasars, x-rays, and pulsars, may provide enabling capabilities in deep space. Surface-based navigation aids, such as systems detecting radio beacons or landmarks, are invited.
• Methods for autonomous position control are also of interest. Technologies that accomplish autonomous relative orbit control among the spacecraft are invited. Control may be accomplished as part of an integrated system that includes one or more of the measurement techniques described above. Of particular interest are autonomous control solutions that do not require operator commanding for individual spacecraft. That is, control solutions should accept as input swarm-level constraints and parameters and provide control for individual spacecraft. Opportunities for innovation include the application of optimization techniques that are feasible for small satellite platforms and do not assume particular orbit eccentricities. State-of-the art in this area is the Cislunar Autonomous Positioning System Technology Operations and Navigation Experiment (CAPSTONE), the first spacecraft to attempt to navigate to and maintain a near-rectilinear halo orbit around the Moon as a precursor for Gateway [Ref. 11].  NASA is also partnering with universities for use of surface-feature-based navigation and timing [Ref. 12].

NOTE: Small-spacecraft propulsion technologies are not included in this subtopic.

Relevance / Science Traceability:

Space communications and position knowledge and control are enabling capabilities required by spacecraft to conduct all NASA missions. The concept of distributed spacecraft missions (DSMs) involves the use of multiple spacecraft to achieve one or more science mission goals.

Several missions are being planned to conduct investigations/observations in the cislunar region and beyond. For example, CLPS; human exploration (Artemis) landing site and resource surveys; and lunar communication and navigation infrastructure, including LunaNet, Mars communications relay, etc.  All of these missions will benefit from improved communications and navigation capabilities.

References:

[1] International Communication System Interoperability Standard (ICSIS): https://www.internationaldeepspacestandards.com

[2] Interagency Operations Advisory Group (IOAG): https://www.ioag.org/

[3] Space Communication Architecture Working Group (SCAWG) (2006) NASA Space Communication and Navigation Architecture Recommendations for  2005-2030: https://www.nas.nasa.gov/assets/pdf/techreports/2006/nas-06-014.pdf

[4] NASA Delay/Disruption Tolerant Networking (DTN): http://www.nasa.gov/content/dtn

[5] "Delay-tolerant networking: an approach to interplanetary internet." IEEE Communications Magazine 41, no. 6 (2003): 128-136.

[6] National Telecommunications and Information Administration Frequency Allocation Chart: https://www.ntia.doc.gov/files/ntia/publications/january_2016_spectrum_wall_chart.pdf

[7] National Telecommunications and Information Administration Tables of Frequency Allocations: https://www.ntia.doc.gov/legacy/osmhome/alloctbl/alloctbl.html

[8] NASA Spectrum Policy and Guidance for Small Satellite Missions: http://www.nasa.gov/directorates/heo/scan/spectrum/policy_and_guidance.html

[9] NASA Space Communications and Navigation networks: https://www.nasa.gov/directorates/heo/scan/services/networks/index.html

[10] NASA Optical Communications: https://www.nasa.gov/directorates/heo/scan/opticalcommunications/overview

[11] Cislunar Autonomous Positioning System Technology Operations and Navigation Experiment (CAPSTONE) https://www.nasa.gov/press-release/nasa-funds-cubesat-pathfinder-mission-to-unique-lunar-orbit

[12] On-Orbit Demonstration of Surface Feature-Based Navigation and Timing: https://www.nasa.gov/feature/ames/nasa-selects-universities-for-collaborative-development

[13] NASA/TP-2018-220027, State of the Art Small Spacecraft Technology Report: https://www.nasa.gov/smallsat-institute/sst-soa

Lead Center: ARC

Participating Centers: GSFC, JPL

Scope Title:

Technologies to Enable Cost and Schedule Reductions for Optical System for CubeSats

Scope Description:

Concepts for optical systems are sought that will enable larger apertures or longer focal lengths than currently available systems, to be deployed from within small spacecraft. Relatively inexpensive small spacecraft offer several advantages over larger, more expensive spacecraft: small spacecraft can perform inspection and repair of larger spacecraft, several can be deployed for more frequent revisit rates over Earth's surface or planetary objects, and multiple craft can achieve affordable mission reliability through redundancy. To date, the utility of small spacecraft in missions involving remote sensing (in any spectral band) has been constrained by their low budget and compact size; optical sensitivity is limited in proportion to the diameter of a telescope's aperture, and magnification is limited by the effective focal length. The cost to produce one-of-a-kind optical assemblies is disproportionate and the production times too long to incorporate into the tight budgets and schedules typical of small-spacecraft missions.

 

The objective of this subtopic is to receive proposals that articulate a demonstrable ability to manufacture, test, and control ultra-low-cost observing systems that can meet the reference mission performance requirements (including infrastructure issues) within a time frame and budget compatible with a small-spacecraft development cycle. For the purposes of this subtopic, small spacecraft are defined as CubeSats of 12U volume. Proposals are sought that will

• Specify observing systems figures of merit for a potential small-spacecraft mission, for example:
  • Earth resource management (commercial).
  • Maritime traffic monitoring.
  • Observations for agricultural industry.
  • Fire, flood, or other emergency monitoring.
  • Lunar exploration precursors or observation of human activity at the Moon.
  • Remote spacecraft health inspection.
  • Near-Earth object detection.
  • In-space or lunar-to-Earth optical communications.
  • Other reference mission to be specified by proposer.
• Include discussion of current state of the art for telescope optical parameters (sensitivity, resolution, and magnification within a spectral band).
• Include production cost and schedule significantly improved by the proposed system design.

Significant areas for proposals are:

• Concept systems that are modular in nature that can be produced in quantities larger than the single-unit production typical of spaceflight hardware builds (for instance, batch or lot production yielding flight units in the quantity range of 30 to 50).
• Concepts that will enable large deployable apertures (optical and related, such as sun shades enabling larger apertures) and/or longer focal lengths than currently available systems that can be implemented from within a 12U small spacecraft. The concepts of large deployable optical apertures and focal lengths for small spacecraft address the fact that small spacecraft are inherently size constrained.

Requirements to be addressed in the project should include specific needs for each wavelength application region, for example:

• For UV/Optical:
  • Wavefront Figure <5 nm rms
  • Wavefront Stability <1 nm/10 min
  • First Mode Frequency >500 Hz
  • Actuator Resolution <1 nm rms
• For EUV:
  • Slope <0.1 microrad

Also needed is ability to fully characterize surface errors and predict optical performance.

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:

• Prototype

Desired Deliverables Description:

In Phase II, the project could build a prototype and complete environmental qualification testing of the optical system (or a single node in the case of a multi-element system), including measuring optical figure before and after vibration testing, acoustic testing, and thermal cycling. It would also demonstrate that the telescope maintains optical figure in a reference thermal environment including thermal gradients.

 

A successful mission-oriented Phase II would yield a credible plan to deliver (in Phase III) flight hardware within the allocated budget for a fully assembled and tested telescope assembly that can be integrated into the potential mission. This plan would demonstrate an understanding of how the engineering specifications of their system meet the performance requirements and operational constraints of the mission (including mechanical and thermal stability analysis). Cost and schedule goals and optical performance goals are listed under State of the Art and Critical Gaps.

 

Requirements to be addressed in the deliverables should include specific needs for each wavelength application region, for example:

• For UV/Optical:
  • Wavefront Figure <5 nm rms
  • Wavefront Stability <1 nm/10 min
  • First Mode Frequency >500 Hz
  • Actuator Resolution <1 nm rms
• For EUV:
  • Slope <0.1 microrad

Also needed is ability to fully characterize surface errors and predict optical performance.

State of the Art and Critical Gaps:

Technical Challenges
Ultrastable, normal incidence mirrors with low mass-to-collecting area ratios, affordably produced and delivered, modular and readily integrated into CubeSat-class form-factor, are desired. 

Affordably Manufactured, Easily Integrated, Readily Available

After performance, affordability is the most important metric for an advanced optical system, and long telescope fabrication times add significant program cost. Current normal incidence space telescopes in the 0.2- to 0.5-m aperture class have lead times of 12 to 18 months and cost $1 million to $5 million. This research effort seeks a 10× reduction in schedule and cost for precision optical components: a lead time of 4 to 6 months and a cost of $100K to $500K for a 0.2- to 0.5-m aperture-class telescope. Options should be offered for modular and easy installation into a CubeSat-style payload enclosure, with considerations for maximizing aperture sizes, reliable deployment (if required), and reliable optical alignment.

 

Large Deployable Apertures for Small Spacecraft

Small spacecraft are inherently size constrained. Given the tight volume constraints of CubeSats and other small spacecraft, deployable systems for these platforms need to be highly volumetrically efficient and may employ novel configurations or deployment mechanisms relative to their larger brethren. Systems that can deploy a larger aperture then nominally available in a fixed system can address this constraint.

 

Affordably Manufactured, Easily Integrated, Readily Available

To accomplish NASA CubeSat-class missions, the mirrors and even entire optical assemblies must be delivered on CubeSat-class schedules after they have been specified. Earth-observing missions and astronomical applications often involve assembling and testing one or many spacecraft within a matter of months from concept to delivery. Optics that can be quickly procured as "catalog items" upon specification must be fabricated and ready for installation—preferably as an assembled module including optics bench and mounting hardware "plug and play" ready—or risk not being available on time for the tight mission schedules.

Relevance / Science Traceability:

A new class of low-cost, optically stable, wide-spectral-range telescopes designed specifically for small spacecraft have application in a variety of exploration, commercial, and science missions. Existing missions can be accomplished in novel and more affordable ways with small spacecraft, and new missions will be enabled by high-performance telescopes in small spacecraft. A few examples include:

• Earth resource management.
• Maritime traffic monitoring.
• Observations for agricultural industry from low Earth orbit.
• Satellite optical crosslinks or lunar satellite-to-Earth optical communications.
• Lunar exploration in situ resource utilization (ISRU) or landing site surveyors or manned surface mission operations observation.
• Remote spacecraft health inspection/monitoring.
• Near-Earth object detection or exoplanet transit detection in deep space.

References:

Types of large sparse systems are described in this SPIE paper: Large sparse aperture space optical systems, Aden B. Meinel, Marjorie P. Meinel, 1 August 2002, Optical Engineering, 41(8), (2002). https://doi.org/10.1117/1.1490557

Nautilus Observatory: a space telescope array based on very large aperture ultralight diffractive optical elements, Dániel Apai, Tom D. Milster, Dae Wook Kim, Alex Bixel, Glenn Schneider, Benjamin V. Rackham, Rongguang Liang, Jonathan Arenberg, Proceedings Volume 11116, Astronomical Optics: Design, Manufacture, and Test of Space and Ground Systems II; 1111608 (2019). https://doi.org/10.1117/12.2529428

Lead Center: MSFC

Participating Centers: AFRC, GRC

Scope Title:

Small Spacecraft Transfer Stage Development

Scope Description:Expected TRL or TRL Range at completion of the Project: 4 to 6
Primary Technology Taxonomy:
    Level 1: TX 01 Propulsion Systems
    Level 2: TX 01.1 Chemical Space Propulsion
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware
• Software
• Analysis

Desired Deliverables Description:

A Phase I effort should provide evidence in the feasibility of key elements of cost, assembly, integration, and operations through fabrication or testing demonstrations. A prototype system should reach a "near CDR" level during Phase I with a mapping of key performance parameters (mass, power, cost, etc.) from the prototype to the flight design, along with potential opportunities for technology demonstration and commercialization.

It is highly desired that the Phase II deliverable include demonstration test of the prototype system along with detailed metrics (mass, power, cost, etc.) traceable to a flight design for the reference mission. Efforts leading to Phase II delivery of integrated prototype systems that could either be ground- or flight-tested as part of a post-Phase II effort are of particular interest.

State of the Art and Critical Gaps:

Many CubeSat/SmallSat propulsion units are designed for low delta-V maneuvers such as orbit maintenance, stationkeeping, or reaction control. Larger delta-V systems are employed for larger satellites and science/exploration missions, but are often costly and integrated as part of the satellite design. Systems typically range from cold-gas to bipropellant storables with electric systems also viable for very small systems. Rocket Lab has recently introduced an upgraded version of their monopropellant kick stage, which includes a bipropellant engine, advanced attitude control, and power subsystems. This system will be used for the first time for NASA's CAPSTONE mission and is suggested to have capability for orbits beyond the lunar environment. At the component level, Aerojet Rocketdyne and Moog, Inc. are prominent suppliers of state-of-the-art (SOA) thrusters, including commonly used variants of the R-4D engine, while companies like Blue Canyon Technologies offer spacecraft bus solutions absent dedicated propulsion elements. Advanced manufacturing, electric pumps and actuators, nontoxic propellants, and electrospray thrusters all offer potential improvements in the flight capabilities of small propulsion systems. System concepts that enable improved spacecraft performance and control, such as dual-mode systems, provide potential advancements to the current SOA, especially those that enable new science missions and those that offer potential improvements to the efficiency, accuracy, and safety of future lunar manned missions. While many of these component technologies are reasonably mature, no integrated system capability has been developed and implemented specifically as a rapid, low-cost solution for translunar or cislunar mission designs.

Relevance / Science Traceability:

This subtopic extends the capabilities of the Flight Opportunities Program and Launch Services Program by seeding potential providers to establish lunar/cislunar transfer capabilities. The Small Spacecraft Technology Program (SSTP) also seeks demonstrations of technical developments and capabilities of small spacecraft to serve as precursor missions (such as landing site investigation or in situ resource utilization (ISRU) prospecting) for human exploration, and as communications and navigation infrastructure for follow-on cislunar missions. SSTP CAPSTONE is an example mission.

 

Many technologies appropriate for this topic area are also relevant to NASA's lunar exploration goals. Small stages developed in this topic area would also be potential flight testbeds for cryogenic management systems, wireless avionics, or advance guidance systems and sensors. Sound rocket capabilities are being improved with options financed through this topic.

 

Small launch vehicles provide direct access for a small spacecraft to the destination or orbit of interest at a time of the small spacecraft mission’s choosing. In support of exploration, science, and technology demonstration missions, further expansion of these vehicle’s reach beyond LEO is needed. To expand the risk-tolerant small spacecraft approach to deep space missions, frequent and low-cost access to destinations of interest beyond Earth is required.

 

In the longer term, technical capabilities of small spacecraft at Venus, Mars, or NEO destinations will be demonstrated by SSTP, and ultimately new kinds of transfer vehicles derived from these capabilities may be needed to propel them there.

References:

• https://www.nasa.gov/sites/default/files/atoms/files/sstp_soa2018_final.pdf
• https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170012214.pdf
• https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20160010585.pdf
• https://www.nasa.gov/directorates/spacetech/small_spacecraft/index.html
• https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20180007439.pdf
• https://www.rocketlabusa.com/photon/
• https://www.rocket.com/space/spacepower-propulsion/bipropellant-space-propulsion
• https://www.rocket.com/sites/default/files/documents/CubeSat%20Mod%20Prop2sided.pdf

Lead Center: ARC

Participating Centers: GRC, JPL, LaRC, MSFC

Scope Title:

Modular and Flexible Wireless Avionics Architectures and Wireless Sensing and Integrated Avionics for Space Applications

Scope Description:

 

The subtopic solicits developments in wireless avionics and wireless sensing for small spacecraft and may include technologies that:

 

1. Improve the reliability and applicability of wireless avionics for small spacecraft with significant improvements in subsystem size, mass, and volume, particularly if the technology can simplify the spacecraft fabrication, test, and integration process.
2. Allow innovative architectures for wireless avionics featuring plug-and-play software supporting modular subsystems that can be easily incorporated into specific small-satellite missions.
3. Improve fault detection aboard spacecraft using wireless sensor systems to augment current wired sensors and which include the capability of adding sensors to address developmental and flight instrumentation use.
4. Use innovative techniques for embedding sensors and other avionics components into a spacecraft to reduce or eliminate large and heavy cables and connectors, or that enable data transfer inside and across rotating mechanisms and pressure interfaces or into remote locations where it is difficult or unfeasible to run cables or where cables are at risk of failure.
5. Use additive manufacturing of wireless components such as antennas, sensors, and processing elements to create new components that may be smaller and lighter than current products. These new components could possibly be embedded into materials and structures that enable in situ structural health management, contributing to the development of smart structures and materials.
6. Include sensors and actuators that can be distributed among cooperative spacecraft to enable automated inspection of space assets or resource detection at the surface of the Moon, Mars, or other celestial bodies.

Key performance parameters (KPPs) would include improvements of at least a factor of 2 over existing technology in size, mass, and power consumption for sensors and associated components for a wireless instrumentation system. Improvements of sensor network throughput greater than 5× the current 2-Mbps performance is desired, along with reduction of latency and incorporation of timing information.

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

• Hardware
• Prototype
• Software
• Research
• Analysis

Desired Deliverables Description: State of the Art and Critical Gaps:

Development of small satellites missions benefit from a growing number of users worldwide. This means there may be a large pool of COTS components available for a specific mission (depending on the type and class of mission). A variety of command and data handling (C&DH) developments for CubeSats have resulted from in-house development, from new companies that specialize in CubeSat avionics, and from established companies who provide spacecraft avionics for the space industry in general. Presently there are a number of commercial vendors who offer highly integrated systems that contain the onboard computer, memory, electrical power system (EPS), and the ability to support a variety of input and output (I/O) for the CubeSat class of small spacecraft.

 

Wireless networks have been incorporated as crew support aboard the International Space Station (ISS). Wireless sensor networks have been flown as demonstrations. Dynamic self-configuring wireless networks have been evaluated in the lab. AIAA has defined the Space Plug-and-Play (SPA) standard, and flight demonstrations are planned.

The maturation of additive manufacturing and 3D-printing technology are making embedded wireless sensors and avionics a possibility. Embedding transmission lines, antennas, connectors, and sensors onto a spacecraft structure turns that structure into a multifunctional system that reduces or eliminates bulky cables and connectors. Embedded passive wireless sensors can greatly increase sensing and telemetry capabilities, including providing low-cost techniques for vehicle health management in future missions. Moreover, flexible embedded passive sensors created with conductive and functional fabrics are enabling new opportunities for sensing in surfaces and systems where sensing has been traditionally absent, such as parachutes and inflatable structures.

Relevance / Science Traceability:

NASA and other space agencies are exploring the application of SmallSats for deep space missions. The availability of modular wireless data connectivity alleviates complexity in testing and integration of systems. Modular components allow easier reconfiguration and late additions to any design. This is a benefit conferred on any spacecraft of any size, with the larger systems benefiting from savings in mass due to a larger reduction in cable harnesses and connectors.

References:

State of the Art: Small Spacecraft Technology: https://www.nasa.gov/sites/default/files/atoms/files/soa2018_final_doc.pdf

Fly-by-Wireless Conference: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20070013704.pdf

Fly-by-Wireless Conference 2007: https://ntrs.nasa.gov/search.jsp?R=20070013704

Fly-by-Wireless Update (2012): https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120010669.pdf

Wireless Aircraft Interconnect (WAIC) Systems: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170000686.pdf

Backscatter Systems for WAIC: https://ntrs.nasa.gov/search.jsp?R=20180004760

NASA Armstrong Patent: https://technology.nasa.gov/patent/DRC-TOPS-42

NASA Trade Study:

https://pdfs.semanticscholar.org/b7d6/e6d92ec78b6bee4cffd5a7f613b90b4508b8.pdf?_ga=2.244696965.1804159109.1563897519-1127952606.1563032260

Passive Wireless Sensor Technology Workshops: https://attend.ieee.org/wisee-2019/program/workshops/

 

 

Lead Center: ARC

Participating Centers: GSFC, JPL

Scope Title:

Artificial Intelligence (AI)/Machine Learning (ML) for Small Spacecraft Swarm Trajectory Control

Scope Description:

Constellations of small spacecraft currently provide unprecedented persistent coverage of the Earth’s surface, but the use of distributed missions for exploration infrastructure and multipoint scientific measurements beyond Earth will require new approaches to operational efficiency. Current commercial constellations use ground-based semiautonomous scheduling and orbital maintenance to decrease the need for spacecraft-by-spacecraft human-in-the-loop decision making. However, each spacecraft is still individually commanded by the ground-based system. For missions operating beyond Earth, the spacecraft will need to be operated as a single unit. Enabling command and control capabilities within the flight element of the distributed mission will allow control of an otherwise impractical number of small spacecraft as well as decreased operational costs for missions with fewer spacecraft.

 

NASA intends to expand the exploitation of small spacecraft swarms (or potentially constellations) in support of exploration and science missions. Multiple numbers of SmallSats, working in concert, offer unique capabilities and benefits to space researchers and satellite operators. For instance, some of these advantages embedded in fractionalized architectures, such as fault-tolerance and continuous repair and upgrades enabling dynamic, agile, adaptable mission plans, are better able to deal with the unplanned and unexpected. However, as the number of spacecraft grows, and destinations of interest move farther away from traditional low Earth orbit (LEO) orbits, operational challenges created by managing large numbers of agents along with significant space-to-ground communication latencies and bandwidth issues require alternative architectures that are able to make time-critical decisions and operate within the maneuvering limitations of each swarm member. Innovative technologies such as AI/ML can contribute significantly to the success of these deep space, multi-spacecraft missions by providing local control less reliant on the ground, as well as optimal use of propulsion.

 

Small spacecraft operating in formation, in close proximity to other objects, or beyond the capacity of human-in-the-loop control will be required to process input onboard and execute correct responses autonomously. These sensor-driven operations will be enabling for safe proximity operations with spacecraft or small bodies as well as the detection and reaction to transient events for observation.

 

This subtopic is interested in software agents and architectures that enable relative stationkeeping, multi-spacecraft orbit determination and prediction, autonomous reactive operations, and interspacecraft timing and communications to enable the above. Proposals should address software applications and/or network applications that enable:

• Efficient information exchange between individual spacecraft.
• Minimal reliance on ground commanding.
• Efficient use of space-qualified computing architectures.
• High-precision swarm navigation and control.
• Asymmetric use of ground assets (emphasizing space side over ground side).

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 10 Autonomous Systems
    Level 2: TX 10.2 Reasoning and Acting
Desired Deliverables of Phase I and Phase II:

• Software
• Prototype

Desired Deliverables Description:

Phase I Deliverables:

1. Software architecture design, including figures of merit (FOMs) for performance.
2. Test-bed environment for software development and testing (identify requirements or develop/describe test bed).
3. Plan to continue through Phase II.
4. Phase I report.

Phase II Deliverables:

1. Software suitable for testing on test-bed environment.
2. Software description/documentation.
3. Final report, including test results.

State of the Art and Critical Gaps:

Currently, the operation of each small spacecraft is individually planned, scheduled, and commanded. Signals from each individual small spacecraft are acquired by ground stations individually. Volumes of unprocessed raw data demand high bandwidth, and this becomes even less practical for greater numbers of cooperative small spacecraft operating at farther distances. As SmallSats are deployed to the Moon, Mars, near-Earth objects (NEOs), and other distant locations, it is costly—in terms of antenna time and human labor—and impractical due to time-of-flight delays at long distance and the fast dynamics of formation flight and cooperative operations to command and control each spacecraft individually, to react to dynamic situations at very remote destinations, and to change the focus of observations, data acquisition, and sample manipulation as areas of interest are discovered at the target. Very little has been done with true swarms of SmallSats, especially in deep space or even lunar locales. Flight processors are becoming more capable, but advanced software, autonomous processes, and adaptive systems that take advantage of increased processing power for SmallSats lag behind. Intelligent, adaptive behaviors and autonomous decision making in reaction to changing conditions at the target, with minimal dependence on Earthbound operators, is needed.

Relevance / Science Traceability:

The Small Spacecraft Technology Program is very interested in developing and demonstrating this technology for a broad list of crosscutting applications within the Science Mission Directorate (SMD) and Human Exploration and Operations Mission Directorate (HEOMD). Examples include:

• Fault-tolerant swarms composed of individual assets capable of reassigning their function to compensate for loss of an individual spacecraft.
• Spacecraft capable of decisions and tactical adaptation for acquiring, preprocessing, and transmitting scientific information (rather than raw data) at distant science or exploration destinations.
• Cooperative groups of spacecraft capable of disaggregated inspection, repair, resource resupply, and other functions autonomously and cooperatively at cislunar or more remote destinations.

Small spacecraft conducting science will need to make observations, process dynamic conditions, make decisions, and adapt their performance without manual intervention (for example, sampling the eruption of a plume at Enceladus).

Suggested use cases:

• Investigation of near-Earth asteroids (NEAs) – Orbit determination and maneuvering around an asteroid.
• Coordinated sensor operations (remote sensing, etc.) between multiple spacecraft at planetary destinations.
• Coordinated observation of distant objects (stars, planets) using multiple sensor platforms.
• Autonomous formation flying and inspection of other space assets.

References:

Achieving Science with CubeSats: Thinking Inside the Box. Committee on Achieving Science Goals with CubeSats; Space Studies Board; Division on Engineering and Physical Sciences; National Academies of Sciences, Engineering, and Medicine, 2016. Committee Chair: Thomas Zurbuchen. (https://doi.org/10.17226/23503)

Mission Design for Deep Space CubeSats. NASA Technical Reports, ARC-E-DAA-TN43486, October 30, 2017. (https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170010358.pdf)

Operating Multiple Satellites as a Single Entity: Introducing SODA. (https://digitalcommons.usu.edu/smallsat/2017/all2017/100/)

Lead Center: ARC

Participating Centers: MSFC

Scope Title:

Modular Open Systems Architectures for Small-Spacecraft Platforms

Scope Description:

This suptopic requests advances within modular open systems architectures for small spacecraft. As the most accessible spacecraft platform logistically and financially, small spacecraft benefit from a heritage based on rapid deployment and cost-effective missions. To further the state of the art (SOA) of both of these considerations, further cost savings may be found by standardizing the system architectures that drive the subsystems for these platforms. Such a realization would enable modular, hot-swappable spacecraft subsystems to accommodate the ever-increasing need for a wider definition of what small spacecraft are capable of and utilized for.

The development of standardized, hot-swappable interfaces should be compliant with and cognizant of NASA spacecraft standards. Of particular interest are designs acquiescent to the Agency standards existing between grounding, thermal, software, and data transfer interfaces. 

The adaptability introduced by an open and modular, interchangeable commercial-off-the-shelf (COTS) architecture furthers the ability to tailor current spacecraft designs for novel applications without requiring significant modifications to existing platforms. Also of interest are advances in modules that minimize complexity in spacecraft manufacturing (such as deterring geometrical modifications by virtue of manufacturing). Advances in additive manufacturing may enable critical enhancements to the performance of small-spacecraft systems by embedding otherwise impractical internal features (such as through holes and cavities for electronics integration). 

Systems that are applicable to small spacecraft (CubeSats up to Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) class), but scalable to large vehicles can result in a significant reduction of risk for more complex and longer duration missions. Near-term missions include:

• Cislunar, lunar orbiting, lunar landed, and exploration precursor.
• Low Earth orbit (LEO) “swarms” for Earth science and heliophysics.
• Disaggregated cooperative ensembles and sustained infrastructure for human exploration.

New applications might include manned spacecraft inspection, repair, communications support, and related areas. Proposals that provide reliable performance in extreme environments and that show a path to a flight demonstration are preferable.

The subtopic solicits developments in open modular architectures for small spacecraft and may include technologies that:

1. Provide interchangeable hardware and software with standardized interfaces.
2. Enable spacecraft to be built up from “plug and play” components.
3. Improve the state of the art of open interfacing platforms suitable for small spacecraft, leveraging COTS wherever possible.
4. Leverage novel manufacturing-in-the-loop considerations for small-spacecraft design standardization.
5. Increase the reliability and durability of small-spacecraft hardware and software by integrating subsystem considerations directly into the design process at the architectural level. 
6. Demonstrate expanded adaptivity for small spacecraft, allowing for platforms to be rapidly varied with respect to altering objectives and variable risk postures.

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

Promising platform architectures that enable the standardization of COTS hardware and software could be demonstrated through benchtop setups validating numerous protocols and compliance with existing NASA design standards for small spacecraft. A demonstration of ease of hot-swapping would be ideal, demonstrating how rapidly such a system could be adapted for altered requirements with new instrumentation and subsystems.

The deliverables should address improvements for ease of integration of varied hardware and software, plug-and-play integration of small-spacecraft subsystems, increased assembly speed of small spacecraft, utilization of advanced manufacturing for ease of integration, automated error assessment for targeted repairability of subsystems, reduced small-spacecraft design complexity, and reduction of small-spacecraft development cost through standardized COTS.

Phase I Deliverable:

Trade study for and demonstration of how NASA small-spacecraft standards, such as thermal, grounding, and software/data normalizations, could be implemented into hot-swappable, modular architecture.

These architectures must be cognizant of:

• NASA thermal interface standards to demonstrate necessary conductivity and respective thermal isolation. 
• NASA grounding interface standards to mitigate unwanted currents through single- or multiple-point grounding framework.
• NASA software and data interfacing standards, complying with Unified S-Band (USB) or Consultative Committee for Space Data Systems (CCSDS) standards.

Phase II Deliverable:

A benchtop hardware demonstration of open and modular architectures at work, exhibiting the standards within Phase I being conserved. The components should take advantage of supply-chain-compliant, heritage-relevant COTS whenever possible.

State of the Art and Critical Gaps:

The current SOA leverages COTS and compiled standards for integrating small spacecraft into a functional system meeting varied mission requirements. A number of in-house developments within NASA have complemented progress in academia and private industry to develop the infrastructure required to expand and normalize the definition of small-spacecraft-compliant subsystems and instrumentation. An issue arises with the software and hardware architecture regulating the agreement of these subsystems with NASA standards. Commercial vendors offering plug-and-play components are often only compliant with a limited number of subsystems, and consequently there exists a need to address this with an open modular architecture to enable more rapid, compliant, and consequently cost-effective small spacecraft that meet NASA's standards. Notable standardization gaps exist within communication gaps (such as wireless systems) and interconnectivity protocols, including but not limited to sustainable (and commonly grounded) power and data transfer with respect to manufacturability considerations.

Relevance / Science Traceability:

NASA and other space agencies are exploring the application of SmallSats for deep space missions. Modular architectures would enable a hot-swap adaptivity to altering mission requirements and serve as low-cost, rapid solutions for emerging destinations as they arise. Modular components allow easier reconfiguration and late additions to any design. Small-spacecraft modularity can be analogous for larger systems as well by virtue of defining and standardizing interconnectivity of universal COTS systems, enabling new objectives to be realized with a wide variety of instrumentation with a wide scope of requirements.

References:

• NASA Common Instrument Interface Project. Hosted Payload Guide for Proposers. Document Number: HPIG0001, Version: Initial. 22 March 2018.
• NASA Common Instrument Interface Project. Hosted Payload Guidelines Document. Document Number: CII-CI-0001, Version: Rev. A. 4 November 2013. Earth System Science Pathfinder Program Office.
• Tracking and Data Relay Satellite (TDRS) Third Generation Capabilities. NASA. 2013. Available from: http://www.nasa.gov/directorates/heo/scan/services/networks/txt_tdrs_gen3.html

Scope Title:

Batch-Producible Small Spacecraft

Scope Description:

The Batch-Producible Small Spacecraft subtopic requests proposals to address the need for industry collaboration to manufacture 30 to 100 small spacecraft for a wide variety of missions, addressing objectives ranging from heliophysics to constellation demonstrations and sensor web applications. The ability to fabricate relatively large "batches" of spacecraft will play an important role with regard to the throughput required for addressing the needs of the mission objectives listed above. As an advent in tandem with small-spacecraft swarms, batch-producible spacecraft are an increasing need as larger spacecraft are replaced with many smaller spacecraft, distributing sensing and collaboratively accomplishing objectives enabled novelly by variable topologies and network-based considerations. 

Advances in batch producibility are in tandem with standardization of rapid manufacturing of small spacecraft by private industry and will likely take advantage of advances in throughput-favorable fabrication methods. The manufacturability of batch-producible small spacecraft would need to consider the required throughput of manufacturing as a factor intrinsic to the small-spacecraft design itself. These systems must still remain compliant with existing NASA small-spacecraft protocols for thermal, electrical, communications, and redundancy considerations. However, batch-producible spacecraft should leverage design methodologies that would decrease the cost and increase the compatibility of these standardized requisites by virtue of the manufacturing process itself, exhibiting design-for-standardization through the engineering process.   

Such a batch-producible set of small spacecraft should leverage supply chain considerations wherever possible and should integrate commercial-off-the-shelf (COTS) components and instrumentation into the design of spacecraft architecture. The end result of rapidly manufacturable batches of spacecraft should demonstrate a significant reduction in manufacturing costs for 30 to 100 buses, with quicker turnaround times than otherwise possible over a range of NASA-relevant projects.

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

An overview and technical description of methods for batch producibility of small spacecraft within the range of 30 to 100 buses, demonstrating the integration of COTS as part of the framework. Successful demonstrations of this deliverable should demonstrate an increase in the competency of the following objectives:

• A standardized high-throughput manufacturing method to enable the fabrication of small spacecraft in batches of 30 to 100 buses (within the scope of CubeSats, up to and including ESPA-class spacecraft).
• A systematic decision tree that addresses fabrication turnaround-time considerations as a factor of spacecraft complexity.
• Demonstrated cost decreases for spacecraft batches with respect to the current state of the art (SOA).
• The integration and normalization of COTS relevant for batch production of small spacecraft as a function of supply chain availability and vendor capabilities.

Phase II Deliverable 

Integrating small-spacecraft standards into batch production and demonstrating an infrastructure that is batch-compliant. Successful demonstrations of this deliverable should demonstrate an increase in the competency of the following objectives:

• The integration of common NASA small-spacecraft standards (such as thermal, grounding, communications) directly into batch producibility.     
• A method for rapid assembly of batch-produced small spacecraft that accounts for manufacturability directly into the architecture of common subsystems (such as power, communications, etc.). 

State of the Art and Critical Gaps:

The current SOA of batch-produced small spacecraft relies heavily on the industry-demonstrated heritage of COTS for small-satellites. These systems have limited throughput considerations and are currently inappropriate for meeting future mission requisites pertaining to small spacecraft requiring the fabrication and integration of 30 to 100 spacecraft at a time (such as those relevant to heliophysics missions, network demonstrations, and swarm considerations).

Relevance / Science Traceability:

Partnership with industry on batch production of spacecraft will be required for distributed missions including synthetic apertures, disaggregated science observations, rapidly established planetary communications architectures, constellations, and sensor web applications; planned heliophysics missions call for 30 to 100 spacecraft. Technology development missions would also benefit from low-cost and shorter lead-time standardized bus platforms.

References:

• State of the Art Small Spacecraft Technology, NASA. 2018. https://www.nasa.gov/sites/default/files/atoms/files/soa2018_final_doc-6.pdf
• https://www.nasa.gov/sites/default/files/atoms/files/nac_march2017_blal_ida_sstp_tagged.pdf
• http://mstl.atl.calpoly.edu/~workshop/archive/2018/Spring/Day%201/Session%201/JimCockrell.pdf
• Manchester, Zachary, and Mason Peck. "Stochastic space exploration with microscale spacecraft." AIAA Guidance, Navigation, and Control Conference. 2011.
• Hopkinson, Neil, and Phill Dickens. "Rapid prototyping for direct manufacture." Rapid Prototyping Journal (2001).
• Kandela, Rami. Developing an Advanced Hardware Testing System to Enable Rapid Spacecraft Manufacturing. Diss. 2019.
• Le Moigne, Jacqueline, John Carl Adams, and Sreeja Nag. "A new taxonomy for distributed spacecraft missions." IEEE Journal of Selected Topics in Applied Earth Observations and Remote Sensing 13 (2020): 872-883.

Power is a ubiquitous technology need across many NASA missions.  Within the SBIR Program, power is represented across a broad range of topics in human exploration, space science, space technology and aeronautics.  New technologies are needed to generate electrical power and/or store energy for future human and robotic space missions and to enable hybrid electric aircraft that could revolutionize air travel.  A key goal is to develop technologies that are multi-use and cross-cutting for a broad range of NASA mission applications. In aeronautics, power technologies are needed to supply large-scale electric power and efficiently distribute the power to aircraft propulsors (see Focus Area 18 – Air Vehicle Technologies).  In the space power domain, mission applications include planetary surface power, large-scale spacecraft prime power, small-scale robotic probe power, and smallsat/cubesat power. Applicable technology options include photovoltaic arrays, radioisotope power systems, nuclear fission, thermal energy conversion, motor/generators, fuel cells, batteries or other energy storage devices, power management, transmission, distribution and intelligent control. An overarching objective is to mature technologies from analytical or experimental proof-of-concept (TRL3) to breadboard demonstration in a relevant environment (TRL5). Successful efforts will transition into NASA Projects where the SBIR/STTR deliverables will be incorporated into ground testbeds or flight demonstrations.

Lead Center: GRC

Participating Centers: JPL

Scope Title:

Photovoltaic Energy Conversion

Scope Description:

This subtopic is seeking photovoltaic cell and blanket technologies that lead to significant improvements in overall solar array performance for missions in areas of scientific interest including high-intensity, high-temperature (HIHT); low-intensity, low-temperature (LILT); and high-radiation environments. Additionally sought are solar power systems that can provide high power in compactly stowed volumes for small spacecraft. 
 

These improvements may be achieved by optimizing the cell technology to operate in HIHT/LILT, increasing end of life (EOL) performance, increasing photovoltaic cell efficiency above 35% at 1 AU, and decreasing solar cell module/blanket stowed volume. Missions at distances of greater than 1 AU may include an inner planetary flyby, as such technologies that optimize solar cell string length to account for the changes in power generation are also of interest. 
 

Photovoltaic energy conversion: advances in, but not limited to, the following: (1) Photovoltaic cell and blanket technologies capable of LILT operation applicable to outer planetary (low solar intensity) missions; (2) Photovoltaic cell and blanket technologies capable of HIHT operation applicable to inner planetary missions; (3) Photovoltaic cell and blanket technologies that enhance and extend performance in lunar applications including orbital, surface, and transfer; and  (4) Solar cell and blanket technologies to support missions in high-radiation, LILT environments near Jupiter and its moons. 
 

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.1 Power Generation and Energy Conservation
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Phase I deliverables include detailed reports with proof of concept and key metrics of components tested and verified.

Phase II deliverables include detailed reports with relevant test data along with proof-of-concept hardware and components developed.

State of the Art and Critical Gaps:

State-of-the-art (SOA) photovoltaic array technology consists of high-efficiency, multijunction cell technology on thick honeycomb panels and, as of late, lightweight blanket system deployable systems. There are very limited demonstrated technology for HIHT and LILT missions. A current solution for high-radiation intensity involves adding thick cover glass to the cells, which increases the overall system mass. 
 

Significant improvements in overall performance are needed to address the current gaps between SOA and many mission requirements for photovoltaic cell efficiency >30%, array mass specific power >200 W/kg, decreased stowed volume, long-term operation in radiation environments, high-power arrays, and a wide range of environmental operating conditions.

Relevance / Science Traceability:

These technologies are relevant to any space science, Earth science, planetary surface, or other science mission that requires affordable high-efficiency photovoltaic power production for orbiters, flyby craft, landers, and rovers.
 

Specific requirements can be found in the References, but include many future Science Mission Directorate (SMD) missions. Specific requirements for orbiters and flybys to Outer planets include: LILT capability (>38% at 10 AU and <140 °C), radiation tolerance (6×10¹⁵ 1 MeV e/cm²), high power (>50 kW at 1 AU), low mass (3× lower than the standard operating procedure (SOP)), low volume (3× lower than SOP), long life (>15 years), and high reliability.

These technologies are relevant and align with any Space Technology Mission Directorate (STMD) or Human Exploration and Operations Mission Directorate (HEOMD) mission that requires affordable high-efficiency photovoltaic power production.
 

NASA outlines New Lunar Science, Human Exploration Missions: https://www.nasa.gov/feature/nasa-outlines-new-lunar-science-human-exploration-missions

NASA Science Missions: https://science.nasa.gov/missions-page?field_division_tid=All&field_phase_tid=3951

 

References:

• Solar Power Technologies for Future Planetary Science Missions: https://solarsystem.nasa.gov/resources/548/solar-power-technologies-for-future-planetary-science-missions/
• NASA outlines New Lunar Science, Human Exploration Missions: https://www.nasa.gov/feature/nasa-outlines-new-lunar-science-human-exploration-missions
• NASA Science Missions: https://science.nasa.gov/missions-page

Lead Center: GRC

Participating Centers: None

Scope Title:

Dynamic Power Conversion

Scope Description:

NASA is developing dynamic radioisotope power systems (DRPSs) for unmanned robotic missions to the Moon and other solar system bodies of interest. This technology directly aligns with the Science Mission Directorate (SMD) strategic technology investment plan for space power and energy storage and could be infused into a highly efficient RPS for missions to dark, dusty, or distant destinations where solar power is not practical. Current work in DRPSs is focused on novel Stirling, Brayton, or Rankine convertors that would be integrated with one or more 250-W_th general-purpose heat source (GPHS) modules or 1-W_th lightweight radioisotope heater unit (RHU) to provide high thermal-to-electric efficiency, low mass, long life, and high reliability for planetary spacecraft, landers, and rovers. Heat is transferred from the radioisotope heat source assembly to the power convertor hot end using conductive or radiative coupling. Power convertor hot-end temperatures would generally range from 300 to 500 °C for RHU applications and 500 to 800 °C for GPHS applications. Waste heat is removed from the cold end of the power convertor at temperatures ranging from 20 to 175 °C, depending on the application, using conductive coupling to radiator panels. The NASA projects target power systems able to produce a range of electrical power output levels based on the available form factors of space-rated fuel sources. These include a very low range of 0.5 to 2.0 W_e that would utilize one or more RHU, a moderate range of 40 to 70 W_e that would utilize a single GPHS Step-2 module, and a high range of 100 to 500 W_e that would utilize multiple GPHS Step-2 modules. For these power ranges, one or more power convertors could be used to improve overall system reliability. The current solicitation is focused on innovations that enable efficient and robust power conversion systems. Areas of interest include:

1. Robust, efficient, highly reliable, and long-life thermal-to-electric dynamic power convertors that would be used to populate a generator of a prescribed electric power output ranges.
2. Electronic controllers applicable to Stirling, Brayton, or Rankine power convertors.   
3. Multilayered metal insulation (MLMI) for minimizing environmental heat losses and maximizing heat transfer from the radioisotope heat source assembly to the power convertor.
4. Advanced dynamic power conversion components and RPS integration components, including efficient alternators able to survive extended exposure to 200 °C, robust high-temperature-tolerant Stirling regenerators, robust highly effective recuperators, integrated heat pipes, and radiators that improve system performance, and improve the margin, reliability, and fault tolerance for existing components.

Expected TRL or TRL Range at completion of the Project: 1 to 5
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.1 Power Generation and Energy Conservation
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Phase I deliverables: results of a feasibility study and analysis, as described in a final report.

 

Phase II deliverables: prototype hardware that has demonstrated basic functionality in a laboratory environment, the appropriate research and analysis used to develop the hardware, and maturation options for flight designs.

State of the Art and Critical Gaps:

Radioisotope power systems are critical for long-duration NASA missions in dark, dusty, or harsh environments. Thermoelectric systems have been used on the very successful RPSs flown in the past, but are limited in efficiency. Dynamic thermal energy conversion provides significantly higher efficiency, and through proper engineering of the noncontact moving components, can eliminate wear mechanisms and provide long life. Although high-efficiency performance of dynamic power convertors has been proven, reliable and robust systems tolerant of off-nominal operation are needed. In addition to convertors appropriate for GPHS RPSs, advances in much smaller and lower power dynamic power conversion systems are sought that can utilize RHUs for applications such as distributed sensor systems, small spacecraft, and other systems that take advantage of lower power electronics for the exploration of surface phenomenon on icy moons and other bodies of interest.  Although the power convertor advances are essential, to develop reliable and robust systems for future flight advances in convertor components as well as RPS integration components are also needed. These would include efficient alternators able to survive 200 °C, robust high-temperature-tolerant regenerators, robust high-efficiency recuperators, heat pipes, radiators, and controllers applicable to Stirling flexure-bearing, Stirling gas-bearing, or Brayton convertors. 

Relevance / Science Traceability:

This technology directly aligns with the Science Mission Directorate - Planetary Science Division for space power and energy storage. Investments in more mature technologies through the Radioisotope Power System Program is ongoing. This SBIR subtopic scope provides a lower TRL technology pipeline for advances in this important power capability that improves performance, reliability, and robustness.

 

References:

• Radioisotope Power Systems (RPS): https://rps.nasa.gov/about-rps/overview/
• Oriti, Salvatore: "Dynamic Power Convertor Development for Radioisotope Power Systems at NASA Glenn Research Center," AIAA Propulsion and Energy (P&E) 2018, AIAA 2018-4498.
• Wilson, Scott D.: "NASA Low Power Stirling Convertor for Small Landers, Probes, and Rovers Operating in Darkness," AIAA P&E 2018, AIAA 2018-4499.
• Wong, Wayne: "Advanced Stirling Convertor (ASC) Technology Maturation," AIAA P&E 2015, AIAA 2015-3806.

Scope Title:

Additive Manufacturing Microfabrication of Stirling Heat Engine Regenerators

Scope Description:

In space applications where solar power is not practical, dynamic power conversion is an effective alternative.  Of the several technologies used for dynamic power conversion, free-piston Stirling heat engines, coupled with alternators, offer high thermal-to-electric conversion efficiency, low mass, and long life.  One component of Stirling heat engines that contributes to their excellent efficiency is the regenerator, which acts as a heat exchanger/storage for the working fluid as it passes from the heat acceptor to the rejector and again as it returns to the acceptor to repeat the cycle.  The current state of the art in the construction of regenerators results in a cylindrical annulus made up of heat- and corrosion-resistant, short metallic fibers in diameters of 20 to 40 µm, packed to form an annulus with a porosity of 80% to 90% (solid fraction 10% to 20%), and sintered to achieve structural stability. 
 

In some instances, these random fiber regenerators have released small particle debris due to less-than-complete sintering of the fiber matrix, presenting a risk of interference in the very small running clearance gaps of the displacer and power pistons, and potentially negatively affecting the performance and robustness of the heat engine.  NASA has engaged in initial studies to determine the feasibility of producing continuous regenerator matrices through additive manufacturing (AM), and while these studies show promise, it has been determined that limitations of selective laser melting in the minimum achievable feature size and spacing between features prevents realization of performance goals.  Sought are advances in AM microfabrication that demonstrate:

1. Applicability to high-temperature, corrosion-resistant metal alloys (Inconel, FeCrAlY, etc.).
2. Capability of creating ligaments in diameters as small as 20 µm, with spacing between ligaments as small as 100 µm.
3. Capability of producing cylindrical annuli on the order of 5.5 cm O.D. and 4.0 cm I.D. in axial lengths of up to 5 cm.  Axial length may be achieved by stacking multiple components of shorter lengths.
4. Reasonable build time to support on-demand production.
5. Ability to create regenerator matrices that are stable and robust in the anticipated vibro-acoustic environments associated with space missions (launch, pyroshock, entry/descent/landing, etc.).

Expected TRL or TRL Range at completion of the Project: 1 to 4
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.1 Power Generation and Energy Conservation
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Phase I deliverables: results of a feasibility study and analysis, as described in a final report.

 

Phase II deliverables: prototype hardware that has demonstrated basic functionality in a laboratory environment, the appropriate research and analysis used to develop the hardware, and maturation options .

State of the Art and Critical Gaps:

Radioisotope power Systems (RPS) are critical for long duration NASA missions to destinations sufficiently far from the Sun that solar power is impractical, and for missions to permanently shadowed areas of planetary bodies and their moons.  Thermoelectric power systems (RTG) have enjoyed much success in past missions, but their efficiency is limited.  Dynamic RPS offer significantly higher efficiency, resulting in lower system mass and reduced radioisotope inventory for a given power output.  Advances in microfabrication of regenerators are needed for reduction of risk associated with stability of the regenerator matrix and improvements in reliability and robustness to support long mission durations.

Relevance / Science Traceability:

The technology described here aligns with the Science Mission Directorate Planetary Science Division (SMD/PSD) requirements for space power and energy storage and provides a low-TRL pathway for technologies that may contribute to a reduction of risk and improvements in reliability and robustness of Stirling heat engines in space-power applications.

References:

• Ibrahim, et al: “A Microfabricated Segmented-Involute-Foil Regenerator for Enhancing Reliability and Performance of Stirling Engines,” NASA Contractor Report 2007-215006, 2007.
• Ibrahim, M.B. and Tew, R.C.: Stirling Convertor Regenerators, CRC Press, Boca Raton, FL, 2017.

 

Lead Center: GRC

Participating Centers: JPL

Scope Title:

Energy Storage for Extreme Environments

Scope Description:

NASA's Planetary Science Division is working to implement a balanced portfolio within the available budget and based on a decadal survey that will continue to make exciting scientific discoveries about our solar system. This balanced suite of missions shows the need for low mass/volume energy storage that can effectively operate in extreme environments for future NASA Science Missions.

Future science missions will require advanced primary and secondary battery systems capable of operating at temperature extremes from -200 °C for outer planet missions to 400 to 500 °C for Venus missions, and a span of -230 to +120 °C for missions to the lunar surface. Operational durations of 30 to 60 days for Venus; 30 to 60 days for deep-space environments such as Europa, Enceladus, and Titan; and 14-day eclipses for lunar night survival and operations on the Moon are of interest. Advancements to battery energy storage capabilities that address operation for one of the listed missions (Venus, deep space, or lunar)  combined with high specific energy and energy density (>250 Wh/kg and >500 Wh/L for rechargeable or >800 Wh/kg and >1000 Wh/L for nonrechargeable at the cell level) are of interest in this solicitation. Novel battery-pack-level designs and technologies that enhance battery reliability and safety and support improved thermal management are also of interest.  Combinations of cell-level improvements and/or battery-system-level improvement for enhanced temperature capability will be considered.

Furthermore, missions that incorporate nonrechargeable (primary) batteries will benefit from instrumentation or modeling that can effectively determine state of charge to a high degree of accuracy and/or state of health, particularly those missions that use cell chemistries with discharge voltage profiles that are a weak function of state of charge or state of health such as lithium carbon monofluoride (Li-CFx) cells. Technologies of interest include: radiation-hardened (to 1 Mrad total ionizing dose) coulomb integration application-specific integrated circuits (ASIC) or hybrid circuits, with >1% accuracy over 1 to 20 A, operating over 24 to 36 V; computational models that can predict state of charge/state of health for primary cells; nondestructive instrumentation that can detect state of charge/state of health for primary and secondary cells.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.2 Energy Storage
Desired Deliverables of Phase I and Phase II:

• Prototype

Desired Deliverables Description:

Research should be conducted to demonstrate technical feasibility in a final report for Phase I and show a path toward a Phase II, and when possible, deliver a demonstration unit for NASA testing at the completion of the Phase II contract. Phase II emphasis should be placed on developing and demonstrating the technology under relevant test conditions. Additionally, a path should be outlined that shows how the technology could be commercialized or further developed into science-worthy systems.

State of the Art and Critical Gaps:

State-of-the-art primary and rechargeable cells are limited in both capacity and temperature range.  Typical primary Li-SO₂ and Li-SOCl₂ operate within a maximum temperature range of -40 to 80 °C but suffer from capacity loss, especially at low temperatures.  At -40 °C, the cells will provide roughly half the capacity available at room temperature. Similarly, rechargeable Li-ion cells operate within a narrow temperature range of -20 to 40 °C and also suffer from capacity loss at lower temperatures. The lower limit of temperature range of rechargeable cells can be extended through the use of low-temperature electrolytes, but with limited rate capability and concerns over lithium plating on charge. There is currently a gap that exists for high-temperature batteries, primary and rechargeable, that can operate at Venus atmospheric temperatures. In addition, there is a gap in the ability to accurately predict or measure the amount of usable capacity of primary battery cells, particularly after a long mission cruise with exposure to varying temperatures and ionizing radiation dose.  This solicitation is aimed at the development of cells that can maintain performance at extreme temperatures to minimize or eliminate the need for strict thermal management of the batteries (which adds complexity and mass to the spacecraft) as well as instrumentation or modeling to predict state of charge/state of health of primary batteries for deep-space missions.

Relevance / Science Traceability:

These batteries are applicable over a broad range of science missions.  Low-temperature batteries are needed for potential NASA decadal missions to ocean worlds (Europa, Enceladus, and Titan) and the icy giants (Neptune, Uranus). These batteries are also needed for science missions on the lunar surface. Low-temperature batteries developed under this subtopic would enhance these missions and could be potentially enabling if the missions are mass or volume limited.  There is also significant interest in a Venus surface mission that will require primary and/or rechargeable batteries that can operate for 60+ days on the surface of Venus. A high-temperature battery that can meet these requirements is enabling for this class of missions.

References:

NASA Science: https://science.nasa.gov/

Solar Electric Propulsion: https://www1.grc.nasa.gov/space/sep/

Lead Center: GRC

Participating Centers: GSFC, JSC

Scope Title:

Innovative Ways to Transmit Power Over Long Distances for Lunar and Mars Missions

Scope Description:Expected TRL or TRL Range at completion of the Project: 3 to 6
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.3 Power Management and Distribution
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

Typically, deliverables under Phase I proposals are geared toward a technology concept with associated analysis and design. A final report usually suffices in summarizing the work, but if a prototype is delivered, that is viewed as a bonus. Phase II hardware prototypes will have opportunities for infusion into NASA technology testbeds and commercial landers.

State of the Art and Critical Gaps:

While high-power terrestrial distribution systems exist, there is no equivalent to a lunar or planetary base. Unique challenges must be overcome in order to enable a realistic power architecture for these future applications, especially when dealing with the environmental extremes that will be encountered. The temperature swings will be a critical requirement on any technology developed, from power converters to cabling or power-beaming concepts. In addition, proposals will have to consider lunar regolith and Mars dust storms. To enable a new Mars transportation capability for human exploration, new technology development must be started soon to address the very unique needs of a mixed AC/DC space-rated power system to prove feasibility and provide realistic performance metrics for detailed vehicle design concepts and mission trade studies. 

Relevance / Science Traceability:

This subtopic would directly address a remaining technology gap in the lunar and Mars surface mission concepts and Mars human transportation needs. There are potential infusion opportunities with SMD (Science Mission Directorate) Commercial Lander Payload Services (CLPS) and HEOMD (Human Exploration and Operations Mission Directorate) Flexible Lunar Exploration (FLEx) Landers. In addition, technologies developed could benefit other NASA missions, including Gateway. The power levels may be different, but the technology concepts could be similar, especially when dealing with temperature extremes.

References:

The Global Exploration Roadmap, January 2018: https://www.nasa.gov/sites/default/files/atoms/files/ger_2018_small_mobile.pdf

Space Policy Directive-1, Reinvigorating America's Human Space Exploration Program, December 11, 2017: https://www.federalregister.gov/documents/2017/12/14/2017-27160/reinvigorating-americas-human-space-exploration-program

Lead Center: GSFC

Participating Centers: GRC, JPL, LaRC

Scope Title:

Radiation-Tolerant High-Voltage, High-Power Electronics

Scope Description:

Other innovative heavy-ion SEE radiation-tolerant, high-power, high-voltage discrete device technologies will be considered that offer significant electrical performance improvement over state-of-the-art heavy-ion SEE radiation-tolerant power devices.

Expected TRL or TRL Range at completion of the Project: 4 to 5
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.3 Power Management and Distribution
Desired Deliverables of Phase I and Phase II:

• Hardware
• Prototype
• Analysis

Desired Deliverables Description:

• For vertical-field power devices: No heavy-ion-induced permanent destructive effects upon irradiation while in blocking configuration (in powered reverse-bias/off state) with ions having a silicon-equivalent surface-incident linear energy transfer (LET) of 40 MeV-cm²/mg and sufficient energy to maintain a rising LET level throughout the epitaxial layer(s).
• For all other devices: No heavy-ion-induced permanent destructive effects upon irradiation while in blocking configuration (in powered reverse-bias/off state) with ions having a silicon-equivalent surface-incident LET of 75 MeV-cm²/mg and sufficient energy to fully penetrate the active volume prior to the ions reaching their maximum LET value (Bragg peak).

Deliverables in Phase II shall include prototype and/or production-ready semiconductor devices (diodes and/or transistors); and device electrical and radiation performance characterization (device electrical performance specifications, heavy-ion SEE radiation test results, and total-dose radiation analyses).

State of the Art and Critical Gaps:Relevance / Science Traceability:

Power transistors and diodes form the building blocks of numerous power circuits for spacecraft and science instrument applications. This subtopic therefore feeds a broad array of space technology hardware development activities by providing SEE (heavy-ion) radiation-hardened state-of-the-art device technologies that achieve higher voltages with lower power consumption and greater efficiency than presently available.

 

Taxonomy Area (TX) 03.3.4, Power Management and Distribution (PMAD) - Advanced Electronic Parts, calls out the need for development of radiation-hardened high-voltage components for power systems. This subtopic serves as a feeder to the subtopic Lunar and Planetary Surface Power Distribution, in which WBG circuits for PMAD applications are solicited. The solicited developments in this subtopic will also feed systems development for the NASA Kilopower project due to the savings in size/mass combined with radiation hardness.

 

TX 08.1.2, Sensors and Instruments - Electronics: Radiation-hardened high-voltage transistors are needed for low-mass, low-leakage, high-efficiency applications such as LIDAR Q-switch drivers, mass spectrometers, and electrostatic analyzers. These applications are aligned with science objectives including Earth science LIDAR needs, Jovian moon exploration, and Saturn missions. Finally, mass spectrometers critical to planetary and asteroid research and in the search for life on other planets such as Mars require high-voltage power systems and will thus benefit from mass and power savings from this subtopic's innovations.

References:

Partial listing of relevant references:

1. E. Mizuta, et al., "Single-Event Damage Observed in GaN-on-Si HEMTs for Power Control Applications," IEEE Transactions on Nuclear Science, vol. 65, pp. 1956-1963, 2018.
2. M. Zerarka, et al., "TCAD Simulation of the Single Event Effects in Normally-OFF GaN Transistors After Heavy Ion Radiation," IEEE Transactions on Nuclear Science, vol. 64, pp. 2242-2249, 2017.
3. C. Abbate, et al., "Experimental Study of Single Event Effects Induced by Heavy Ion Irradiation in Enhancement Mode GaN Power HEMT," Microelectronics Reliability, vol. 55, pp. 1496-1500, 2015.
4. J. Kim, et al., "Radiation damage effects in Ga2O3 materials and devices," Journal of Materials Chemistry C, vol. 7, pp. 10-24, 2019.
5. S. J. Pearton, et al., "Perspective: Ga2O3 for ultra-high power rectifiers and MOSFETS," Journal of Applied Physics, vol. 124, p. 220901, 2018.
6. D. R. Ball, et al., “Ion-Induced Energy Pulse Mechanism for Single-Event Burnout in High-Voltage SiC Power MOSFETs and Junction Barrier Schottky Diodes,” IEEE Transactions on Nuclear Science, vol. 67, no. 1, pp. 22-28, 2020.
7. J. McPherson, et al., "Mechanisms of Heavy Ion Induced Single Event Burnout in 4H-SiC Power MOSFETs," Materials Science Forum, 1004, 889–896, 2020.
8. S. Kuboyama, et al., "Thermal Runaway in SiC Schottky Barrier Diodes Caused by Heavy Ions," IEEE Transactions on Nuclear Science, vol. 66, pp. 1688-1693, 2019.
9. C. Abbate, et al., "Gate Damages Induced in SiC Power MOSFETs During Heavy-Ion Irradiation--Part I," IEEE Transactions on Electron Devices, vol. 66, no. 10, pp. 4235-4242, Oct. 2019. [See also Part II ]
10. J.-M. Lauenstein, “Getting SiC Power Devices Off the Ground: Design, Testing, and Overcoming Radiation Threats,” Microelectronics Reliability and Qualification Working (MRQW) Meeting, El Segundo, CA, February 2018. https://ntrs.nasa.gov/search.jsp?R=20180006113

Lead Center: GRC

Participating Centers: MSFC

Scope Title:

Megawatt-Class Nuclear Power System

Scope Description:

Recent Mars transportation assessments identify megawatt-class nuclear electric propulsion (NEP) systems as a reasonable approach for use in a crewed mission to Mars. A critical subcomponent of the reference NEP concept is a dynamic thermal-to-electric power convertor. Dynamic power convertors are needed that address the following technical challenges:

• Robust, efficient, high-reliability, long-life thermal-to-electric power conversion and controller technology in the minimum range of 100 to 500 kWe. Brayton, Rankine, and Stirling convertors are of primary interest. Multiple parallel/redundant convertors may be used to achieve the megawatt-class power level.
  • Includes subcomponents such as efficient turbomachinery, bearings, alternators, recuperators, and heat exchangers.
• Convertors must be capable of interfacing with pumped liquid-metal loops: one for thermal energy input and one for heat rejection.

In addition, liquid metal pumps that address the following technical challenges are also needed:

• Can operate for long periods of time (years) at relevant in-space (in vacuum and zero g) reactor/power generation temperatures.
• Can withstand liquid metal freeze-thaw transition during initial reactor startup in zero g.
• Low mass and high efficiency at fluid throughputs that are relevant for in-space nuclear power generation.
• Wetted surfaces of the pump composed of materials that are compatible with liquid metals under consideration for in-space nuclear power generation (NaK, Li, etc.).

The desired deliverables are primarily prototype hardware, research, and analysis to demonstrate concept feasibility and a Technology Readiness Level (TRL) range of 3 to 5. There is a strong desire for hardware that can operate or has a clear path for operation in the relevant space environment. The specified higher power levels are of priority, but demonstrations at lower power levels may be considered as long as the scaling to higher power levels is straightforward and does not require significant new technology or configuration change. The prototype hardware may include one (or more) of the following:

• Power convertor (hot-end temperature = 850 ºC, cold-end temperature = 100 to 200 ºC).
• Controller electronics.
• Convertor subcomponent(s).
• Liquid metal pump and/or subcomponent(s).

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.1 Power Generation and Energy Conservation
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

The desired deliverables for Phase I include monthly progress reports and a comprehensive final report.

For Phase II, we are primarily interested in component and/or breadboard hardware that demonstrates concept feasibility in a laboratory or relevant environment. The appropriate research and analysis required to develop the hardware are also desired. The Phase II deliverables should include hardware, monthly progress reports, and a comprehensive final report.

State of the Art and Critical Gaps:

Multikilowatt electric propulsion systems are well developed and have been used on commercial and military satellites for several years. Higher power electric propulsion systems are currently being considered to support crewed missions to near-Earth asteroids and as cargo transport for sustained lunar or Mars exploration, and for very high power crewed missions to Mars and the outer planets. One of the key technologies required in a NEP system is the power convertor. A recent Mars Transportation Assessment Study was completed that included Brayton, Stirling, Rankine, thermoelectric, and thermionic technologies in the trade space. The study identified HeXe Brayton as the baseline dynamic power convertor in the reference NEP concept, with supercritical CO₂ Brayton and K-Rankine as the primary options. Current state-of-the-art Brayton technology has been demonstrated in a relevant space environment (ground test) in the 10s of kWe range.  Convertor scaleup to the 100s of kWe per unit is required.

Relevance / Science Traceability:

This technology directly aligns with the Space Technology Mission Directorate (STMD) roadmap for space power and energy storage.

References:

Mason, Lee S., “Dynamic Energy Conversion: Vital Technology for Space Nuclear Power,” Journal of Aerospace Engineering, April 2013.

https://www.researchgate.net/publication/275186601_Dynamic_Energy_Conversion_Vital_Technology_for_Space_Nuclear_Power

Gilland, James H., LaPointe, Michael R., et al., “MW-Class Electric Propulsion System Designs for Mars Cargo Transport,” AIAA 2011-7253, September 2011. 

https://arc.aiaa.org/doi/abs/10.2514/6.2011-7253

 

This focus area includes development of robotic systems technologies (hardware and software) to improve the exploration of space. Robots can perform tasks to assist and off-load work from astronauts. Robots may perform this work before, in support of, or after humans. Ground controllers and astronauts will remotely operate robots using a range of control modes, over multiple spatial ranges (shared-space, line of sight, in orbit, and interplanetary) and with a range of time-delay and communications bandwidth. Technology is needed for robotic systems to improve transport of crew, instruments, and payloads on planetary surfaces, on and around small bodies, and in-space. This includes hazard detection, sensing/perception, active suspension, grappling/anchoring, legged locomotion, robot navigation, end-effectors, propulsion, and user interfaces.

In the coming decades, robotic systems will continue to change the way space is explored. Robots will be used in all mission phases: as independent explorers operating in environments too distant or hostile for humans, as precursor systems operating before crewed missions, as crew helpers working alongside and supporting humans, and as caretakers of assets left behind. As humans continue to work and live in space, they will increasingly rely on intelligent and versatile robots to perform mundane activities, freeing human and ground control teams to tend to more challenging tasks that call for human cognition and judgment.

Innovative robot technologies provides a critical capability for space exploration. Multiple forms of mobility, manipulation and human-robot interaction offer great promise in exploring planetary bodies for science investigations and to support human missions. Enhancements and potentially new forms of robotic systems can be realized through advances in component technologies, such as actuation and structures (e.g. 3D printing). Mobility provides a critical capability for space exploration. Multiple forms of mobility offer great promise in exploring planetary bodies for science investigations and to support human missions. Manipulation provides a critical capability for positioning crew members and instruments in space and on planetary bodies, it allows for the handling of tools, interfaces, and materials not specifically designed for robots, and it provides a capability for drilling, extracting, handling and processing samples of multiple forms and scales. This increases the range of beneficial tasks robots can perform and allows for improved efficiency of operations across mission scenarios. Manipulation is important for human missions, human precursor missions, and unmanned science missions.  Sampling, sample handling, transport, and distribution to instruments, or instrument placement directly on in-place rock or regolith, is important for robotic missions to locales too distant or dangerous for human exploration.

Future space missions may rely on co-located and distributed teams of humans and robots that have complementary capabilities. Tasks that are considered "dull, dirty, or dangerous" can be transferred to robots, thus relieving human crew members to perform more complex tasks or those requiring real-time modifications due to contingencies. Additionally, due to the limited number of astronauts anticipated to crew planetary exploration missions, as well as their constrained schedules, ground control will need to remotely supervise and assist robots using time-delayed and limited bandwidth communications.  Advanced methods of human-robot interaction over time delay will enable more productive robotic exploration of the more distant reaches of the solar system.  This includes improved visualization of alternative future states of the robot and the terrain, as well as intuitive means of communicating the intent of the human to the robotic system.

Lead Center: JPL

Participating Centers: ARC, GRC, GSFC

Scope Title:

Robotic Mobility, Manipulation, and Sampling

Scope Description:Expected TRL or TRL Range at completion of the Project: 2 to 4
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
• Hardware
• Software

Desired Deliverables Description:

Hardware, software, and designs for component robotic systems.

• Phase I: proof of concept to include research and analysis along with design in a final report.
• Phase II: prototype for further testing.

State of the Art and Critical Gaps:

Scoops, powder drills, and rock core drills and their corresponding handling systems have been developed for sample acquisition on Mars and asteroids. Non-flight systems have been developed for sampling on comets, Venus, and Earth's Moon.  Some of these environments still present risk and have gaps that need to be addressed (i.e. Venus). 

 

Ocean worlds exploration presents new environments and unique challenges not met by existing mobility and sampling systems.   New mobility, manipulation, and sampling technologies are needed to enable new types of missions and missions to different and challenging environments. 

Relevance / Science Traceability:

The subtopic supports multiple programs within Science Mission Directorate (SMD). The Mars program has had infusion of technologies such as a force-torque sensor in the Mars 2020 mission. Recent awards would support the Ocean Worlds program with surface and deep drills for Europa, and future awards could include technologies to support missions to Enceladus, Titan, and other planetary bodies with subsurface oceans.  Sample-return missions could be supported such as from Ceres, comets, and asteroids. Products from this subtopic have been proposed for New Frontiers program missions. With renewed interest in return to Earth's Moon, the mobility and sampling technologies could support future robotic missions to the Moon.

References:

• Mars Exploration/Program & Missions, https://mars.nasa.gov/programmissions/
• Solar System Exploration, https://solarsystem.nasa.gov/
• Ocean Worlds website: https://www.nasa.gov/specials/ocean-worlds/
• Ocean Worlds article: https://science.nasa.gov/news-articles/ocean-worlds

 

Lead Center: ARC

Participating Centers: JSC

Scope Title:

Improve the Capability or Performance of Intravehicular Activity Robots

Scope Description:

To support human exploration beyond Earth orbit, NASA is developing Gateway, which will be an orbiting facility near the Moon. This facility will serve as a starting point for missions to cislunar space and beyond. It could enable assembly and servicing of telescopes and deep space exploration vehicles. It could also be used as a platform for astrophysics, Earth observation, heliophysics, and lunar science.

In contrast to the International Space Station (ISS), which is continuously manned, Gateway is expected to be occupied by humans only intermittently—perhaps only 1 month per year. Consequently, there is a significant need for Gateway to have autonomous capabilities for performing payload operations and spacecraft caretaking, particularly when astronauts are not present. Similar capabilities are needed for future lunar or planetary surface habitats. Intravehicular activity (IVA) robots can potentially perform a wide variety of tasks, including systems inspection, monitoring, diagnostics and repair, logistics and consumables stowage, exploration capability testing, aggregation of robotically returned destination surface samples, and science measurements and operations.

The objective of this subtopic is to develop technologies that can improve the capability or performance of IVA robots to perform payload operations and spacecraft caretaking. Proposals are specifically sought to create technologies that can be integrated and tested with the NASA Astrobee, Robonaut 2, or other NASA robots in the following areas:

• Sensors and perception systems for performing contact tasks; manipulation; and/or interior environment monitoring, inspection, modeling, and navigation.
• Robotic tools for manipulating logistics and stowage or performing maintenance, housekeeping, or emergency management operations (e.g., fire detection and suppression in multiple constrained locations or cleaning lunar dust out of air filters).
• Operational subsystems that enable extended robot operations (power systems, efficient propulsion, etc.); increase robot autonomy via computationally efficient methods (planning, scheduling, and task execution); or improve human-robot interaction between IVA robots and human teams on the ground under communications constraints, including low bandwidth and extended loss-of-signal periods (software architecture, remote operations methods, etc.).

This subtopic also seeks to advance technologies that will enable the next generation of IVA robots to operate in lunar surface habitats, including:

• Novel robotic end effectors capable of reliably performing fine grasping tasks, such as plugging and unplugging MIL-STD-38999 electrical connectors and fluid quick-disconnect connectors.
• Compact, reliable, modular robotic actuators and controllers for IVA robots.
• Software that enables autonomous management of robot operational and hardware faults such that the robot can “fail operational.” For example, the software may use algorithms to determine how to automatically respond to a failure in a motion planner for move to a commanded location by taking into account a projected collision and replanning to the next closest point not in collision.

Expected TRL or TRL Range at completion of the Project: 4 to 5
Primary Technology Taxonomy:
    Level 1: TX 04 Robotics Systems
    Level 2: TX 04.X Other Robotic Systems
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Proposals must describe how the technology will make a significant improvement over the current state of the art, rather than just an incremental enhancement, for a specific IVA robot application.

Deliverables should focus on prototype components, subsystems, and the demonstration thereof. Specifically, Phase I awards shall deliver an interim and final report discussing these results. Phase II awards shall deliver demonstration reports along with supporting software, design information, and documentation.

State of the Art and Critical Gaps:

The technology developed by this subtopic would both enable and enhance the Astrobee free-flying robot and Robonaut 2 humanoid robot, which are the state of the art for IVA robots. SBIR technology would improve the capability and performance of these robots to routinely and robustly perform IVA tasks, particularly internal spacecraft payload operations and logistics. New technology created by 2021 SBIR awards could potentially be tested with these, or other, robots in ground testbeds at Ames Research Center (ARC) and Johnson Space Center (JSC) in follow-on awards. Likewise, on-orbit testing on ISS may be possible during follow-on awards.

 

The technology developed by this subtopic would also fill technical gaps identified by the proposed Game Changing Development (GCD) Integrated System for Autonomous and Adaptive Caretaking (ISAAC) project, which will mature autonomy technology to support the caretaking of human exploration spacecraft. In particular, the SBIR technology would help provide autonomy and robotic capabilities that are required for in-flight maintenance (both preventive and corrective) of Gateway during extended periods when crew are not present.

Relevance / Science Traceability:

This subtopic is directly relevant to the following STMD (Space Technology Mission Directorate) investments:

• Astrobee freeflying robot, GCD
• Integrated System for Autonomous and Adaptive Caretaking (ISAAC), GCD
• Smart Deep Space Habitats (SmartHabs), Space Technology Research Institutes (STRI)

This subtopic is directly relevant to the following HEOMD (Human Exploration and Operations Mission Directorate) investments:

• SPHERES (Synchronized Position Hold, Engage, Reorient, Experimental Satellite)/Astrobee facility, ISS
• Robonaut 2 humanoid robot, ISS
• Gateway program, Advanced Exploration Systems (AES)
• Logistics Reduction project, AES
• Autonomous Systems Operations project, AES

References:

The NASA Science Missions Directorate seeks technology for cost-effective high-performance advanced space telescopes for astrophysics and Earth science. Astrophysics applications require large aperture light-weight highly reflecting mirrors, deployable large structures and innovative metrology, control of unwanted radiation for high-contrast optics, precision formation flying for synthetic aperture telescopes, and cryogenic optics to enable far infrared telescopes. A few of the new astrophysics telescopes and their subsystems will require operation at cryogenic temperatures as cold a 4 K. This topic will consider technologies necessary to enable future telescopes and observatories collecting electromagnetic bands, ranging from UV to millimeter waves, and also include gravity waves. The subtopics will consider all technologies associated with the collection and combination of observable signals. Earth science requires modest apertures in the 2 to 4 meter size category that are cost effective. New technologies in innovative mirror materials, such as silicon, silicon carbide and nanolaminates, innovative structures, including nanotechnology, and wavefront sensing and control are needed to build telescopes for Earth science.

Lead Center: JPL

Participating Centers: GSFC

Scope Title:

Control of scattered starlight with coronagraphs and starshades

Scope Description:

The goal of this subtopic is to address the unique problem of imaging and spectroscopic characterization of faint astrophysical objects that are located within the obscuring glare of much brighter stellar sources. Examples include planetary systems beyond our own, the detailed inner structure of galaxies with very bright nuclei, binary star formation, and stellar evolution. Contrast ratios of one million to ten billion over an angular spatial scale of 0.05 to 1.5 arcsec are typical of these objects. Achieving a very low background requires control of both scattered and diffracted light. The failure to control either amplitude or phase fluctuations in the optical train severely reduces the effectiveness of starlight cancellation schemes.

This innovative research focuses on advances in coronagraphic instruments, starlight cancellation instruments, and potential occulting technologies that operate at visible and near-infrared wavelengths. The ultimate application of these instruments is to operate in space as part of a future observatory mission concepts such as the Habitable Exoplanet Observatory (HabEx) and Large Ultraviolet Optical Infrared Surveyor (LUVOIR). Measurement techniques include imaging, photometry, spectroscopy, and polarimetry. There is interest in component development and innovative instrument design, as well as in the fabrication of subsystem devices to include, but not limited to, the following areas:

Starlight Suppression Technologies:

• Hybrid metal/dielectric and polarization apodization masks for diffraction control of phase and amplitude for coronagraph-scaled starshade experiments.
• Low-scatter, low-reflectivity, sharp, flexible edges for control of solar scatter in starshades.
• Low-reflectivity coatings for flexible starshade optical shields.
• Methods to distinguish the coherent and incoherent scatter in a broadband speckle field.

Wavefront Measurement and Control Technologies:

• Small-stroke, high-precision, deformable mirrors and associated driving electronics scalable to 10,000 or more actuators (both to further the state of the art towards flightlike hardware and to explore novel concepts). Multiple deformable mirror technologies in various phases of development and processes are encouraged to ultimately improve the state of the art in deformable mirror technology. Process improvements are needed to improve repeatability, yield, and performance precision of current devices.
• Multiplexers with ultralow power dissipation for electrical connection to deformable mirrors.
• Low-order wavefront sensors for measuring wavefront instabilities to enable real-time control and postprocessing of aberrations.
• Thermally and mechanically insensitive optical benches and systems.

Optical Coating and Measurement Technologies:

• Instruments capable of measuring polarization crosstalk and birefringence to parts per million.
• Polarization-insensitive coatings for large optics.
• Methods to measure the spectral reflectivity and polarization uniformity across large optics.

In addition this subtopic solicits proposals to develop components that improve the footprint, robustness, power consumption, reliability, and wavefront quality of high-contrast, low-temporal bandwidth, adaptive optics systems. These include application-specific integrated circuit (ASIC) drivers that easily integrate with the deformable mirrors, improved connectivity technologies, as well as high-actuator-count deformable mirrors  with high-quality, ultrastable wavefronts.

It also seeks coronagraph masks that can be tested in ground-based high-contrast testbeds in place at a number of institutions, as well as devices to measure the masks to inform optical models. The masks include transmissive scalar, polarization-dependent, and spatial apodizing masks, including those with extremely low reflectivity regions that allow them to be used in reflection.

The subtopic seeks samples of optical coatings that reduce polarization and can be applied to large optics as well as methods and instruments to characterize them over large optical surfaces.

Finally, for starshades, the subtopic seeks low-reflectivity and potentially diffraction-controlling edges that minimize scattered sunlight while also remaining robust to handling and cleaning.  Low-reflectivity optical coatings that can be applied to the surfaces for the large (hundreds of square meters) optical shield are also desired.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Under this Subtopic a concept study provided as a final report in Phase I is acceptable and a prototype for Phase II is acceptable.

 

State of the Art and Critical Gaps:

Coronagraphs have been demonstrated to achieve high contrast in moderate bandwidth in laboratory environments. Starshades will enable even deeper contrast over broader bands but to date have demonstrated deep contrast in narrow band light. The extent to which the telescope optics will limit coronagraph performance is a function of the quality of the optical coating and the ability to control polarization over the full wavefront. Neither of these technologies is well characterized at levels required for 10¹⁰ contrast. Wavefront control using deformable mirrors is critical. Controllability and stability to picometer levels is required. To date, deformable mirrors have been up to the task of providing contrast approaching 10¹⁰, but they require thousands of wires, and overall wavefront quality and stroke remain concerns.

Relevance / Science Traceability:

These technologies are directly applicable to the Nancy Grace Roman Space Telescope (NGRST) coronagraph instrument (CGI), and the HabEx and LUVOIR concept studies.

References:

See SPIE conference papers and articles published in the Journal of Astronomical Telescopes and Instrumentation on high-contrast coronagraphy, segmented coronagraph design and analysis, and starshades.

Websites:

• Exoplanet Exploration - Planets Beyond Our Solar System: https://exoplanets.jpl.nasa.gov
• Exoplanet Exploration Program: https://exoplanets.nasa.gov/exep/
• Goddard Space Flight Center: https://www.nasa.gov/goddard

Lead Center: JPL

Participating Centers: GSFC

Scope Title:

Assembled Deployable Optical Metering Structures and Instruments

Scope Description:

Future space astronomy missions from ultraviolet to millimeter wavelengths will push the state of the art in current optomechanical technologies. Size, dimensional stability, temperature, risk, manufacturability, and cost are important factors, separately and in combination. The Large Ultraviolet Optical Infrared Surveyor (LUVOIR) calls for deployed apertures as large as 15 m in diameter; the Origins Space Telescope (OST), for operational temperatures as low as 4 K; and LUVOIR and the Habitable Exoplanet Observatory (HabEx), for exquisite optical quality. Methods to construct large telescopes in space are also under development.  Additionally, sunshields for thermal control and starshades for exoplanet imaging require deployment schemes to achieve 30- to 70-m-class space structures.

This subtopic addresses the need to mature technologies that can be used to fabricate 10- to 20-m-class, lightweight, ambient or cryogenic flight-qualified observatory systems and subsystems (telescopes, sunshields, starshades). Proposals to fabricate demonstration components and subsystems with direct scalability to flight systems through validated models will be given preference. The target launch volume and expected disturbances, along with the estimate of system performance, should be included in the discussion. Novel metrology solutions to establish and maintain optical alignment will also be accepted.

Technologies including, but not limited to, the following areas are of particular interest:

Precision structures/materials:

• Low coefficient of thermal expansion/coefficient of moisture expansion (CTE/CME) materials/structures to enable highly dimensionally stable optics, optical benches, metering structures.
• Materials/structures to enable deep-cryogenic (down to 4 K) operation.
• Novel athermalization methods to join materials/structures with differing mechanical/thermal properties.
• Lightweight materials/structures to enable high-mass-efficiency structures.
• Precision joints/latches to enable submicron level repeatability.
• Mechanical connections providing microdynamic stability suitable for robotic assembly.

Deployable technologies:

• Precision deployable modules for assembly of optical telescopes (e.g., innovative active or passive deployable primary or secondary support structures).
• Hybrid deployable/assembled architectures, packaging, and deployment designs for large sunshields and external occulters (20 to 50 m class).
• Packaging techniques to enable more efficient deployable structures.

Metrology:

• Techniques to verify dimensional stability requirements at subnanometer level precisions (10 to 100 pm).
• Techniques to monitor and maintain telescope optical alignment for on-ground and in-orbit operation.

A successful proposal shows a path toward a Phase II delivery of demonstration hardware scalable to 5-m diameter for ground test characterization. Proposals should show an understanding of one or more relevant science needs, and present a feasible plan to fully develop the relevant subsystem technologies and to transition into future NASA program(s).

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

For Phase I, a successful deliverable would include a demonstration of the functionality and/or performance of a system/subsystem with model predictions to explain observed behavior as well as make predictions on future designs.

For Phase II this should be demonstrated on units that can be scaled to future flight sizes.

State of the Art and Critical Gaps:

The James Webb Space Telescope, currently set to launch in 2021, represents the state of the art in large deployable telescopes. The Wide Field Infrared Survey Telescope’s (WFIRST) coronagraph instrument (CGI) will drive telescope/instrument stability requirements to new levels. The mission concepts in the upcoming Astro2020 decadal survey will push technological requirements even further in the areas of deployment, size, stability, lightweighting, and operational temperature. Each of these mission studies have identified technology gaps related to their respective mission requirements.

Relevance / Science Traceability:

These technologies are directly applicable to the WFIRST CGI and the HabEx, LUVOIR, and OST mission concepts.

References:

• Large UV/Optical/IR Surveyor (LUVOIR): https://asd.gsfc.nasa.gov/luvoir/
• Habitable Exoplanet Observatory (HabEx): https://www.jpl.nasa.gov/habex/
• Origins Space Telescope: https://asd.gsfc.nasa.gov/firs/
• What is an Exoplanet? https://exoplanets.nasa.gov/what-is-an-exoplanet/technology/
• NASA in-Space Assembled Telescope (iSAT) Study: https://exoplanets.nasa.gov/exep/technology/in-space-assembly/iSAT_study/

Lead Center: MSFC

Participating Centers: GRC, GSFC, JPL, LaRC

Scope Title:

Optical Components and Systems for Large Telescope Missions

Scope Description:

4. For extreme ultraviolet (EUV):

• Surface slope: <0.1 µrad

 

Also needed is the ability to fully characterize surface errors and predict optical performance.

 

Proposals must show an understanding of one or more relevant science needs and present a feasible plan to develop the proposed technology for infusion into a NASA program: suborbital rocket or balloon; competed SMEX or MIDEX; or Decadal-class mission. Successful proposals will demonstrate an ability to manufacture, test, and control ultra-low-cost optical systems that can meet science performance requirements and mission requirements (including processing and infrastructure issues). Material behavior, process control, active and/or passive optical performance, and mounting/deploying issues should be resolved and demonstrated.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Prototype
• Hardware

Desired Deliverables Description:

• An ideal Phase I deliverable would be a precision optical system of at least 0.25 m; a relevant subcomponent of a system; a prototype demonstration of a fabrication, test, or control technology leading to a successful Phase II delivery; or a reviewed preliminary design and manufacturing plan that demonstrates feasibility. While detailed analysis will be conducted in Phase 2, the preliminary design should address how optical, mechanical (static and dynamic), and thermal designs and performance analyses will be done to show compliance with all requirements. Past experience or technology demonstrations that support the design and manufacturing plans will be given appropriate weight in the evaluation.
• An ideal Phase II project would further advance the technology to produce a flight-qualifiable optical system greater than 0.5 meters or relevant subcomponent (with a TRL in the 4 to 5 range) or a working fabrication, test, or control system. Phase I and Phase II mirror system or component deliverables would be accompanied by all necessary documentation, including the optical performance assessment and all data on processing and properties of its substrate materials. A successful mission-oriented Phase II would have a credible plan to deliver for the allocated budget a fully assembled and tested telescope assembly that can be integrated into the potential mission as well as demonstrate an understanding of how the engineering specifications of their system meets the performance requirements and operational constraints of the mission (including mechanical and thermal stability analyses).

State of the Art and Critical Gaps:

Current normal incidence space mirrors cost $4 million to $6 million per square meter of optical surface area. This research effort seeks a cost reduction for precision optical components by 5× to 50×, to between $100K/m² and $1M/m².

Relevance / Science Traceability:

S2.03 primary supports potential Astrophysics Division missions. S2.03 has made optical systems in the past for potential balloon experiments. Future potential Decadal missions include Laser Interferometer Space Antenna (LISA), Habitable Exoplanet Observatory (HabEx), Large UV/Optical/Near-IR Surveyor (LUVOIR) and the Origins Space Telescope (OST).

References:

The HabEx and LUVOIR space telescope studies are developing concepts for UVOIR space telescopes for exo-Earth discovery and characterization, exoplanet science, general astrophysics and solar system astronomy.

• The HabEx Interim Report is available at: https://www.jpl.nasa.gov/habex/pdf/interim_report.pdf.
• The LUVOIR Interim Report is available at: https://asd.gsfc.nasa.gov/luvoir/.
• The OST is a single-aperture telescope concept for the Far-Infrared Surveyor mission described in the NASA Astrophysics Roadmap, "Enduring Quests, Daring Visions: NASA Astrophysics in the Next Three Decades": https://smd-prod.s3.amazonaws.com/science-pink/s3fs-public/atoms/files/secure-Astrophysics_Roadmap_2013_0.pdf 
• The OST mission is described on the website: https://origins.ipac.caltech.edu
• The Space Infrared Interferometric Telescope (SPIRIT) and its optical system requirements description: https://asd.gsfc.nasa.gov/cosmology/spirit/
• LISA mission description: https://lisa.nasa.gov/

Scope Title:

Balloon Planetary Telescope

Scope Description:

Astronomy from a stratospheric balloon platform offers numerous advantages. At typical balloon cruise altitudes (100,000 to 130,000 ft.), 99%+ of the atmosphere is below the balloon, and the attenuation due to the remaining atmosphere is small, especially in the near-ultraviolet (NUV) band and in the infrared (IR) bands near 2.7 and 4.25 µm. The lack of atmosphere nearly eliminates scintillation and allows the resolution potential of relatively large optics to be realized, and the small amount of atmosphere reduces scattered light and allows observations of brighter objects even during daylight hours.

 

Potential balloon science missions are either in the UV/optical (UVO) or in the infrared/far-infrared (IR/FIR). 

• UVO science missions require a 1-m-class telescope diffraction limited at 500 nm or a primary mirror system that can maintain <10 nm rms surface figure error for elevation angles ranging from 0° to 60° over a temperature range of 220 to 280 K.
• IR science missions require 1.5-m-class telescopes diffraction limited at 5 µm. 
• FIR missions require 2-m-class (or larger) telescopes diffraction limited at 50 µm. 

 

In all cases, the telescopes need to achieve:

• Mass: <300 kg
• Shock: 10G without damage
• Elevation: 0° to 60°
• Temperature: 220 to 280 K

 

For packaging reasons, the primary mirror assembly must have a radius of curvature 3 m (nominal) and a mass <150 kg.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

• Phase I will produce a preliminary design and report including initial design requirements such as wavefront error budget, mass allocation budget, structural stiffness requirements, etc. as well as trade studies performed and and analysis that compares the design to the expected performance over the specified operating range. Development challenges shall be identified during Phase I, including trade studies and challenges to be addressed during Phase II with subsystem proof-of-concept demonstration hardware.
• If Phase II can only produce a subscale component, then it should also produce a detailed final design, including final requirements (wavefront error budget, mass allocation, etc.) and a performance assessment over the specified operating range.

State of the Art and Critical Gaps:

Current SOA (state-of-the-art) mirrors made from Zerodur^(C) or ULE^(C), for example, require lightweighting to meet balloon mass limitations and cannot meet diffraction limited performance over the wide temperature range due to the coefficient of thermal expansion limitations.

Relevance / Science Traceability:

“Vision and Voyages for Planetary Science in the Decade 2013-2022”

• Page 22, last paragraph of NASA Telescope Facilities within the Summary Section:
  Balloon- and rocket-borne telescopes offer a cost-effective means of studying planetary bodies at wavelengths inaccessible from the ground. Because of their modest costs and development times, they also provide training opportunities for would-be developers of future spacecraft instruments. Although NASA’s Science Mission Directorate regularly flies balloon missions into the stratosphere, there are few funding opportunities to take advantage of this resource for planetary science because typical planetary grants are too small to support these missions. A funding line to promote further use of these suborbital observing platforms for planetary observations would complement and reduce the load on the already oversubscribed planetary astronomy program.
• Page 203, 5th paragraph of section titled Earth and Space-Based Telescopes:
  Significant planetary work can be done from balloon-based missions flying higher than 45,000 ft. This altitude provides access to electromagnetic radiation that would otherwise be absorbed by Earth’s atmosphere and permits high-spatial-resolution imaging unaffected by atmospheric turbulence. These facilities offer a combination of cost, flexibility, risk tolerance, and support for innovative solutions that is ideal for the pursuit of certain scientific opportunities, the development of new instrumentation, and infrastructure support. Given the rarity of giant-planet missions, these types of observing platforms (high-altitude telescopes on balloons and sounding rockets) can be used to fill an important data gap.

 

Potential advocates include planetary scientists at Goddard Space Flight Center (GSFC), Johns Hopkins Applied Physics Laboratory (APL), and Southwest Research Institute, etc.

References:

• For additional discussion of the advantages of observations from stratosphere platforms, refer to: 
  Dankanich et. al.: “Planetary Balloon-Based Science Platform Evaluation and Program Implementation - Final Report,” available from:  https://ntrs.nasa.gov/ (search for "NASA/TM-2016-218870").
• Additional information about scientific balloons can be found at: https://www.csbf.nasa.gov/docs.html

Scope Title:

Large Ultraviolet/Optical/near-IR (LUVOIR) Surveyor and Habitable Exoplanet (HabEx) Missions

Scope Description:

Potential ultraviolet/optical (UVO) missions require 4- to 16-m monolithic or segmented primary mirrors with <5 nm rms surface figures. Active or passive alignment and control is required to achieve system-level diffraction-limited performance at wavelengths less than 500 nm (<40-nm rms wavefront error, WFE). Additionally, a potential exoplanet mission, using an internal coronagraph, requires total telescope wavefront stability on the order of 10 pm rms per 10 min. This stability specification places severe constraints on the dynamic mechanical and thermal performance of 4-m and larger telescope. Potential enabling technologies include: active thermal control systems, ultrastable mirror support structures, athermal telescope structures, athermal mirror struts, ultrastable joints with low coefficient of thermal expansion (CTE) and high stability, and vibration compensation.

 

Mirror areal density depends upon available launch vehicle capacities to Sun-Earth L2 (i.e., 15 kg/m² for a 5-m-fairing Evolved Expendable Launch Vehicle (EELV) versus 150 kg/m² for a 10-m-fairing Space Launch (SLS)). Regarding areal cost, a good goal is to keep the total cost of the primary mirror at or below $100M. Thus, an 8-m-class mirror (with 50 m² of collecting area) should have an areal cost of less than $2M/m². And, a 16-m-class mirror (with 200 m² of collecting area) should have an areal cost of less than $0.5M/m².
 

Key technologies to enable such a mirror include new and improved:

• Mirror substrate materials and/or architectural designs.
• Processes to rapidly fabricate and test UVO quality mirrors.
• Mirror support structures, joints, and mechanisms that are athermal or have zero CTE at the desired scale.
• Mirror support structures, joints, and mechanisms that are ultrastable at the desired scale.
• Mirror support structures with low mass that can survive launch at the desired scale.
• Mechanisms and sensors to align segmented mirrors to <1 nm rms precisions.
• Thermal control (<1 mK) to reduce wavefront stability to <10 pm rms per 10 min.
• Dynamic isolation (>140 dB) to reduce wavefront stability to <10 pm rms per 10 min.

 

Also needed is the ability to fully characterize surface errors and predict optical performance via integrated optomechanical modeling.

 

Potential solutions for substrate material/architecture include, but are not limited to: ultra-uniform low-CTE glasses, silicon carbide, nanolaminates, or carbon-fiber-reinforced polymer. Potential solutions for mirror support structure material/architecture include, but are not limited to: additive manufacturing, nature-inspired architectures, nanoparticle composites, carbon fiber, graphite composite, ceramic or SiC materials, etc. Potential solutions for new fabrication processes include, but are not limited to: additive manufacture, direct precision machining, rapid optical fabrication, roller embossing at optical tolerances, slumping, or replication technologies to manufacture 1- to 2-m- (or larger) precision quality components. Potential solutions for achieving the 10-pm wavefront stability include, but are not limited to: metrology, passive, and active control for optical alignment and mirror phasing; active vibration isolation;  metrology; and passive and active thermal control.

Expected TRL or TRL Range at completion of the Project: 2 to 4
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Hardware
• Software

Desired Deliverables Description:

• An ideal Phase I deliverable would be a precision optical system of at least 0.25 m; a relevant subcomponent of a system; a prototype demonstration of a fabrication, test, or control technology leading to a successful Phase II delivery; or a reviewed preliminary design and manufacturing plan that demonstrates feasibility. While detailed analysis will be conducted in Phase II, the preliminary design should address how optical, mechanical (static and dynamic), and thermal designs and performance analyses will be done to show compliance with all requirements. Past experience or technology demonstrations that support the design and manufacturing plans will be given appropriate weight in the evaluation.
• An ideal Phase II project would further advance the technology to produce a flight-qualifiable optical system greater than 0.5 m or relevant subcomponent (with a TRL in the 4 to 5 range) or a working fabrication, test, or control system. Phase I and Phase II mirror system or component deliverables would be accompanied by all necessary documentation, including the optical performance assessment and all data on processing and properties of its substrate materials. A successful mission-oriented Phase II would have a credible plan to deliver for the allocated budget a fully assembled and tested telescope assembly that can be integrated into the potential mission as well as demonstrate an understanding of how the engineering specifications of their system meets the performance requirements and operational constraints of the mission (including mechanical and thermal stability analyses).

State of the Art and Critical Gaps:

The precision fabrication of large mirrors is a daunting task. The fabrication process needs to be scaled from the state-of-the-art (SOA) Hubble mirror at 2.4 m both in precision and dimensions of the mirrors.

Relevance / Science Traceability:

S2.03 primary supports potential Astrophysics Division missions. S2.03 has made optical systems in the past for potential balloon experiments. Future potential Decadal missions include Laser Interferometer Space Antenna (LISA), Habitable Exoplanet Observatory (HabEx), Large UV/Optical/Near-IR Surveyor (LUVOIR) and the Origins Space Telescope (OST).

References:
Scope Title:

Near-Infrared Lidar Beam Expander Telescope

Scope Description:

Potential airborne coherent lidar missions need compact 15-cm diameter 20× magnification beam expander telescopes. Potential space-based coherent lidar missions need at least 50-cm 65× magnification beam expander telescopes. Candidate coherent lidar systems (operating with a pulsed 2-µm laser) have a narrow, almost diffraction-limited field-of-view, close to 0.8 lambda/D half angle. Aberrations, especially spherical aberration, in the optical telescope can decrease the signal.
 

Additionally, the telescope beam expander should maintain the laser beam’s circular polarization. The incumbent telescope technology is a Dall-Kirkham beam expander. Technology advance is needed to make the beam expander more compact with less mass while retaining optical performance, and to demonstrate the larger diameter. Additionally, technology for nonmoving scanning of the beam expander output is needed.

Expected TRL or TRL Range at completion of the Project: 3 to 4
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware

Desired Deliverables Description:

• An ideal Phase I deliverable would be a precision optical system of at least 0.15 m or a relevant subcomponent of a system.  While detailed analysis will be conducted in Phase II, the preliminary design should address how optical, mechanical (static and dynamic), and thermal designs and performance analysis will be done to show compliance with all requirements. Past experience or technology demonstrations that support the design and manufacturing plans will be given appropriate weight in the evaluation.
• An ideal Phase II project would further advance the technology to produce a flight-qualifiable optical system greater than 0.5 m or relevant subcomponent (with a TRL in the 4 to 5 range). Phase I and Phase II system or component deliverables would be accompanied by all necessary documentation, including the optical performance assessment and all data on processing and properties of its substrate materials. A successful mission-oriented Phase II would have a credible plan to deliver for the allocated budget a fully assembled and tested telescope assembly that can be integrated into the potential mission as well as demonstrate an understanding of how the engineering specifications of their system meets the performance requirements and operational constraints of the mission (including mechanical and thermal stability analyses).

State of the Art and Critical Gaps:Relevance / Science Traceability:

Science Mission Directorate (SMD) desires both an airborne coherent-detection wind-profiling lidar systems and space-based wind measurement. The space mission has been recommended to SMD by both the 2007 and 2017 Earth Science Decadal Surveys. SMD has incorporated the wind lidar mission in its planning and has named it "3-D Winds". SMD recently held the Earth Venture Suborbital competition for 5 years of airborne science campaigns. The existing coherent wind lidar at Langley, Doppler Aerosol Wind (DAWN), was included in three proposals that are under review. Furthermore, SMD is baselining DAWN for a second Convective Processes Experiment (CPEX-) type airborne science campaign and for providing calibration/validation assistance to the European Space Agency (ESA) Aeolus space mission. DAWN flies on the DC-8, and it is highly desired to fit DAWN on other NASA and National Oceanic and Atmospheric Administration (NOAA) aircraft. DAWN needs to lower its mass for several of the aircraft, and a low-mass telescope retaining the required performance is needed. Additionally, an electronic remote control of telescope focus is needed to adapt to aircraft cruise altitude and weather conditions during science flights.

References:

• NRC Decadal Surveys at: http://sites.nationalacademies.org/DEPS/ESAS2017/index.htm
• "Workshop Report on Scientific Challenges and Opportunities in the NASA Weather Focus Area," https://smd-prod.s3.amazonaws.com/science-pink/s3fs-public/atoms/files/Weather_Focus_Area_Workshop_Report_2015_0.pdf
• A. K. DuVivier, J. J. Cassano, S. Greco and G. D. Emmitt: “A Case Study of Observed and Modeled Barrier Flow in the Denmark Strait in May 2015,” Monthly Weather Review, 145, 2385–2404 (2017), https://doi.org/10.1175/MWR-D-16-0386.1 
  • See also supplemental material at: http://dx.doi.org/10.1175/MWR-D-16-0386.s1
• M. J. Kavaya, J. Y. Beyon, G. J. Koch, M. Petros, P. J. Petzar, U. N. Singh, B. C. Trieu, and J. Yu: “The Doppler Aerosol Wind Lidar (DAWN) Airborne, Wind-Profiling, Coherent-Detection Lidar System: Overview and Preliminary Flight Results,” J. of Atmospheric and Oceanic Technology 34 (4), 826-842 (2014), https://doi.org/10.1175/JTECH-D-12-00274.1
• Scott A. Braun, Ramesh Kakar, Edward Zipser, Gerald Heymsfield, Cerese Albers, Shannon Brown, Stephen L. Durden, Stephen Guimond, Jeffery Halverson, Andrew Heymsfield, Syed Ismail, Bjorn Lambrigtsen, Timothy Miller, Simone Tanelli, Janel Thomas, and Jon Zawislak: “NASA’s Genesis and Rapid Intensification Processes (GRIP) Field Experiment,” Bull. Amer. Meteor. Soc. (BAMS) 94(3), 345-363 (2013), https://doi.org/10.1175/BAMS-D-11-00232.1

Scope Title:

Fabrication, Test, and Control of Advanced Optical Systems

Scope Description:

Future ultraviolet (UV)/optical/near-infrared (NIR) telescopes require mirror systems that are very precise and ultrastable.

 

Regarding precision, this subtopic encourages proposals to develop technology that makes a significant advance in the ability to fabricate and test an optical system.

One area of current emphasis is the ability to nondestructively characterize coefficient of thermal expansion (CTE) homogeneity in 4-m-class Zerodur and 2-m-class ULE mirror substrates to an uncertainty of 1 ppb/K and a spatial sampling of 100×100. This characterization capability is needed to select mirror substrates before they undergo the expense of turning them into a lightweight space mirror.

 

Regarding stability, to achieve high-contrast imaging for exoplanet science using a coronagraph instrument, systems must maintain wavefront stability to <10 pm rms over intervals of ~10 min during critical observations. The ~10-min time period of this stability is driven by current wavefront sensing and control techniques that rely on stellar photons from the target object to generate estimates of the system wavefront. This subtopic aims to develop new technologies and techniques for wavefront sensing, metrology, and verification and validation of optical system wavefront stability.

 

Current methods of wavefront sensing include image-based techniques such as phase retrieval, focal-plane contrast techniques such as electric field conjugation and speckle nulling, and low-order and out-of-band wavefront sensing that use nonscience light rejected by the coronagraph to estimate drifts in the system wavefront during observations. These techniques are limited by the low stellar photon rates of the dim objects being observed (~5 to 11 Vmag), leading to 10s of minutes between wavefront control updates.

 

New methods may include: new techniques of using out-of-band light to improve sensing speed and spatial frequency content, new control laws incorporating feedback and feedforward for more optimal control, new algorithms for estimating absolute and relative wavefront changes, and the use of artificial guide stars for improved sensing signal-to-noise ratio and speed.

 

Current methods of metrology include edge sensors (capacitive, inductive, or optical) for maintaining segment cophasing, and laser distance interferometers for absolute measurement of system rigid body alignment. Development of these techniques to improve sensitivity, speed, and component reliability is desired. Low-power, high-reliability electronics are also needed.

 

Finally, metrology techniques for system verification and validation at the picometer level during integration and test (I&T) are needed. High speed spatial and speckle interferometers are currently capable of measuring single-digit picometer displacements and deformations on small components in controlled environments. Extension of these techniques to large-scale optics and structures in typical I&T environments is needed.

Expected TRL or TRL Range at completion of the Project: 2 to 4
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Hardware
• Software

Desired Deliverables Description: State of the Art and Critical Gaps:

Wavefront (WF) sensing using star images, including dispersed-fringe and phase-retrieval methods, is at TRL 6, qualified for space by the James Webb Space Telescope (JWST). WF sensing and control for coronagraphs, including electric field conjugation and low-order WF sensing (LOWFS), is at TRL4 and is being developed and demonstrated by Wide Field Infrared Survey Telescope Coronagraph Instrument (WFIRST/CGI).

 

Laser-distance interferometers for point-to-point measurements with accuracies from nanometers to picometers have been demonstrated on the ground by the Space Interferometry Mission and other projects, and on orbit by the LISA Pathfinder and Grace Follow-On mission. Application to telescope alignment metrology has been demonstrated on testbeds, to TRL4 for nanometer accuracy. Picometer accuracy for telescopes awaits demonstration.

 

Edge sensors are in use on segmented ground telescopes but are not yet on space telescopes. New designs are needed to provide picometer sensitivity and millimeter range in a space-qualified package.

 

Higher order WF sensing for coronagraphs using out-of-band light is beginning development, with data limited to computer simulations.

Relevance / Science Traceability:

These technologies are enabling for coronagraph-equipped space telescopes, segmented space telescopes, and others that utilize actively controlled optics. The  Large UV/Optical/IR Surveyor (LUVOIR) and Habitable Exoplanet Observatory (HabEx) mission concepts currently under study provide good examples.

References:

• The HabEx interim report is available at: https://www.jpl.nasa.gov/habex/  
• The LUVOIR interim report is available at: https://asd.gsfc.nasa.gov/luvoir/reports/

Scope Title:

Optical Components and Systems for Potential Infrared/Far-Infrared Missions

Scope Description:

Far-infrared surveyor mission described in NASA's Astrophysics Roadmap, "Enduring Quests, Daring Visions":

 

In the context of subtopic S2.03, the challenge is to take advantage of relaxed tolerances stemming from a requirement for long-wavelength (30 µm) diffraction-limited performance in the fully integrated optical telescope assembly to minimize the total mission cost through innovative design and material choices and novel approaches to fabrication, integration, and performance verification.

 

A far-infrared surveyor is a cryogenic far-infrared (IR) mission, which could be either a large single-aperture telescope or an interferometer. There are many common and a few divergent optical system requirements between the two architectures.

Common requirements:

• Telescope operating temperature of ~4 K.
• Telescope diffraction-limited at 30 µm at the operating temperature.
• Mirror survivability at temperatures ranging from 315 to 4 K.
• Mirror substrate thermal conductivity at 4 K of >2 W/m·K.
• Zero or low CTE mismatch between mirror substrate and backplane.

Divergent requirements:

• Large single-aperture telescope:
  • Segmented primary mirror, circular. or hexagonal.
  • Primary mirror diameter 5 to 10 m.
  • Possible 3 degree-of-freedom (tip, tilt, and piston) control of mirror segments on orbit.
• Interferometer:
  • Monolithic primary mirrors.
  • Afocal, off-axis telescope design.
  • Primary mirror diameter 1 to 4 m.

Success metrics:

• Areal cost <$500K/m².
• Areal density <15 kg/m² (<40 kg/m² with backplane).
• Production rate >2 m² per month.
• Short time span for optical system integration and test.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Prototype
• Hardware

Desired Deliverables Description: State of the Art and Critical Gaps:

Current state of the art (SOA) is represented by the Herschel Space Observatory (3.5-m monolith; SiC) and James Webb Space Telescope (6.5-m segmented primary mirror; beryllium). Technologies are needed to advance the fabrication precision and the size of the mirrors, both monolithic and segmented, beyond the current SOA.

Relevance / Science Traceability:

The technology is relevant to the Far-Infrared Surveyor mission described in NASA's Astrophysics Roadmap and prioritized in NASA's Program Annual Technology Reports for Cosmic Origins and Physics of the Cosmos. A future NASA far-IR astrophysics mission will answer compelling questions, such as:

• How common are life-bearing planets? 
• How do the conditions for habitability develop during the process of planet formation? 
• How did the universe evolve in response to its changing ingredients (buildup of heavy elements and dust over time)?

To answer these questions, NASA will need telescopes and interferometers that reach fundamental sensitivity limits imposed by astrophysical background photon noise. Only telescopes cooled to a cryogenic temperature can provide such sensitivity.

 

Novel approaches to fabrication and test developed for a far-IR astrophysics mission may be applicable to far-IR optical systems employed in other divisions of the NASA Science Mission Directorate (SMD), or to optical systems designed to operate at wavelengths shorter than the far-IR.

References:

• The Far-Infrared Surveyor is described in NASA's Astrophysics Roadmap, "Enduring Quests, Daring Visions," which can be downloaded from: https://smd-prod.s3.amazonaws.com/science-pink/s3fs-public/atoms/files/secure- Astrophysics_Roadmap_2013_0.pdf
• The Origins Space Telescope (OST) final report is at: https://asd.gsfc.nasa.gov/firs/  
• Program Annual Technology Reports (PATR) can be downloaded from the NASA Physics of the Cosmos and Cosmic Origins (PCOS/COR) Technology Development website at: https://apd440.gsfc.nasa.gov/technology/

Scope Title:

Low-Cost Compact Reflective Telescope for CubeSAT Missions

Scope Description:

The need exists for a low-cost, compact (e.g., CubeSAT-class), scalable, diffraction-limited, athermalized, off-axis reflective telescopes.  Typically, specialty optical aperture systems are designed and built as “one-offs,” which are inherently high in cost and often out of scope for smaller projects.  A Phase I would investigate current compact off-axis reflective designs and develop a trade space to identify the most effective path forward.  The work would include a strategy for aperture diameter scalability, athermalization, and low-cost fabrication.  Detailed optical designs would be developed along with detailed structural, thermal, optical performances (STOP) analyses confirming diffraction limited operation across a wide range of operational disturbances, both structural dynamic and thermal. Phase II may follow up with development of prototypes, built at multiple aperture diameters and fidelities.

 

This Scope topic solicits solutions for two applications:  near-infrared- and short-wave-infrared- (NIR/SWIR-) band communication and the Lightning Imaging Sensor.

 

NIR/SWIR optical-communication-support hardware should be assumed towards an integrated approach, including fiber optics, fast-steering mirrors, and applicable detectors.

 

The Lightning Imaging Sensor application requires a telescope that will fit inside a 6U or smaller CubeSAT with an 80° field-of-view, is diffraction limited at 500 nm (nominal), and has high spectral transmission at both 337 and 777 nm.

Expected TRL or TRL Range at completion of the Project: 2 to 4
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware
• Analysis

Desired Deliverables Description:

• An ideal Phase I deliverable would be a prototype unobscured telescope with the required performance and size, or a reviewed preliminary design and manufacturing plan that demonstrates feasibility. While detailed analysis will be conducted in Phase II, the preliminary design should address how optical, mechanical (static and dynamic), and thermal designs and performance analyses will be done to show compliance with all requirements. Past experience or technology demonstrations that support the design and manufacturing plans will be given appropriate weight in the evaluation.
• An ideal Phase II project would further advance the technology to produce a flight-qualifiable optical system with the required performance for a CubeSAT mission.  Phase I and Phase II mirror system or component deliverables would be accompanied by all necessary documentation, including the optical performance assessment and all data on processing and properties of its substrate materials. A successful mission-oriented Phase II would have a credible plan to deliver for the allocated budget a fully assembled and tested telescope assembly that can be integrated into the potential mission as well as demonstrate an understanding of how the engineering specifications of their system meets the performance requirements and operational constraints of the mission (including mechanical and thermal stability analyses).

State of the Art and Critical Gaps:

Currently, the state of the art for reflective optical system for communications applications are:

1. On-axis or axisymmetric designs are typically used for (space) optical communications and imaging, which inherently are problematic due to the central obscuration.
2. Off-axis designs provide superior optical performance due to the clear aperture, however, are rarely considered due to complex design, manufacturing, and metrology procedures needed.
3. Currently flying Lightning Imaging Sensor is a large refractive lens optimized for single-wavelength operation.  A reflective system is required for dual-wavelength operation.  Also, a compact design is required to fit inside a CubeSAT.

Relevance / Science Traceability:

Optical communications enable high-data-rate downlink of science data. The initial motivation for this scalable off-axis optical design approach is for bringing high-performance reflective optics within reach of laser communication projects with limited resources. However, this exact optical hardware is applicable for any diffraction-limited, athermalized, science imaging application. Any science mission could potentially be able to select from a “catalog” of optical aperture systems that would already have (flight) heritage and reduced risks.

References:

• An example of an on-axis design has been utilized in the Lunar Laser Communications Demonstration (LLCD): https://www.spiedigitallibrary.org/conference-proceedings- of-spie/10563/105630X/NASAs-current-activities-in-free-space-optical-communications/10.1117/12.2304175.full?SSO=1
• An example of an off-axis design is being developed by the Jet Propulsion Laboratory (JPL) for Deep Space Optical Communications (DSOC): https://www.spiedigitallibrary.org/conference-proceedings-of-spie/10096/100960V/Discovery-deep-space-optical- communications-DSOC-transceiver/10.1117/12.2256001.full
• Information about NASA's current (large-scale) Lightning Imaging Sensor can be found at: https://gpm.nasa.gov/missions/TRMM/satellite/LIS

Lead Center: GSFC

Participating Centers: JPL, MSFC

Scope Title:

X-Ray Mirror Systems Technology, Coating Technology for X-Ray-UV-OIR, and Free-Form Optics

Scope Description:

The National Academy Astro2010 Decadal Report identifies studies of optical components and ability to manufacture, coat, and perform metrology needed to enable future x-ray observatory missions.
 

The Astrophysics Decadal specifically calls for optical coating technology investment for future ultraviolet (UV), optical, exoplanet, and infrared (IR) missions, and the Heliophysics 2009 Roadmap identifies the coating technology for space missions to enhance rejection of undesirable spectral lines and improve space/solar-flux durability of extreme UV (EUV) optical coatings, as well as coating deposition to increase the maximum spatial resolution.
 

Future optical systems for NASA's low-cost missions, CubeSat, and other small-scale payloads, are moving away from traditional spherical optics to nonrotationally symmetric surfaces with anticipated benefits of free-form optics such as fast wide-field and distortion-free cameras.
 

This subtopic solicits proposals in the following three focus areas:

• X-ray manufacturing, coating, testing, and assembling complete mirror systems in addition to maturing the current technology.
• Coating technology including carbon nanotubes (CNTs) for a wide range of wavelengths from x-ray to IR (x-ray, EUV, Lyman UV (LUV), vacuum UV (VUV), visible, and IR).
• Free-form optics design, fabrication, and metrology for CubeSat, SmallSat, and various coronagraphic instruments.

Expected TRL or TRL Range at completion of the Project: 3 to 6
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Typical deliverables based on sub-elements of this subtopic:

Phase I: 

• X-ray optical mirror system: Analysis, reports, prototype.
• Coating: Analysis, reports, software, demonstration of the concept and prototype.
• Free-form optics: Analysis, design, software and hardware prototype of optical components.

Phase II:

• X-ray optical mirror system: Analysis and prototype.
• Coating: Analysis, reports, software, demonstration of the concept and prototype.
• Free-form optics: Analysis, design, software and hardware prototype of optical components

State of the Art and Critical Gaps:

This subtopic focuses on three areas of technology development:

• This work is a very costly and time consuming. Most of SOA (state of the art) requiring improvement is ~10 arcsec angular resolution. SOA straylight suppression is bulky and ineffective for wide-field of view telescopes. We seek significant reduction in both expense and time. Reduce the areal cost of telescope by 2× such that the larger collecting area can be produced for the same cost or half the cost.
• Coating technology for wide range of wavelengths from x-ray to IR (x-ray, EUV, LUV, VUV, visible, and IR). The current x-ray coating is defined by NuSTAR. Current EV is defined by Heliophysics (80% reflectivity from 60 to 200 nm). Current UVOIR is defined by Hubble. MgFI2 over coated aluminum on 2.4-m mirror. This coating has birefringence concerns and marginally acceptable reflectivity between 100 to 200 nm.
• Free-form optics design, fabrication, and metrology for package constrained imaging systems. This field is in early stages of development. Improving the optical surfaces with large field of view and fast F/#s is highly desirable.

Relevance / Science Traceability:

S2.04 supports variety of Astrophysics Division missions. The technologies in this subtopic encompasses fields of x-ray, coating technologies ranging from UV to IR, and free-form optics in preparation for Decadal missions such as HabEx, LUVOIR, and OST.
 

Optical components, systems, and stray light suppression for x-ray missions: The 2010 National Academy Decadal Report specifically identifies optical components and the ability to manufacture and perform precise metrology on them needed to enable several different future missions (Next Generation x-ray Optics, NGXO). The National Research Council (NRC) NASA Technology Roadmap Assessment ranked advanced mirror technology for new x-ray telescopes as the #1 Object C technology requiring NASA investment.
 

Free-form optics: NASA missions with alternative low-cost science and small-size payload are increasing. However, the traditional interferometric testing as a means of metrology are unsuited to free-form optical surfaces because of changing curvature and lack of symmetry. Metrology techniques for large fields of view and fast F/#s in small-size instruments is highly desirable, specifically if they could enable cost-effective manufacturing of these surfaces (CubeSat, SmallSat, NanoSat, various coronagraphic instruments).
 

Coating for x-ray, EUV, LUV, UV, visible, and IR telescopes: Astrophysics Decadal specifically calls for optical coating technology investment for: Future UV/Optical and Exoplanet missions (Habitable Exoplanet Observatory (HabEx) or Large Ultraviolet Optical Infrared Surveyor (LUVOIR)). Heliophysics 2009 Roadmap identifies optical coating technology investments for: Origins of Near-Earth Plasma (ONEP); Ion-Neutral Coupling in the Atmosphere (INCA); Dynamic Geospace Coupling (DGC); Fine-scale Advanced Coronal Transition-Region Spectrograph (FACTS); Reconnection and Micro-scale (RAM); and Solar-C Nulling polarimetry/coronagraph for exoplanet imaging and characterization, dust and debris disks, extra-galactic studies, and relativistic and nonrelativistic jet studies.

References:

The Habitable Exoplanet Observatory (HabEx) is a concept for a mission to directly image planetary systems around Sun-like stars. HabEx will be sensitive to all types of planets; however, its main goal is, for the first time, to directly image Earth-like exoplanets, and characterize their atmospheric content. By measuring the spectra of these planets, HabEx will search for signatures of habitability such as water, and be sensitive to gases in the atmosphere possibility indicative of biological activity, such as oxygen or ozone.

 

The study pages are available at:

• Habitable Exoplanet Observatory (HabEx): https://www.jpl.nasa.gov/habex/
• LUVOIR: https://asd.gsfc.nasa.gov/luvoir/
• Origins Space Telescope: https://asd.gsfc.nasa.gov/firs/
• The LYNX Mission Concept: https://wwwastro.msfc.nasa.gov/lynx/
• The Large UV/Optical/IR Surveyor (LUVOIR) is a concept for a highly capable, multiwavelength space observatory with ambitious science goals. This mission would enable great leaps forward in a broad range of science, from the epoch of re-ionization, through galaxy formation and evolution, star and planet formation, to solar system remote sensing. LUVOIR also has the major goal of characterizing a wide range of exoplanets, including those that might be habitable—or even inhabited. The LUVOIR Interim Report is available at: 
  https://asd.gsfc.nasa.gov/luvoir/
• The Origins Space Telescope (OST) is the mission concept for the Far-IR Surveyor study. NASA's Astrophysics Roadmap, Enduring Quests, Daring Visions, recognized the need for an Origins Space Telescope mission with enhanced measurement capabilities relative to those of the Herschel Space Observatory, such as a 3-order-of-magnitude gain in sensitivity, angular resolution sufficient to overcome spatial confusion in deep cosmic surveys or to resolve protoplanetary disks, and new spectroscopic capability. The community report is available at:
   https://science.nasa.gov/science-committee/subcommittees/nac-astrophysics-subcommittee/astrophysics-roadmap

Scope Title:

X-Ray Mirror Systems Technology

Scope Description:

NASA large x-ray observatory requires low-cost, ultrastable, lightweight mirrors with high-reflectance optical coatings and effective stray-light suppression. The current state of the art of mirror fabrication technology for x-ray missions is very expensive and time consuming. Additionally, a number of improvements such as 10 arcsec angular resolutions and 1 to 5 m² collecting area are needed for this technology. Likewise, the stray-light suppression system is bulky and ineffective for wide-field-of-view telescopes.
 

In this area, we are looking to address the multiple technologies including: improvements to manufacturing (machining, rapid optical fabrication, slumping, or replication technologies), improved metrology, performance prediction and testing techniques, active control of mirror shapes, new structures for holding and actively aligning of mirrors in a telescope assembly to enable x-ray observatories while lowering the cost per square meter of collecting aperture and effective design of stray-light suppression in preparation for the Decadal Survey of 2020. Additionally, we need epoxies to bond mirrors that are made of silicon. The epoxies should absorb infrared (IR) radiation (with wavelengths between 1.5 and 6 µm that traverses silicon with little or no absorption) and therefore can be cured quickly with a beam of IR radiation. Currently, x-ray space mirrors cost $4 million to $6 million per square meter of optical surface area. This research effort seeks a cost reduction for precision optical components by 5 to 50 times, to less than $1M to $100K per square meter.
 

Additionally, proposals are solicited to develop new advanced-technology computer-numerical-control (CNC) machines to polish inside and/or outside surfaces of full-shell (between 100 and 1,000 mm in height, 100 to 2,800 mm in diameter, varying radial prescription along azimuth, and ~2 mm in thickness), grazing-incidence optics to x-ray quality surface tolerances (with surface figure error <1 arcsec half-power diameter (HPD), radial slope error <1 µrad, and out-of-round <2 µm). Current state-of-the-art technology in CNC polishing of full-shell, grazing-incidence optics yields 2.5 arcsec HPD on the outside of a mandrel used for replicating shells. Technology advances beyond current state of the art include application of CNC and deterministic polishing techniques that (1) allow for direct force closed-loop control, (2) reduce alignment precision requirements, and (3) optimize the machine for polishing cylindrical optics through simplifying the axis arrangement and the layout of the cavity of the CNC polishing machine.

Expected TRL or TRL Range at completion of the Project: 3 to 6
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Typical deliverable based on sub-elements of this subtopic:
X-ray optical mirror system—Demonstration, analysis, reports, software, and hardware prototype:

• Phase I deliverables:  Reports, analysis, demonstration, and prototype
• Phase II deliverables: Analysis, demonstration, and prototype

State of the Art and Critical Gaps:

X-ray optics manufacturing, metrology, coating, testing, and assembling complete mirror systems in addition to maturing the current technology. This work is very costly and time-consuming. Most of the SOA (state of the art) requiring improvement is ~10 arcsec angular resolution. SOA stray-light suppression is bulky and ineffective for wide-field of view telescopes. We seek a significant reduction in both expense and time. Reduce the areal cost of a telescope by 2× such that the larger collecting area can be produced for the same cost or half the cost.
 

The gaps to be covered in this track are:

• Lightweight, low-cost, ultrastable mirrors for large x-ray observatory.
• Stray-light suppression systems (baffles) for large advanced x-ray observatories.
• Ultrastable, inexpensive lightweight x-ray telescope using grazing-incidence optics for high-altitude balloon-borne and rocket-borne mission.

Relevance / Science Traceability:

The 2010 National Academy Decadal Report specifically identifies optical components and the ability to manufacture and perform precise metrology on them needed to enable several different future missions (Lynx and Advanced X-ray Imaging Satellite (AXIS)).
 

The National Research Council NASA Technology Roadmap Assessment ranked advanced mirror technology for new x-ray telescopes as the #1 Object C technology requiring NASA investment.

References:

NASA High Energy Astrophysics (HEA) mission concepts including x-ray missions and studies are available at:

• https://heasarc.gsfc.nasa.gov/docs/heasarc/missions/concepts.html

Scope Title:

Coating Technology for X-Ray-UV-OIR

Scope Description:

The optical coating technology is a mission-enabling feature that enhances the optical performance and science return of a mission. Lowering the areal cost of coating determines if a proposed mission could be funded in the current cost environment. The most common forms of coating used on precision optics are antireflective (AR) coating and high-reflective (HR) coating.
 

The current coating technology of optical components needed to support the 2020 Astrophysics Decadal process. Historically, it takes 10 years to mature mirror technology from TRL 3 to 6.

To achieve these objectives requires sustained systematic investment.
 

The telescope optical coating needs to meet low-temperature operation requirement. It’s desirable to achieve 35 K in future.
 

A number of future NASA missions require suppression of scattered light. For instance, the precision optical cube utilized in a beam-splitter application forms a knife-edge that is positioned within the optical system to split a single beam into two halves. The scattered light from the knife-edge could be suppressed by carbon nanotube (CNT) coating. Similarly, the scattered light for gravitational-wave application and lasercom system where the simultaneous transmit/receive operation is required, could be achieved by a highly absorbing coating such as CNT. Ideally, the application of CNT coating needs to:

• Achieve broadband (visible plus near infrared (IR)) reflectivity of 0.1% or less.
• Resist bleaching of significant albedo changes over a mission life of at least 10 years.
• Withstand launch conditions such as vibration, acoustics, etc.
• Tolerate both high continuous-wave (CW) and pulsed power and power densities without damage: ~10 W for CW and ~0.1 GW/cm² density, and 1-kW/nsec pulses.
• Adhere to the multilayer dielectric or protected metal coating, including ion beam sputtering (IBS) coating.

NASA's Laser Interferometer Space Antenna (LISA) mission on-axis design telescope operates both in transmission and reception simultaneously where the secondary mirror sends the transmitted beam directly back at the receiver. The apodized petal-shaped mask inherently suppress the diffraction once patterned at the center of the secondary mirror.  The emerging cryogenic etching of black-silicon has demonstrated bidirectional reflectance distribution function (BRDF) ultralow specular reflectance of 1×10^-7 in the range of 500 to 1,064 nm. The advancement of this technology is desired to obtain ultralow reflectivity.

• Improve the specular reflectance to 1×10^-10 and hemispherical reflectance better than 0.1%.
• Improve the cryogenic etching process to provide a variation of the reflectance (apodization effect) by increasing or decreasing the height of the grass.
• Explore etching process and duration.

Expected TRL or TRL Range at completion of the Project: 3 to 6
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Coating—Analysis, reports, software, demonstration of the concept, and prototype:

• Phase I deliverables: Report, analysis, demonstration, and prototype.
• Phase II deliverables: Analysis, demonstration and prototype.

State of the Art and Critical Gaps:

Coating Technology (for wide range of wavelengths from x-ray to IR: x-ray, extended UV (EUV), Lyman UV (LUV), vacuum UV (VUV), visible, and IR):

• The current x-ray coating is defined by Nuclear Spectroscopic Telescope Array (NuSTAR).
• Current EUV is defined by Heliophysics (80% reflectivity from 60 to 200 nm).
• Current UVOIR is defined by Hubble. MgFI2 overcoated aluminum on 2.4-m mirror. This coating has birefringence concerns and marginally acceptable reflectivity between 100 and 200 nm.

Metrics for X-Ray:

• Multilayer high-reflectance coatings for hard x-ray mirrors.
• Multilayer depth gradient coatings for 5 to 80 keV with high broadband reflectivity.
• Zero-net-stress coating of iridium or other high-reflectance elements on thin substrates (<0.5 mm).

Metrics for EUV:

• Reflectivity >90% from 6 to 90 nm onto a <2 m mirror substrate.

Metrics for Large UV/Optical/IR Surveyor (LUVOIR):

• Broadband reflectivity >70% from 90 to 120 nm (LUV) and >90% from 120 nm to 2.5 µm (VUV/visible/IR).
• Reflectivity non-uniformity <1% 90 nm to 2.5 µm.
• Induced polarization aberration <1% 400 nm to 2.5 µm spectral range from mirror coating applicable to a 1- to 8-m substrate.

Metrics for LISA:

• HR: Reflectivity >99% at 1,064+/-2 nm with very low scattered light and polarization-independent performance over apertures of ~0.5 m.
• AR: Reflectivity <0.005% at 1,064+/-2 nm.
  • Low-absorption, low-scatter, laser-line optical coatings at 1,064 nm.
  • High reflectivity, R > 0.9995.
  • Performance in a space environment without significant degradation over time, due for example to radiation exposure or outgassing.
  • High polarization purity, low optical birefringence over a range of incident angles from ~5° to ~20°.
  • Low coating noise (thermal, photothermal, etc.) for high-precision interferometric measurements.
  • Ability to endure applied temperature gradients (without destructive effects, such as delamination from the substrate).
  • Ability to clean and protect the coatings and optical surfaces during mission integration and testing. Cleaning should not degrade the coating performance.

Nonstationary Optical Coatings:

• Used in reflection and transmission that vary with location on the optical surface.

CNT Coatings:

• Broadband visible to NIR, total hemispherical reflectivity of 0.01% or less, adhere to the multilayer dielectric or protected metal coating.

Black-Silicon Cryogenic Etching (new):

• Broadband UV+visible+NIR+IR, reflectivity of 0.01% or less, adhere to the multilayer dielectric (silicon) or protected metal.

Software tools to simulate and assist the anisotropic etching by employing variety of modeling techniques such as rigorous coupled wave analysis (RCWA), method of moments (MOM), finite-difference time domain (FDTD), finite element method (FEM), transfer matrix method (TMM), and effective medium theory (ETM).

Relevance / Science Traceability:

• Coating for x-ray, EUV, LUV, UV, visible, and IR telescopes: Astrophysics Decadal specifically calls for optical coating technology investment for: Future UV/optical and exoplanet missions.
• Heliophysics 2009 Roadmap identifies optical coating technology investments for: Origins of Near-Earth Plasma (ONEP), Ion-Neutral Coupling in the Atmosphere (INCA), Dynamic Geospace Coupling (DGC), Fine-scale Advanced Coronal Transition-Region Spectrograph (FACTS), Reconnection and Micro-scale (RAM), and Solar-C.
• LISA requires low-scatter HR coatings and low reflectivity coatings for scatter suppression near 1064 nm. Polarization-independent performance is important.
• Nulling polarimetry/coronagraphy for exoplanets imaging and characterization, dust and debris disks, extra-galactic studies, and relativistic and nonrelativistic jet studies.

References:

Laser Interferometer Space Antenna (LISA) is a space-based gravitational wave observatory building on the success of LISA Pathfinder and Laser Interferometer Gravitational-Wave Observatory (LIGO). Led by the European Space Agency (ESA), the new LISA mission (based on the 2017 L3 competition) is a collaboration between ESA and NASA.

• More information can be found at: https://lisa.nasa.gov

Scope Title:

Free-Form Optics

Scope Description:

Future NASA science missions demand wider fields of view in a smaller package. These missions could benefit greatly by free-form optics as they provide nonrotationally symmetric optics, which allow for better packaging while maintaining desired image quality. Currently, the design and fabrication of free-form surfaces is costly. Even though various techniques are being investigated to create complex optical surfaces, small-size missions highly desire efficient small packages with lower cost that increase the field of view and expand operational temperature range of un-obscured systems. In addition to the free-form fabrication, the metrology of free-form optical components is difficult and challenging because of the large departure from planar or spherical shapes accommodated by conventional interferometric testing. New methods such as multibeam low-coherence optical probe and slope sensitive optical probe are highly desirable.
 

Specific metrics are:

• Design: Innovative design methods/tools for free-form systems, including applications to novel reflective optical designs with large fields of view (>30°) and fast F/#s (<2.0).
• Fabrication: 10-cm-diameter optical surfaces (mirrors) with free-form optical prescriptions >1 mm, spherical departure with surface figure error <10 nm rms, and roughness <5 Angstroms.  10-cm-diameter blazed optical reflective gratings on free-form surface shapes with >1 mm departure from a best-fit-sphere, and grating spacings from 1 to 100 µm.  Larger mirrors are also desired for flagship missions for ultraviolet (UV) and coronagraphy applications, with 10-cm- to 1-m-diameter surfaces having figure error <5 nm rms and roughness <1 Angstroms rms.
• Metrology: Accurate metrology of "free-form" optical components with large spherical departures (>1 mm), independent of requiring prescription-specific null lenses or holograms.

Expected TRL or TRL Range at completion of the Project: 3 to 6
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Optical components—Demonstration, analysis, design, metrology, software, and hardware prototype:

• Phase I deliverables: Report, analysis, demonstration, and prototype.
• Phase II deliverables: Analysis, demonstration, and prototype.

State of the Art and Critical Gaps:

Free-form optics design, fabrication, and metrology for package constrained imaging systems. This field is in early stages of development. Improving the optical surfaces with large field-of-view and fast F/#s is highly desirable.

Relevance / Science Traceability:

NASA missions with alternative low-cost science and small-size payload are increasing. However, the traditional interferometric testing as a means of metrology is unsuited to freeform optical surfaces due to changing curvature and lack of symmetry. Metrology techniques for large fields-of-view and fast F/#s in small size instruments are highly desirable specifically if they could enable cost-effective manufacturing of these surfaces. (CubeSat, SmallSat, and NanoSat). Additionally, design studies for large observatories such as Origins Space Telescope (OST) and Large UV/Optical/IR Surveyor (LUVOIR, currently being proposed for the 2020 Astrophysics Decadal Survey) have demonstrated improved optical performance over a larger field-of-view afforded by free-form optics. Such programs will require advances in free-form metrology to be successful.

References:

A presentation on application of Freeform Optics at NASA is available at: 

• Applications for Freeforms Optics at NASA, https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170010419.pdf
• Alignment and Testing for a Freeform Telescope, https://ntrs.nasa.gov/citations/20180007557
• Freeform Surface Characterization and Instrument Alignment for Freeform Space Applications, https://ntrs.nasa.gov/citations/20190025929

 

Lead Center: JPL

Participating Centers: GSFC

Scope Title:

Components, Assemblies, and Subsystems for Extreme Precision Radial Velocity Measurements and Detection of Extrasolar Planets

Scope Description:

Astronomical spectrographs have proven to be powerful tools for exoplanet searches. When a star experiences periodic motion due to the gravitational pull of an orbiting planet, its spectrum is Doppler modulated in time. This is the basis for the precision radial velocity (PRV) method, one of the first and most efficient techniques for detecting and characterizing exoplanets. Because spectrographs have their own drifts, which must be separated from the periodic Doppler shift, a stable reference is always needed for calibration. Optical frequency combs (OFCs) and line-referenced etalons are capable of providing the spectral rulers needed for PRV detection of exoplanets. Although “stellar jitter” (a star’s photospheric velocity contribution to the RV signal) is unavoidable, the contribution to the error budget from Earth’s atmosphere would be eliminated in future space missions. Thus, there is a need to develop robust spectral references, especially at visible wavelengths to detect and characterize Earth-like planets in the habitable zone of their Sun-like host stars, with size, weight, and power (SWaP) suitable for space-qualified operation to calibrate the next generation of high-resolution spectrographs with precision corresponding to <~1 cm/s over multiple years of observations.

 

This subtopic solicits proposals to develop cost-effective component and subsystem technology for low-SWaP, long-lived, robust implementation of RV measurement instruments both on the ground and in space. Research areas of interest include but are not limited to:

• Integrated photonic spectrographs.
• Spectrograph gratings.
• PRV spectrograph calibration sources.
• High efficiency photonic lanterns.
• Advanced optical fiber delivery systems and subsystems with high levels of image scrambling and modal noise reduction.
• Software for advanced statistical techniques to mitigate effects of telluric absorption and stellar jitter on RV precision and accuracy.

Expected TRL or TRL Range at completion of the Project: 3 to 5
Primary Technology Taxonomy:
    Level 1: TX 08 Sensors and Instruments
    Level 2: TX 08.2 Observatories
Desired Deliverables of Phase I and Phase II:

• Hardware
• Software

Desired Deliverables Description: State of the Art and Critical Gaps:

High-resolving-power spectrographs (R ~ 150,000) with simultaneous UV, visible, and NIR coverage and exquisite long-term stability are required for PRV studies. Classical bulk optic spectrographs traditionally used for PRV science impose architectural constraints due to their large mass and limited optical flexibility. Integrated photonic spectrographs are wafer-thin devices that could reduce instrument volume by up to 3 orders of magnitude. Spectrometers that are fiber fed, with high illumination stability, excellent wavelength calibration, and precise temperature and pressure control represent the immediate future of precision RV measurements.

 

Traditional RV spectrographs would benefit from improvements in grating technology. Diffraction-limited PRV spectrographs require echelle gratings with low wavefront error and high efficiency, both of which are very challenging to achieve. Echelle spectrographs are designed to operate at high angle-of-incidence and very high diffraction order. Hence, the grating must have very accurate groove placement (for low wavefront error) and very flat groove facets (for high efficiency). For decades, echelle gratings have been fabricated by diamond ruling, but it is difficult to achieve all aspects of the performance required for PRV instruments. Newer grating fabrication techniques using lithographic methods to form the grooves may be a promising approach. As spectrograph stability imposes limits on how precisely RV can be measured, spectral references play a critical role in characterizing and ensuring this precision. Only laser frequency combs (LFCs) and line-referenced Fabry-Pérot etalons are capable of providing the broad spectral coverage and long-term stability needed for extreme PRV detection of exoplanets. Although both frequency combs and etalons can deliver high-precision spectrograph calibration, the former requires relatively complex hardware in the visible portion of the spectrum.

 

Commercial fiber laser astrocombs covering 450 to 1400 nm at 25 GHz line spacing and <3-dB intensity variations over the entire bandwidth are available for ground-based astronomical spectrographs. However, the cost for these systems is often so prohibitive that recent RV spectrograph projects either do not use a LFC or include it only as a future upgrade. Alternatively, astrocombs produced by electro-optic modulation (EOM) of a laser source have been demonstrated in the NIR. EOM combs produce modes spaced at a radiofrequency (RF) modulation frequency, typically 10 to 30 GHz. Significantly, EOM combs avoid the line filtering step required by commercial mode-locked fiber laser combs. Comb frequency stabilization can be accomplished by referencing the laser pump source to a molecular absorption feature or another frequency comb. Where octave spanning EOM combs are available, f-2f self-referencing provides the greatest stability. EOM combs must be spectrally broadened to provide the bandwidth necessary for PRV applications. This is accomplished through pulse amplification followed by injection into highly nonlinear fiber or nonlinear optical waveguides. 

Power consumption of the frequency comb calibration system will be a significant driver of mission cost for space-based PRV systems and motivates the development of a comb system that operates with less than 20 W of spacecraft power. Thus, for flight applications, it is highly desirable to develop frequency comb technology with low power consumption; ~10 to 30 GHz mode spacing; compact size; broad (octave spanning) spectral grasp across both the visible and NIR; low phase noise; stability traceable to the International System of Units definition of the second; and importantly, long life.

 

The intrinsic illumination stability of the spectrometer also sets a fundamental measurement floor. As the image of the star varies at the entrance to the spectrometer because of atmospheric effects and telescope guiding errors, so too does the recorded stellar spectrum, leading to a spurious RV offset. Current seeing-limited PRV instruments use multimode optical fibers, which provide some degree of azimuthal image scrambling, to efficiently deliver stellar light from the telescope focal plane to the spectrometer input. Novel-core-geometry fibers, in concert with dedicated optical double-scramblers, are often used to further homogenize and stabilize the telescope illumination pattern in both the image and pupil planes. However, these systems still demonstrate measurable sensitivity to incident illumination variations from the telescope and atmosphere. Furthermore, as spectral resolution requirements increase, the commensurate increase in instrument size becomes impractical. Thus, the community has turned to implementing image and pupil slicers to reformat the near or far fields of light entering the spectrometer by preferentially redistributing starlight exiting the fiber to maintain high spectral resolution, efficiency, and compact spectrometer size.

Relevance / Science Traceability:References:

Precision Radial Velocity:

• Fischer et al. (2016): State of the Field: Extreme Precision Radial Velocities, http://adsabs.harvard.edu/abs/2016PASP..128f6001
• Plavchan et al. (2015): Radial Velocity Prospects Current and Future: A White Paper Report prepared by the Study Analysis Group 8 for the Exoplanet Program Analysis Group (ExoPAG), http://adsabs.harvard.edu/abs/2015arXiv150301770P
• Plavchan et al. (2019): EarthFinder Probe Mission Concept Study (Final Report), https://smd-prod.s3.amazonaws.com/science-red/s3fs-public/atoms/files/Earth_Finder_Study_Rpt.pdf

Photonic Lanterns:

• Gris-Sanchez et al. (2018): Multicore fibre photonic lanterns for precision radial velocity Science, https://academic.oup.com/mnras/article/475/3/3065/4769655
• Jvanovic, N. et al. (2012): Integrated photonic building blocks for next-generation astronomical instrumentation I: the multimode waveguide. Optics Express, 20:17029.

Astrocombs:

• Yi, X. et al. (2016): Demonstration of a near-IR line-referenced electro-optical laser frequency comb for precision radial velocity measurements in astronomy. Nature Communications, 7:10436.
• Halverson, S. et al. (2014): The habitable-zone planet finder calibration system. Proc. SPIE 9147, Ground-based and Airborne Instrumentation for Astronomy V, 91477Z, https://doi.org/10.1117/12.2054967
• Suh, M.-G., et al. (2019): Searching for exoplanets using a microresonator astrocomb. Nature Photonics, 13(1):25–30.
• Obrzud, E., et al. (2019): A Microphotonic Astrocomb. Nature Photonics, 13 (1):31–35.
• Metcalf, A., et al. (2019): Stellar Spectroscopy in the Near-infrared with a Laser Frequency Comb, Optica, Vol 6, issue 2.
• Lee, S.H., Oh, D.Y., Yang, Q. et al. (2017): Towards visible soliton microcomb generation. Nat Commun 8, 1295.

Nonlinear Waveguides:

• Chang, L., et al. (2018): Heterogeneously integrated GaAs waveguides on insulator for efficient frequency conversion. Laser Photonics Reviews, 12, 1800149, https://doi.org/10.1002/lpor.201800149
• Halir, R., et al. (2012): Ultrabroadband supercontinuum generation in a CMOS-compatible platform. Optics letters, 37, 1685, https://di.org/10.1364/OL.37.001685

Spectral Flattening:

• Probst, R.A., et al. (2015): Spectrally Flattened, Broadband Astronomical Frequency Combs, https://di.org/10.1364/CLEO_SI.2015.SW4G.7

his focus area includes tools and technologies that contribute to both the Advanced Air Vehicles Program (AAVP) and the Transformative Aeronautics Concepts Program (TACP) encompassing technologies in all six Strategic Thrusts within the NASA Aeronautics Mission Directorate (ARMD). AAVP studies, evaluates and develops technologies and capabilities for new aircraft systems, and also explores far-future concepts that hold promise for revolutionary air-travel improvements. Innovative AAVP design concepts for advanced vehicles integrate technologies focus on fuel burn, noise, emissions and intrinsic safety. The goal: to enable new aircraft to fly safer, faster, cleaner, quieter, and use fuel far more efficiently. Partnering with industry, academia, and other government agencies, AAVP pursues mutually beneficial collaborations to leverage opportunities for effective technology transition. TACP encourages revolutionary concepts, creates the environment for researchers to experiment with new ideas, performs ground and small-scale flight tests, and drives rapid turnover into potential future concepts to enable aviation transformation. Research is organized to aggressively engage both the traditional aeronautics community and non-traditional partners. Although TACP focuses on sharply focused studies, the program provides flexibility for innovators to assess new-technology feasibility and provide the knowledge base for radical aeronautics advances.in noise reduction technology.

Lead Center: ARC

Participating Centers: AFRC, LaRC

Scope Title:

Integration of Flight Control with Aircraft Multidisciplinary Design Optimization

Scope Description:

Successful design tools for advanced aircraft concepts typically require close interaction among the various design disciplines involved: aerodynamics, structures, propulsion, flight control, etc. This is particularly true for advanced modern transports such as the Boeing 787 and nontraditional aircraft concepts currently being studied at NASA, such as the Transonic Truss-Braced Wing or the blended wing body, where the level of coupling across the various disciplines can be very complex. It has been shown that substantial benefits could be realized by including many of these disciplines within a single multidisciplinary design optimization (MDO) process, rather than optimizing each discipline in isolation. Furthermore, nontraditional flight control design philosophies are becoming increasingly more common as the aircraft industry is responding to the new urban air mobility (UAM) market. The NASA X-57, for example, could employ many modes of flight control distributed throughout the aircraft enabled by distributed electric propulsion. Thus, integration of flight control into an aircraft MDO process can produce more advanced aircraft design capabilities through improved and integrated methods and tools.

 

This subtopic seeks proposals that develop new methods and tools that enable the inclusion of flight control in an aircraft MDO process. The proposed control approach should directly contribute to aircraft overall design objectives, such as fuel burn, aircraft takeoff gross weight, or other aircraft performance metrics. The resulting tightly coupled MDO environment should be capable of handling the interactions among relevant disciplines, such as structures, aerodynamics, flight control (including the control effector design parameters and actuator weight/power), and propulsion. Design variables may include structural details (e.g., skin thickness, structural layout, and materials), aerodynamic details (planform shape, jig twist, and airfoil shape distributions), control law details (type, objective function, and control gains), control effector and sensor design parameters (size, layout, locations, and quantity), propulsion design parameters (thrust distribution, integrated allocation, and scheduling), etc. The control law may be parameterized in such a way that the resulting parameters are suitable to a numerical optimizer. Control design tools may utilize advanced distributed sensors such as fiber optic shape sensors or surface pressure sensors to achieve improved aircraft design objectives.

 

Appropriate design constraints should be imposed within the integrated MDO tool to ensure proper aerodynamic limits, such as stall and buffet; structural limits, such as maneuver and gust load response and flutter margins; as well as control-centric constraints, such as stability margins, control actuator power and authority, position and rate limits, and sensor and actuator requirements that dictate controllability and observability.

 

Applications of these integrated MDO approaches could be implemented for realistic and relevant aircraft configurations to demonstrate the impact of the proposed methods. Performance gains should be actively derived from the system response to a control policy such as cruise-drag optimization, gust load alleviation, flutter suppression, flight-propulsion control, and multiobjective flight control. To integrate flight control into an MDO environment with a large number of design variables and constraints, new tools that would enable increased computational efficiency, such as adjoint-based optimization, are of interest using appropriate-fidelity aircraft models capable of capturing nonlinear aerodynamics. The proposed tools could be stand-alone or added capabilities to available open-sourced NASA MDO environments.

 

This subtopic does not seek proposals that propose vehicle conceptual design and analysis studies or control methods that have no clear transition paths to commercial applications.

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

• Research
• Analysis
• Software
• Prototype
• Hardware

Desired Deliverables Description:

For Phase I, the deliverables include flight control integrated MDO methods and tools.
 

For Phase II, the deliverables include more mature and validated flight control integrated MDO methods and tools which are to be integrated with open-sourced NASA MDO environments.

State of the Art and Critical Gaps:

Increasingly, more advanced aircraft concepts utilize innovative flight control design philosophies, such as the modern Boeing 787 with trailing edge variable camber, and Airbus A350 with adaptive drooped hinge flaps, as well as a wide variety of UAM aircraft design concepts with distributed electric propulsion. The current state of the art in research does not usually consider integration of flight control systems into an aircraft MDO process. Some low-level research effort of addressing flight control integration in an MDO process exists within the Aeronautics Research Mission Directorate (ARMD) Advanced Air Transport Technology (AATT) Project, but the effort does not consider a wide range of integration of flight control systems including actuators, sensors, and flight control laws. This critical research gap area could be filled by this subtopic.

Relevance / Science Traceability:

Under the NASA Advanced Air Vehicle Program (AAVP), the AATT Project is conducting research in distributed electric propulsion and adaptive wing technologies. Both of these research elements could benefit from this subtopic. Also, under the NASA AAVP, the Revolutionary Vertical Lift Technologies (RVLT) Project is conducting research in the area of UAM aircraft using distributed electric propulsion for electrical vertical takeoff and landing (eVTOL). This subtopic could complement the research conducted under the RVLT Project.

References:

NASA Aeronautics Research Mission Directorate (ARMD), https://www.nasa.gov/aeroresearch

Advanced Air Vehicles Program (AAVP), https://www.nasa.gov/aeroresearch/programs/aavp

Advanced Air Transport Technology (AATT) project, https://www.nasa.gov/aeroresearch/programs/aavp/aatt

Revolutionary Vertical Lift Technology (RVLT) project, https://www.nasa.gov/aeroresearch/programs/aavp/rvlt

Lead Center: LaRC

Participating Centers: GRC

Scope Title:

Airframe Noise Reduction

Scope Description:Expected TRL or TRL Range at completion of the Project: 2 to 5
Primary Technology Taxonomy:
    Level 1: TX 15 Flight Vehicle Systems
    Level 2: TX 15.1 Aerosciences
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Concepts, technologies, and tools that enable rapid assessment of the noise impact of novel engine/airframe configurations, mitigate component noise issues associated with novel aircraft configurations, and/or aid in the development and optimization of noise control approaches for component noise sources that enable new aircraft configurations such as truss-braced wing and small-core turbofan engines. Example Phase I deliverables include laboratory demonstrations that establish proof of concept of noise reduction technologies, or applications of novel computational tools with limited scope that demonstrate the potential for success on problems of greater scope. Example Phase II deliverables include system or subsystem demonstrations concurrent with the establishment of a realistic path to concept production, or incorporation of novel computational tools into existing modeling toolchains with validation cases to document capabilities.

State of the Art and Critical Gaps:

State-of-the-art technologies for noise reduction on conventional transport aircraft are generally passive and do not incorporate advanced material systems or adaptive mechanisms that can modify their performance based on the noise state of the vehicle. Advanced material systems for airframe noise control are still in their infancy, especially in the context of certifiability and robustness. Novel material systems that could be applied to component noise sources on the aircraft are needed, such as shape memory alloy actuators, or active or adaptive systems. In addition, future aircraft designs, such as high-aspect ratio, truss-braced configurations, are envisioned that either leverage noise benefits of complex geometrical configurations or introduce noise challenges with engine/airframe integration. Efficient computational tools that enable rapid-turn evaluations of multiple configurations at the design stage are lacking. Numerical methods to study complex engine/airframe configurations are complex and difficult to leverage at the aircraft design stage where configuration details are not specified. Improvements to numerical methods and models for studying the noise aspects of advanced airframe configurations, including engine integration, would ease consideration of acoustics in the design, rather than leaving acoustics to the late design stage where noise control solutions are costly and less effective. Improved tools would also enable more rapid evaluation and development of novel noise control approaches that may be needed for these novel aircraft configurations.    

Relevance / Science Traceability:

AAVP: The Advanced Air Transport Technology (AATT) and Commercial Supersonic Technology (CST) Projects would benefit from noise reduction technologies that could reduce the aircraft noise footprint at landing and takeoff. Configurations with novel engine placement, such as above the fuselage, can reduce the noise footprint, but technologies are needed to efficiently model the performance and noise impacts of these novel engine installations. In addition, novel configurations and technologies such as truss-braced wing and small-core turbofan engines will introduce new noise challenges that must be addressed to enable their successful deployment.

 

TACP: The Transformational Tools and Technologies (TTT) Project would benefit from tool developments to enhance the ability to consider acoustics earlier in the aircraft design process. The TTT project would also benefit from the development and demonstration of simple material systems, such as advanced liner concepts with reduced drag or adaptive material and/or structures that reduce noise, as these component technologies could have application in numerous vehicle classes in the AAVP portfolio, including subsonic and supersonic transports as well as vertical lift vehicles.

References:

AAVP - Advanced Air Transport Technology (AATT) Project: https://www.nasa.gov/aeroresearch/programs/aavp/aatt

AAVP - Commercial Supersonic Technology (CST) Project: https://www.nasa.gov/aeroresearch/programs/aavp/cst

TACP - Transformational Tools and Technologies (TTT) Project: https://www.nasa.gov/aeroresearch/programs/tacp/ttt

Lead Center: GRC

Participating Centers: None

Scope Title:

Advanced Materials and Structures Technologies Enabling New Highly Efficient Propulsion Systems for Subsonic Transport Vehicles

Scope Description:

Materials and structures research and development (R&D) contributes to NASA’s ability to achieve its long-term aeronautics goals, including the development of advanced propulsion systems. Proposals are sought for advanced materials and structures technologies that will be enabling for new propulsion systems for subsonic transport vehicles with high levels of thermal, transmission, and propulsive efficiency. Integrated computational and experimental approaches are needed that can reduce the time necessary for development, testing, and validation of new materials systems and components.
 

Advanced high-pressure-ratio compact gas turbine engines will include components of sufficiently compact size that new approaches to processing and advanced manufacturing will be needed. Temperature capability, thermomechanical performance, environmental durability, reliability, and cost-effectiveness are important considerations. 
 

The increased use of various types of modeling to improve R&D effectiveness and enable more rapid and revolutionary materials design has been identified as critical. NASA recently sponsored a study to define a potential 25-year goal for integrated, multiscale modeling of materials and systems to accelerate the pace and reduce the expense of innovation in future aeronautical systems. Through a series of surveys, workshops, and validation exercises, this study identified critical cultural changes and gaps facing the multiscale modeling community. The results of this study were published in a NASA report, "Vision 2040: A Roadmap for Integrated, Multiscale Modeling and Simulation of Materials and Systems" [Ref. 1]. Some of the critical gaps identified in this report are: (1) under-development of physics-based models that link length and time scales, (2) inability to conduct real-time characterization at appropriate length and time scales, (3) lack of optimization methods that bridge scales, (4) lack of models that compute input sensitivities and propagate uncertainties, and (5) lack of verification and validation methods and data. 
 

Proposals emphasizing modeling can address topics which shall address gaps in that 2040 vision. The range of topics could include data management, data analytics, machine learning, linkage and integration across spatiotemporal scales, and characterization of materials over their lifecycle. Proposals may address any material class associated with aeronautics propulsion for subsonic transport vehicles, multiscale modeling and measurements, multiscale optimization methods, and verification and validation of models and methods. However, approaches should rely on iterative, predictive methods that integrate experiments and simulations to describe the behavior and response of materials at various length and time scales.

Technology areas of interest this year include:

• Computational materials and multiscale modeling tools, including methods to predict properties, and/or durability of propulsion materials based upon chemistry and processing for conventional as well as functionally graded, nanostructured, multifunctional, and adaptive materials.
• Robust and efficient methods/tools to design and model advanced propulsion system materials and structures at all scale levels, including approaches that are adaptable for a multiscale framework. 
• Multiscale design tools that integrate novel materials, mechanism design, and structural subcomponent design into system level designs.
• Advancing technology for ceramic matrix composites (CMCs) and their environmental barrier coatings (EBCs) for gas turbine engine components operating at 1,482 °C (2,700 °F) or higher. Focus areas include increased thermomechanical durability, increased resistance to environmental interactions, cost-effectiveness of processing and manufacturing, and improved approaches to component fabrication and integration. Computational tools and integrated experimental/computational methods are sought, including models/tools to predict degradation and failure mechanisms.
• Additive manufacturing and other advanced processing/manufacturing approaches for structural components or materials to enable improved engine efficiency through decreasing weight and/or improving component design, properties, and performance.
• In support of future aircraft with hybrid electric or all electric propulsion systems, advanced cross-cutting materials technologies are needed. For example: (1) soft magnetic material with high magnetic saturation and/or lower losses for 100 to 300 kHz operation, (2) hard magnetic materials with an energy product greater than neodymium iron boron, (3) conductors for power cables with a specific resistivity less than copper or aluminum, and (4) novel materials systems and structures to enable functionality, such as power harvesting, thermal management, self-sensing, and actuation.
• Design and development of unique materials such as shape memory alloys and high entropy alloys for subsonic transport vehicle propulsion system structures and components. 
• Propulsion aeromechanics, damping devices, and analysis and mistuning analysis for turbomachinery rotating blades.

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:

• Research
• Analysis
• Prototype
• Software

Desired Deliverables Description:

NASA’s intent is to select proposals that have the potential to move a critical technology beyond Phase II SBIR funding and transition it to Phase III, where NASA’s aeronautics programs, another Government agency, or a commercial entity in the aeronautics sector can fund further maturation as-needed, leading to actual usage in an enhanced propulsion system. The Phase I outcome should establish the scientific, technical, and commercial feasibility of the proposed innovation in fulfillment of NASA needs. Phase I should demonstrate advancement of a specific technology, supported by analytical and experimental studies that are documented in a final report. Phase IIs could yield: (1) models supported with experimental data, (2) software related to a model that was developed, (3) a material system or subcomponent that has been demonstrated to have better properties/performance (ability to operate at a higher temperature, carry more current, etc.), and (4) modeling tools for incorporation in software, etc. that can be infused into a NASA project or lead to commercialization of the technology. Consequently, Phase II efforts are strengthened when they include a partnership with a potential end-user of the technology.  

State of the Art and Critical Gaps:

This subtopic would support R&D on advanced materials and structures technologies that will be enabling for new propulsion systems for subsonic transport vehicles with high levels of thermal, transmission, and propulsive efficiency. The needs are specified in the scope description. One of the major NASA Glenn Research Center core competencies is Materials and Structures for Extreme Environments. This subtopic supports that type of research—enabling materials and structures research that allows more efficient propulsion systems.

In general, integrated computational and experimental approaches are needed that can reduce the time necessary for development, testing, and validation of new materials systems and components. The increased use of various types of modeling to improve R&D effectiveness and enable more rapid and revolutionary materials design has been identified as critical. NASA recently conducted a study that identified critical cultural changes and gaps facing the multiscale modeling community.

For future aircraft with hybrid electric or all electric propulsion systems, advanced materials technology is needed for power components including electric machines and power cables.

Advanced high-pressure-ratio compact gas turbine engines will include components of sufficiently compact size that new approaches to processing and advanced manufacturing will be needed. Improvements in temperature capability, thermomechanical performance, environmental durability, reliability, and cost-effectiveness are important considerations.

Relevance / Science Traceability:

Aeronautics Research Mission Directorate (ARMD) projects that would/could support each of the specified areas of interest are listed below, along with advocates for the technologies. The technologies would lead to improved propulsion efficiencies (subsonic transport vehicles).

• Computational materials and multiscale modeling tools, including methods to predict properties, and/or durability of propulsion materials based upon chemistry and processing for conventional as well as functionally graded, nanostructured, multifunctional, and adaptive materials. TTT (Transformational Tools and Technology) Project. 
• Robust and efficient methods/tools to design and model advanced propulsion system materials and structures at all scale levels, including approaches that are adaptable for a multiscale framework. TTT Project. 
• Multiscale design tools that integrate novel materials, mechanism design, and structural subcomponent design into system level designs. TTT Project. 
• Advancing technology for CMCs and their EBCs for gas turbine engine components operating at 1,482 °C (2,700 °F) or higher. TTT and AATT (Advanced Air Transport Technology) Projects.
• Additive manufacturing and other advanced processing/manufacturing approaches for structural components or materials to enable improved engine efficiency through decreasing weight and/or improving component design, properties, and performance. TTT Project.
• Soft magnetic material with high magnetic saturation and/or lower losses for 100 to 300 kHz operation, hard magnetic materials with an energy product greater than neodymium iron boron, conductors with a specific resistivity less than copper or aluminum, and cable insulation materials with increased dielectric breakdown strength, and significantly higher thermal conductivity (≥1 W/m·K) and resistance to ageing effects such as corona, ozone, humidity, and dust operating at greater than 3 kV. TTT, AATT, CAS (Convergent Aeronautics Solutions) Projects/HGEP (Hybrid Gas-Electric Propulsion) subproject.
• Novel materials systems and structures to enable functionality, such as power harvesting, thermal management, self-sensing, and actuation. Approaches may include use of nanotechnology and novel processing to tailor and control properties such as thermal conductivity, electrical conductivity, thermoelectric response, microstructure and porosity, and shape-memory behavior. TTT Project.
• Design and development of unique materials such as shape-memory alloys and high-entropy alloys for aeronautics structures and components. TTT and CAS Projects.
• Propulsion aeromechanics, damping devices, and analysis and mistuning analysis for turbomachinery rotating blades. TTT Project/AATT - UPAI (Unconventional Propulsion Airframe Integration)

References:

1. Vision 2040: A Roadmap for Integrated, Multiscale Modeling and Simulation of Materials and Systems, NASA/CR-2018-219771, https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20180002010.pdf
 

HGEP:
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170004515.pdf

Hybrid Electric Propulsion:
https://www1.grc.nasa.gov/aeronautics/hep/
https://www.grc.nasa.gov/www/cdtb/aboutus/workshop2016/HybridElectricPropulsionReportOut-LCCPCDRoadmapWorkshop.pdf
 

TTT:
https://www.nasa.gov/aeroresearch/programs/tacp/ttt/description
https://www.nasa.gov/aeroresearch/programs/tacp/ttt
 

CAS:
https://www.nasa.gov/aeroresearch/programs/tacp/cas
https://nari.arc.nasa.gov/
https://www.nasa.gov/aeroresearch/programs/tacp
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170006909.pdf

Lead Center: GRC

Participating Centers: AFRC, LaRC

Scope Title:

Electrified Aircraft Propulsion

Scope Description:Expected TRL or TRL Range at completion of the Project: 2 to 6
Primary Technology Taxonomy:
    Level 1: TX 01 Propulsion Systems
    Level 2: TX 01.3 Aero Propulsion
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype
• Hardware
• Software

Desired Deliverables Description:

Deliverables vary considerably within the topic, but ideally proposals would identify a technology pull area (with a market size estimate), how the proposed idea addresses the needs of the technology pull area and then deliver a combination of analysis and prototypes that substantiate the idea's merit. For Phase I, it is desirable that the proposed innovation clearly demonstrates that it is commercially feasible and addresses NASA's needs. Deliverables for a Phase II should be focused on the maturation, development, and demonstration of the proposed technical innovation.

State of the Art and Critical Gaps:

The critical technical need is for lightweight, high-efficiency power distribution systems and energy storage that have flight-critical reliability. Typically, the weight needs to be reduced by a factor of 2 to 3 and efficiency needs to be improved. Higher efficiency reduces losses and makes thermal management more achievable in an aircraft. Another need for medium to large aircraft is the ability to operate at voltages above 600 V. This capability results in reduced weight, however, is called out specifically because it impacts all of the power system components.

 

Technologies that address these gaps enable Electrified Aircraft Propulsion (EAP) which enables new aircraft configurations and capabilities for the point-to-point on-demand mobility market and a new type of innovation for transport aircraft to reduce fuel consumption and emissions.

Relevance / Science Traceability:

EAP is an area of strong and growing interest in Aeronautics Research Mission Directorate (ARMD). There are emerging vehicle level efforts in Urban On-Demand Mobility, the X-57 electric airplane being built to demonstrate EAP advances applicable to thin and short haul aircraft markets and an ongoing technology development subproject to enable EAP for single aisle aircraft. Additionally, NASA is starting the new Electrified Powertrain Flight Demo (EPFD) project to enable a MW-class aircraft.

 

 

Key outcomes NASA intends to achieve in this area are:

• Outcome for 2015-2025: markets will begin to open for electrified small aircraft.
• Outcome for 2025-2035: certified small-aircraft fleets enabled by electrified aircraft propulsion will provide new mobility options. The decade may also see initial application of electrified aircraft propulsion on large aircraft.
• Outcome for >2035: The prevalence of small-aircraft fleets with electrified propulsion will provide improved economics, performance, safety, and environmental impact, while growth in fleet operations of large aircraft with cleaner, more efficient alternative propulsion systems that will substantially contribute to carbon reduction.

 

Projects working in the vehicle aspects of EAP include:

• Advanced Air Vehicles Program (AAVP)/Advanced Air Transport Technology (AATT) Project
• Integrated Aviation Systems Program (IASP)/Flight Demonstrations and Capabilities (FDC) Project
• AAVP/Revolutionary Vertical Lift Technology (RVLT) Project
• Transformative Aeronautics Concepts Program (TACP)/Convergent Aeronautics Solutions (CAS) Project
• TACP/Transformational Tools and Technologies (TTT) Project

References:

EAP is called out as a key part of Thrust 3 in the ARMD strategic plan: https://www.nasa.gov/aeroresearch/strategy 

Overview of NASA's EAP Research for Large Subsonic Aircraft:  https://ntrs.nasa.gov/search.jsp?R=20170006235

NASA X-57 Project:  https://www.nasa.gov/aeroresearch/X-57/technical/index.html

Lead Center: LaRC

Participating Centers: ARC, GRC

Scope Title:

Computational Tools and Methods

Scope Description:

 

Expected TRL or TRL Range at completion of the Project: 3 to 6
Primary Technology Taxonomy:
    Level 1: TX 15 Flight Vehicle Systems
    Level 2: TX 15.1 Aerosciences
Desired Deliverables of Phase I and Phase II:

• Software
• Research
• Analysis

Desired Deliverables Description:

For focused area 1, the suggested research and development (including deliverable) during Phase I and II include:

Phase I:

• Single-zone demonstrations using airfoil geometry for O-grid or C-grid topologies with precomputed error-metric scalar/vector fields.
• Structured multiblock (without overset) grid for wing geometry.
• Structured overset grids for the wing geometry.
• Demonstrate capability of grid redistribution with up to 200 million grid cells within 20 minutes of wall-clock time.
• Provide an executable or an API for independent assessment by NASA teams.

Phase II:

• Collaborate with NASA team to develop API for coupled RANS-based grid adaptation.
• Assessment on four open-source complex structured grids: (1) Drag Prediction Workshop 6 (Wing-body-nacelle-pylon), (2) nozzle with baseline ramp (5^th AIAA Propulsion Aerodynamics Workshop), (3) Sonic-Boom Prediction Workshop (C608 geometry), and (4) High Lift Prediction Workshop (HLPW-4).
• Demonstrate distributed memory strong scalability on grid sizes up to 1 billion with granularity at 500,000 points per core in under 5 minutes of wall-clock time. 

For focused area 2, the suggested research and development during Phase I and II include:

Phase I:

• Given a bounding surface mesh (tri/quad/poly), demonstrate fully automated body-fitted polyhedral volume mesh generation for canonical geometries as proof of concept.
• Ability to export grid in CFD General Notation System (CGNS) (version greater than 4.1) file format.
• Demonstrate required cell quality metrics:
  • For each face, vector connecting the left and right cell centroids are aligned with face normal.
  • For each face, face centroid is half-way between left and right cell centroids.
• User defined refinement criteria should allow for:
  • Spatially varying wall-distance specified on the surface.
  • Minimum cell-centroid to face-centroid distance criteria specified by the user.
• Provide executable to NASA teams for preliminary testing.

Phase II:

• Demonstrations on three complex topologies: (1) HLPW-4, (2) Benchmark for Airframe Noise Computations (BANC-4) landing gear (PDCC-NLG), and (3) Multistream Chevron nozzle (TMP17) satisfying the requirements established in Phase 1.
• Performance assessment on a HLPW-4 geometry:
  • Demonstrate distributed memory weak scaling up to 10 billion cells at better than 10 million cells per core hour.
  • Demonstrate strong scaling on 10 billion cells mesh up to a granularity of 100,000 cells per core.
• Demonstrate the ability to conform to a user specified set of cell centroids.
• Develop API in collaboration with NASA CFD code developers.
• Demonstrate ability to regrid based on modified refinement criteria accessible through the API.

For focused area 3, demonstrations of the UQ methodologies and tools on a single NASA-FUN3D flutter problem are desired during Phase I. Phase II efforts would require more complete development of an automated and user-friendly toolbox.

State of the Art and Critical Gaps:

NASA's CFD Vision 2030 Study identified several impediments in computational technologies and this solicitation addresses one of those related to application of scale resolving simulations needed for expanding the scope of application of CFD across the aircraft flight envelope, particularly in the prediction of maximum lift.

Relevance / Science Traceability:

Various programs and projects of NASA missions use CFD for advanced aircraft concepts, launch vehicle design, and planetary entry vehicles. The developed technology will enable design decisions by Aeronautics Research Mission Directorate (ARMD) and Human Exploration and Operations Mission Directorate (HEOMD).

References:

https://www.nasa.gov/aeroresearch/programs/aavp

https://www.nasa.gov/aeroresearch/programs/tacp

NASA’s CFD Vision 2030 Study: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140003093.pdf

Lead Center: GRC

Participating Centers: AFRC, ARC, LaRC

Scope Title:

Electric Vertical Takeoff and Landing (eVTOL) Electric Propulsion Safety and Reliability

Scope Description:

The expanding Urban Air Mobility (UAM) vehicle industry has generated a significant level of enthusiasm among aviation designers and manufacturers, resulting in numerous vehicle configurations. The majority of the prototype UAM vehicles have more than 4 rotors or propellers, have electric propulsion, carry 2 to 6 passengers, fly more like a helicopter (vertical takeoff and landing) than a fixed-wing aircraft and will fly relatively close to the ground and near buildings. There are many technical challenges facing industry‘s development of safe, quiet, reliable, affordable, comfortable, and certifiable UAM vehicles and vehicle operations. One of those challenges is the subject of this SBIR subtopic, namely, safety and reliability of the electric power system and electric powertrain for these UAM /eVTOL vehicles.
 

This solicitation is seeking advancements in technologies that will improve the safety and reliability of the electrical power system/powertrain of UAM vehicles. There are three areas of main interest for this solicitation focused on the safety and reliability of the electrical power/powertrain system: (1) prognostic and diagnostic technologies for electric motors, (2) fault protection of high-voltage power system, and (3) advancements in thermal management technologies for electric motor systems. Based on industrial motor research and field data, use of continuous monitoring of conditions, such as temperature and vibration, can reduce the failure rate of motors by about two-thirds [Ref. 1]. This conclusion is based on data from large motors, but the physics of aviation motor degradation may be similar and future service experience may produce a similar overall trend. Due to the power levels envisioned for UAM vehicles, most will require high voltage (>540 V) operation, with the corresponding high-voltage direct current (DC) protection devices to ensure safe systems [Ref. 2]. Advancements in thermal management technologies, noted below, are also of interest given that temperature influences the overall safety and reliability of the power/powertrain system.  
 

The application of the requested technologies should be relevant to the NASA Revolutionary Vertical Lift Technology (RVLT) project’s reference concept vehicles [Refs. 3-4], which embody the key vehicle characteristics of the UAM vehicle configurations being designed throughout industry. Technologies proposed for this solicitation should be relevant to 100 kW class motor-rotor powertrain elements with scalability in the 20 to 500 kW class.
 

Through this solicitation, NASA is seeking advanced technologies supporting UAM electric/hybrid-electric propulsion in the areas of:

• Prognostic/diagnostic technologies: Operational performance diagnostics and prognostics technologies for electric motors and motor controllers to improve life and reliability of the system and reduce unscheduled maintenance for improved affordability. Technologies of interest related to motors/motor controllers include novel sensor and/or sensing technologies, data analysis algorithms, and methods to classify, fuse, or otherwise combine multiple sources of information for diagnostics, prognostics, and remaining useful life assessments. Proposers should consider the anticipated vehicle usage that will require differing motor power requirements for different phases of likely UAM flight profiles and maneuvers (power transients). The proposers should quantify any additional mass and/or power required (for example, mass of any additional sensors and associated cabling). Solutions that add minimal or no extra mass are of higher value and interest. The topic scope for the operational performance diagnostics and prognostics technologies is limited to the powertrain, motor controllers, and associated supporting systems for lubrication and cooling. The scope does not include batteries, fuel cells, fuel burning engines, avionics, communications equipment, or airframe structures.
• High-voltage protection devices: The high voltages expected in these systems will require properly rated protection devices. These do not currently exist for aeronautics-rated applications. While DC and alternating current (AC) architectures are both under consideration by designers, the focus here is on DC systems as these are the first being investigated by NASA. For this call, advanced technology high-voltage protection devices proposed should be applicable for systems with these characteristics: Voltage of 540 to 1,000 V, 100 to 200 A continuous operation with <5 msec clearing time. Other aspects of interest are current-limiting capability, high efficiency, and high reliability.   
• Thermal management technologies: Lightweight, durable, and efficient thermal management solutions for 100 kW continuous-class electric motors are sought. Preferable are solution technologies that could be scaled and applied to aviation motors in a continuous power range of 20 to 500 kW. The most prevalent motor topology proposed and studied for this application are the permanent magnet motor family. Both passive and active approaches for thermal management may be applicable. Examples of technologies sought include integrated channels in rotor components, integrated heat pipes, and solutions enabled by additive manufacturing. Intended outcomes are effective and invasive solutions for motor rotors and windings with minimal mass and efficiency penalty. Solutions that demonstrate an improvement in reliability of the cooling system will be given extra consideration. Proposals should address the continuous motor specific power enabled by the innovations and the associated mass and efficiency penalties.

Expected TRL or TRL Range at completion of the Project: 2 to 4
Primary Technology Taxonomy:
    Level 1: TX 01 Propulsion Systems
    Level 2: TX 01.3 Aero Propulsion
Desired Deliverables of Phase I and Phase II:

• Analysis
• Research
• Prototype

Desired Deliverables Description:

Phase I of the SBIR should develop design concepts for specific technology advancements that address safety and reliability supported by analytical studies including modeling and simulation. Phase I effort should establish Phase II goals and should quantify projections of improvements to safety, reliability, and unscheduled maintenance assuming success of Phase II goals.
 

Phase II of the SBIR should further develop the designs and validate achievement of goals through additional analysis, modeling, and simulation and through system/component functionality experiments. Phase II incorporates experiments with aircraft relevant hardware available commercially or through partnership with an aircraft component supplier and modified with innovative technology from this SBIR effort.

State of the Art and Critical Gaps:

There are over 200 UAM vehicle concepts in varying stages of development. The immediate focus of the vehicle developers is overcoming obstacles on the path to certification. The public has experience flying in large transport aircraft and regional fixed-wing aircraft and are calibrated to associated safety levels for commercial air transportation. Detailed certification requirements for UAM vehicles are still under development by the relevant certifying authorities. For UAM aircraft, research is needed that addresses safety and reliability expectations of the traveling public and certifying authorities.

Relevance / Science Traceability:

This subtopic is relevant to the Aeronautics Research Mission Directorate (ARMD) Revolutionary Vertical Lift Technology (RVLT) Project under the Advanced Air Vehicle Program. The goal of the RVLT Project is to develop and validate tools, technologies, and concepts to overcome key barriers for vertical lift vehicles. The project scope encompasses technologies that address noise, speed, mobility, payload, efficiency, environment, and safety for both conventional and nonconventional vertical lift configurations. This subtopic directly aligns with the mission goals and scope in addressing safety and reliability of nonconventional vertical lift configurations.

 

References:

Lead Center: LaRC

Participating Centers: GRC

Scope Title:

Thermionic Energy Harvetsing

Scope Description:

When employing a nonrotating propulsive device (engine), conventional generators for electrical power are not generally a viable design option; therefore, an alternative option is to utilize direct thermal-to-electric power conversion methods. One such technique is thermionic energy extraction applied to the primary flow path. Thermionic devices have been shown to be efficient in exploiting extreme temperature differentials, like those that typically characterize hypersonic flight environments, and also have a higher/larger energy density than that of standard batteries. The applications of thermionic devices in hypersonic vehicles are numerous, and broaden the design space, since 10 to 100 kW/m² of electrical power are compatible with realistic hardware implementation strategies. Significant, prior efforts in the 1960s, by both the U.S. and Soviet Union, were explored for space applications, but hypersonic atmospheric applications have not been extensively investigated.

 

For hypersonic atmospheric applications, the operational environments, relevant to the thermionic conversion process, are characterized by free-stream flight trajectories having dynamic pressures bounded by 750 to 3,500 (lbf/ft)/ft, and internal-flow path, peak static temperatures associated with adiabatic-flame temperatures (of a combusting hydrocarbon/air stoichiometric mixture). Additionally, volumetric and weight constraints of a novel thermionic device need to be competitive, and/or exceed, existing state-of-the-art battery technology in order to create viable design alternatives for system-design exploration. Although U.S. governmental restrictions typically apply to hypersonic-flight systems, this is not deemed to be a limiting factor in this Phase I activity, due to the low TRL; however, this issue will need to be further addressed during a Phase II award cycle, via coordination with appropriate U.S. governmental and industrial entities.

Expected TRL or TRL Range at completion of the Project: 1 to 3
Primary Technology Taxonomy:
    Level 1: TX 03 Aerospace Power and Energy Storage
    Level 2: TX 03.X Other Aerospace Power and Energy Storage
Desired Deliverables of Phase I and Phase II:

• Research
• Analysis
• Prototype

Desired Deliverables Description:

Phase I - Design of a testable prototype, capable of generating a few W/cm².

 

Phase II - Generation of experimental data in a relevant test environment with a thermionic device to quantify basic performance parameters and initiate scaling study.

State of the Art and Critical Gaps:

The specifics of designing a device with acceptable work functions for both emitter and collector surfaces is a significant technical issue. Additionally, manufacturing a volumetrically efficient device is not codified.

Relevance / Science Traceability:

The applications of power generator in harsh environments is relevant to atmospheric flight and numerous other commercial applications associated with energy conversion needs.

References:

(1) Herring, C. and Nichols, M. H., "Thermionic Emission," Reviews of Modern Physics, Vol. 21, No. 2, 1949.

(2) Rasor, N.S., "Thermionic Energy Converter," Fundamentals Handbook of Electrical and Computer Engineering," pg. 668 ISBN-0471-86213-4 (1983).

(3) Mahefkey, T., "Thermionics Quo Vadis?", National Academy Press (2001).

(4) SECOND INTERNATIONAL CONFERENCE ON THERMIONIC ELECTRICAL POWER GENERATION, Stresa, Italy - May 27-31, 1968.

Lead Center: LaRC

Participating Centers: ARC, GRC

Scope Title:

Boundary Layer Velocity Profile Measurement On Models in Wind Tunnels

Scope Description:

NASA's aeroscience ground test facilities include wind tunnels, air-breathing engine test facilities, and simulation and loads laboratories. They play an integral role in the design, development, evaluation, and analysis of advanced aerospace technologies and vehicles. These facilities provide critical data and fundamental insight required to understand complex phenomena and support the advancement of computational tools for modeling and simulation. The primary objective of the Aeronautics Ground Test and Measurements Technologies subtopic is to develop innovative tools and technologies that can be applied in NASA’s aeroscience ground test facilities to revolutionize testing and measurement capabilities and improve utilization and efficiency. Of primary interest are technologies which can be applied to NASA’s portfolio of large-scale ground test facilities.
 

Spatially and temporally resolved velocity measurements are sought to study boundary layers on test articles in NASA’s large-scale ground test facilities. Proposed measurement capabilities could be used to measure boundary layer thickness, velocity profiles, and transition from laminar to turbulent flow, as well as detecting flow separation that could potentially be used for flow control. Measurement systems should be reliable and robust enough to be implemented in various wind tunnel facilities such as the National Transonic Facility, the 11-Foot Transonic Unitary Plan Facility, the Transonic Dynamics Tunnel, and 8- by 6-Foot Supersonic Wind Tunnel. These facilities operate using a variety of media, including air, nitrogen, and other gases such as R134a.  
 

Technology for measuring boundary layers on models in cryogenic wind tunnels is relatively undeveloped so innovative techniques which can operate at cryogenic conditions are especially of interest. In one potential application in a transonic cryogenic wind tunnel, the boundary layer thickness can be as small as 0.1 mm (e.g., on a wing) but could be as large as 10 mm (e.g., on a fuselage). In this same application, the flow velocity could range from -50 to 200 m/sec.  While light projection and remote detection capabilities will be considered, fiber optic light delivery could be advantageous, particularly if the fiber optic transmission could utilize an existing pressure tap (or multiple taps) in the wind tunnel model. Scattering from light transmitted by a fiber optic embedded in the model could be detected by a remote, high magnification detection system or locally on the model (though that would result in additional cost to modify the wind tunnel model, which is less attractive).
 

Discrete sensor systems could be proposed that both transmit and collect light from a single location. Considerations for selection include the following: the ability to measure at multiple spatial locations, with higher spatial resolution, with higher temporal resolution, over a wider range of conditions, with higher accuracy and higher precision, and probes with potential for measuring multiple parameters (multiple velocity components or temperature, density, or other). Molecular- and particle-based approaches will also be considered. Such a sensor may have medical or industrial applications. Plans for device calibration and risk mitigation should be provided. Expected accuracy, precision, and spatial and temporal resolution should be estimated.

Expected TRL or TRL Range at completion of the Project: 2 to 5
Primary Technology Taxonomy:
    Level 1: TX 13 Ground, Test, and Surface Systems
    Level 2: TX 13.X Other Ground, Test, and Surface Systems
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware

Desired Deliverables Description:

Phase I: The desired deliverable would be a detailed system design for measuring thin boundary layers on model surfaces in both air and cryogenic conditions. This concept or system design should preferably be verified by simulation or testing and capable of being developed into a prototype system as part of a Phase II effort.

 

Phase II: Desired deliverables include a prototype system to measure boundary layers on a representative aerodynamic surface at room and cryogenic temperatures, a robust process for calibrating the system, and well-documented results from functional/operation tests including quantified uncertainty. It is anticipated that the technology would be demonstrated in a NASA ground-based test facility as part of the Phase II effort or shortly thereafter under Phase III funding.

State of the Art and Critical Gaps:

As NASA continues in the quest to advance computational methods and predictive tools as part of CFD 2030, one of the nemeses to significant advancement and validation is the lack of velocity information in the boundary layer. At both low and high speeds, making this measurement requires probe-based systems to be positioned close to the surface or optical-based systems to be focused near the surface. This is quite challenging due to (1) the thickness of the boundary layer, which can be as small as 0.1 mm, (2) conditions on the surface of the test article, (3) vibration of the test article, and (4) drastic temperature changes in the wind tunnel, which can range from room temperature to as low as -200 °C at cryogenic temperatures. There is currently no measurement system available to measure thin boundary layers, especially at cryogenic conditions, and NASA’s inability to obtain this critical data in ground-based facilities at near-flight Reynolds numbers limits the validation of numerical simulations and predictive tools for use in design and certification by analysis. The proposed technology could possibly have application on flight vehicles and in the medical field as well.

Relevance / Science Traceability:

The Transformational Tools and Technologies Project would use this technology to provide critical data to validate computational tools at near-flight Reynolds numbers as it pursues technical challenges to realize the CFD 2030 Vision and enable new aerospace vehicles to be certified by analysis. The technology would also have application for tests supporting the Advanced Air Transport Technology Project, the Commercial Supersonic Technology Project, the Revolutionary Vertical Lift Technology Project, and the Aerosciences Evaluation and Test Capabilities (AETC) Portfolio.

References:

https://www.nasa.gov/aeroresearch/programs/aavp/aetc/ground-facilities
https://ntrs.nasa.gov/search.jsp?R=20140003093

Scope Title:

Measurement Technologies for Vertical Lift Configurations and Concepts

Scope Description:

NASA is currently evaluating new vertical lift configurations and concepts to enable Advanced Air Mobility (AAM). The measurement capabilities needed for AAM research are varied depending on whether model- or full-scale test articles are involved, whether wind tunnel or flight testing is conducted, and whether customized or off-the-shelf hardware is utilized. Nevertheless, vertical lift concepts will require new methods of providing real-time or near real-time measurements of rotor blade and vehicle performance including blade surface pressure distributions, boundary layer transition, rotor performance and blade position, rpm/phasing measurements, rotor control positions (cyclic pitch and flapping) and off-body flow field measurements for examining multiblade vortex interactions and wake/fuselage interaction. Current capabilities for obtaining these data are limited due to lower sampling rates and spatial resolutions using remote sensing. Technologies are needed to (1) increase the temporal and spatial resolutions of remotely sensed/wireless measurements and (2) bring these remotely sensed/wireless measurements onboard the vehicle or test article, where possible, to allow data to be efficiently acquired across a range of operational envelopes. Instrumentation that can be embedded in the blades and fuselage or added to off-the-shelf vehicle or model hardware with efficient data telemetry are desired.

Expected TRL or TRL Range at completion of the Project: 2 to 5
Primary Technology Taxonomy:
    Level 1: TX 13 Ground, Test, and Surface Systems
    Level 2: TX 13.X Other Ground, Test, and Surface Systems
Desired Deliverables of Phase I and Phase II:

• Prototype
• Hardware
• Research

Desired Deliverables Description:

Phase 1: The desired deliverable would be a detailed system design to enable remotely sensed/wireless measurements of blade surface pressure distributions, boundary layer transition, rotor performance and blade position, and off-body measurements. The system should enable efficient acquisition of performance data across the operating envelope of a representative AAM vehicle or model.

 

Phase II: Desired deliverables for Phase II include demonstration of a prototype system on a representative AAM vehicle, model, or test stand of the offeror’s choosing or in collaboration with NASA during a Revolutionary Vertical Lift Technology- (RVLT-) sponsored test.

State of the Art and Critical Gaps:

As part of the AAM mission, NASA is seeking to develop vehicle concepts and technologies to define requirements and standards that address key challenges such as safety, affordability, passenger acceptability, noise, automation, etc. Evaluation of these concepts and technologies will inherently depend on the same type of vehicle performance data typically used to assess the performance of larger-scale vehicles. Unfortunately, the same measurement techniques used for larger-scale vehicles are not directly applicable to AAM vehicles due to the reduced scale and unique design. As such, new measurement capabilities are needed that can be applied or added to vehicles or models during and after development.

Relevance / Science Traceability:

This technology would help fulfill the vision of the Revolutionary Vertical Lift Technology (RVLT) Project to create a future where vertical takeoff and landing (VTOL) configurations operate quietly, safely, efficiently, affordably, and routinely as an integral part of everyday life. This scope has been endorsed by the RVLT Project Manager as well as the Technical Leads in RVLT for Validation Test Campaigns.
 

The system might also have application to other types of flight and ground-based testing within the NASA Enterprise. This would fall under the purview of the Aerosciences Evaluation and Test Capabilities (AETC) Project.

References:

https://www.nasa.gov/aam

https://www.nasa.gov/aeroresearch/programs/aavp/rvlt