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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2015

Formation Flying for Satellites and Unmanned Aerial Vehicles

The shrinking size of satellites and unmanned aerial vehicles (UAVs) is enabling lower cost missions. As sensors and electronics continue to downsize, the next step is multiple vehicles providing different perspectives or variations for more precise measurements. While flying a single satellite or UAV autonomously is a challenge, flying multiple vehicles in a precise formation is even more challenging. The goal of this project is to develop a scalable mesh network between vehicles (satellites or UAVs) to share real-time position data and maintain formations autonomously. Newly available low-cost, commercial off-the-shelf credit card size computers will be used as the basis for this network. Mesh networking techniques will be used to provide redundant links and a flexible network. The Small Projects Rapid Integration and Test Environment Lab will be used to simulate formation flying of satellites. UAVs built by the Aero-M team will be used to demonstrate the formation flying in the West Test Area. The ability to test in flight on NASA-owned UAVs allows this technology to achieve a high Technology Readiness Level (TRL) (TRL-4 for satellites and TRL-7 for UAVs). The low cost of small UAVs and the availability of a large test range (West Test Area) dramatically reduces the expense of testing. The end goal is for this technology to be ready to use on any multiple satellite or UAV mission.

Merrill, Garrick↗
NASA NTRS2017

Point-Focus Concentration Compact Telescoping Array: Extreme Environments Solar Power Base Phase Final Report

Orbital ATK, in partnership with Mark ONeill LLC (MOLLC), has developed a novel solar array platform, PFC-CTA, which provides a significant advance in performance and cost reduction compared to all currently available space solar systems. PFC refers to the Point Focus Concentration of light provided by MOLLCs thin, flat Fresnel optics. These lenses focus light to a point of approximately 100 times the intensity of the ambient light, onto a solar cell of approximately 125th the size of the lens. CTA stands for Compact Telescoping Array, which is the solar array blanket structural platform originally devised by NASA and currently being advanced by Orbital ATK and partners under NASA and AFRL funding to a projected TRL 5+ by late-2018.The NASA Game Changing Development Extreme Environment Solar Power (EESP) Base Phase study has enabled Orbital ATK to refine component designs, perform component level and system performance analyses, and test prototype hardware of the key elements of PFC-CTA, and increased the TRL of PFC-specific technology elements to TRL 4. Key performance metrics currently projected are as follows: Scalability from 5 kW to 300 kW per wing (AM0); Specific Power 500 Wkg (AM0); Stowage Efficiency 100 kWm3; 5:1 margin on pointing tolerance vs. capability; 50 launched cost savings; Wide range of operability between Venus and Saturn by active andor passive thermal management.

McEachen, Michael E.↗
NASA NTRS2017

100 kW Nested Hall Thruster System Development

Large scale cargo transportation to support human missions to the Moon and Mars will require very high power Solar Electric Propulsion (SEP) systems operating between 200 and 400 kW. Aerojet Rocketdyne's NextSTEP program is developing and demonstrating a 100 kW EP system, the XR-100, using a Nested Hall Thruster (NHT) designed for powers up to 200 kW, a modular power processor and a modular flow controller. The three year program objective is to operate the integrated EP system continuously at 100 kW for 100 h, advancing this very high power Electric Propulsion (EP) system to Technology Readiness Level (TRL) 5. With our University of Michigan, Jet Propulsion Laboratory and NASA Glenn Research Center teammates, Aerojet Rocketdyne has completed the initial phase of the program, including operating the thruster at up to 30 kW to validate the thermal models and developing and operating multiple power processor modules in the required seriesparallel configuration. The current phase includes completing a TRL 4 integrated system test at reduced power to validate all system operating phases. Design upgrades to demonstrate the TRL 5 capabilities are underway. This paper will present the high power XR-100 capabilities, overall program and design approach and the latest test results for the 100 kW EP system demonstration program.

Jackson, Jerry↗
NASA NTRS2016

JPL Technology Readiness Assessment Guideline

New capabilities in spaceflight missions are enabled by new technologies. Transitioning new technology to spaceflight elements is difficult and introduces risk, but finding the right balance between benefit and risk leads to scientific advancements and novel space missions. A clear understanding of the risks of new technology can create an environment where innovation is nurtured rather than avoided. The Technology Readiness Level (TRL) was developed as a metric for the maturity of new technology, but, in the past, assessing the TRL was often done informally and inconsistently. This frequently led to discrepancies between the TRL as perceived by the technologist and that perceived by a project. JPL has developed a guideline for their projects to provide a basis for a consistent Technology Readiness Assessment (TRA). Highlights of this guideline are presented here. It is anticipated that the implementation of this guideline will enable the hand-off from technologists to project engineers leading to greater acceptance of technologies by flight projects. On completion of a satisfactory TRA, an agreement can be made between the parties on the maturation plan required for successful infusion of the technology into a flight mission.

Frerking, Margaret A.↗
NASA NTRS2019

Point-Focus Concentration Compact Telescoping Array: EESP Option 1 Phase Final Report for Public Release

Orbital ATK, in partnership with Mark O’Neill LLC (MOLLC) and SolAero Technologies Corp., has developed a novel solar array platform, PFC-CTA, which provides a significant advance in performance and cost reduction compared to all currently available space solar systems. “PFC” refers to the Point Focus Concentration of light provided by MOLLC’s thin, flat Fresnel optics. These lenses focus light to a point of approximately 100 times the intensity of the ambient light, onto a solar cell of approximately 1/25th the size of the lens. “CTA” stands for Compact Telescoping Array1, which is the solar array blanket structural platform originally devised by NASA and currently being advanced by Orbital ATK and partners under NASA and AFRL funding to a projected TRL 5+ by late-2018. The NASA Game Changing Development Extreme Environment Solar Power (EESP) Option 1 Phase study has enabled Orbital ATK to generate and refine component designs, perform component level and system performance analyses, and test prototype hardware of the key elements of PFC-CTA, and increased the TRL of PFC-specific technology elements to TRL ~5. Key performance metrics currently projected are as follows: Scalability from < 5 kW to >300 kW per wing (AM0); Specific Power > 250 W/kg (BoL, AM0); Stowage Efficiency > 60 kW/m3; 5:1 margin on pointing tolerance vs. capability; >50% launched cost savings; Wide range of operability between Venus and Saturn by active and/or passive thermal management.

Mceachen, Michael E.↗
NASA NTRS2010

TPSAS-NF1676L-10746-DND

This paper presents the design, development, and field testing of a high sensitivity ground-based Differential Absorption Lidar (DIAL) system that was developed under the NASA Instrument Incubator Program. The investigation presents a significant advancement towards the development of future CO2 profiling capability as it incorporates key elements of technologies needed for a future development of global CO2 measuring systems including: (1) 2-μm laser technologies that have been developed under a number of NASA programs including the Laser Risk Reduction Program (LRRP) (2) A novel high quantum efficiency (QE), high gain (without excess noise factor), and low noise phototransistor, and (3) Direct detection DIAL system using a large collection area receiver that is insensitive to speckle and coherence length effects from atmospheric turbulence that influences heterodyne detection systems. The objective of the project was a system TRL of 4, and the goal was TRL 5. Development and testing of the laser, new detector, and receiver systems during the project, integration into a complete lidar system into a trailer, field testing of system at West Branch, IA and comparison of the lidar CO2 measurements with in situ sensors advanced the system to a TRL of 5. The system demonstrated high vertical resolution CO2 profiling capability within the boundary layer and column measurements to long ranges. This is the first direct detection demonstration of a 2-micron CO2 DIAL high vertical resolution capability from instrument concept to field demonstration.

Syed Ismail↗
NASA NTRSDate not supplied

Thermal Protection System to Enable Ice Giant Aerocapture Mission for Delivering Both an Orbiter and an In Situ Probe

The Ice Giants have been identified as high priority science destinations in the last Decadal Survey [1] and could benefit from aerocapture as the primary method for orbit insertion [2]. A mass-efficient aerocapture system will enable the delivery of an orbiter along with an atmospheric probe (for in situ measurements to anchor global data collected by the orbiter) and possibly a lander at Triton [3]. Aerocapture could be executed either using low L/D rigid aeroshell with lift modulation (LMA) [4] or using deployable aeroshell using drag modulation (DMA) [3]. Nearly two decades ago, a NASA-funded team performed Neptune-Triton aerocapture studies with a mid-L/D lifting configuration [5] for achieving orbit using LMA. This study showed aerocapture challenges. Due to very high peak entry conditions combined with very high heat-load, a suite of TPS materials was required and this suite was deemed problematic from a qualification perspective, due test facility limitations. In the past 20 years, progress made in GN&C for lift-guided entry missions such as MSL, Orion EFT1, Mars 2020 and the upcoming Artemis missions, and the expertise in blunt body aerodynamics at large scale (~ 5m) has led the EDL community to conclude that aerocapture is a “go do” engineering activity and most technologies are in hand to propose missions with aerocapture [6] [7]. Aerocapture using DMA, currently in development, is an option for Ice Giant Missions. While DMA is simpler in some sense, due to ballistic entry and no need for lift-guided maneuvering, it has challenges and it’s maturity is lower. LMA and DMA both require one or more ablative Thermal Protection System (TPS) materials for the rigid aeroshell element. The ablative TPS needs to be robust and mass efficient due to the high heat loads and size of the rigid aeroshell. Currently, there are capable ablative thermal protection materials, e.g., Heatshield for Extreme Entry Environments Technology (HEEET), 3-D woven Mid-Density Carbon- Phenolic (3MDCP), and PICA (Phenolic-Impregnated Carbon Ablator) that are mature, i.e., at TRL 6 or higher. NASA also invested in Conformal PICA that was matured to TRL 5. Our goal is to evaluate the applicability of high TRL TPS and consider other design options. We first establish bounding aerocapture trajectories for a wide range of arrival conditions and the associated aerothermal environment. Based on the environments, we then determine the predicted TPS mass for the aeroshell [4]. In this presentation, we will outline the process by which we establish bounding aerocapture trajectories for hyperbolic excess velocities ranging from 27 km/s to 35 km/s, which are shown to be a range of velocities that can reduce the trip time from ~14 years to 8 years. The above velocity range translates to ~12 km/s to ~24 km/s arrival velocities at the planetary entry interface [2]. The velocity reduction required to achieve orbit ranges between ~2.5 km/s to 9.5 km/s for both Neptune and Triton. Propulsive insertion alone, due to the amount of fuel required to achieve the required velocity reduction, limits the science returned [2]. We establish the bounding aerocapture trajectories for a low L/D (~ 0.4) configuration for three different ballistic coefficients. The ballistic coefficient range is determined from three different aeroshell diameters of 3m, 4m and 5m and with an entry system mass of 2200 kg. With the above range of design parameters, we then determine conservative/bounding estimates of aerothermal environments by using a combination of CFD simulations and stagnation point heating estimates [7]. This engineering approach allows us to first assess the TPS need vs. TPS capability and determine the applicability of existing TPS. Once an applicable suite of TPS is determined, the TPS thickness and mass are computed. We show that the TPS mass fraction can be as low as 5% to as high as 20%, depending on total trip time reduction and other design parameters for a range of TPS. This is a large range for TPS mass fraction. We show PICA and HEEET can indeed enable aerocapture missions, but the missions incur a mass penalty. TPS mass savings, can be further reduced with the use of conformal PICA. Advancing the development of Conformal PICA to make it robust across the entire aerothermal environment (peak heat-flux, pressure and shear) range will result in TPS mass fractions of < 10% for Ice Giant aerocapture missions such as the Neptune-Triton mission. Aerocapture allows for not only shortening the trip time but enables larger mass to be placed in orbit. Furthermore, probes deployed from orbit will benefit in reduced entry environments allowing for a lower risk TPS implementation as compared to mission designs where the probe is released prior to orbit insertion. One of the challenges for the Ice Giant community is to ensure mission designs that maximize science and allow flexibility in the placement of the entry probe. The traditional approach to release the probe ahead of the orbiter may not optimize returned science. In this presentation, we will make the case for mature TPS such as HEEET and PICA. While these materials can enable aerocapture missions, completing the development of conformal PICA and extending Conformal PICA to be more robust, will have significant impact to TPS mass efficiency and significantly enhance science return for future Gas- and Ice-Giant missions.

E Venkatapathy↗
NASA NTRS2021

NASA / NIST / FAA Technical Interchange Meeting on Computational Materials Approaches for Qualification by Analysis for Aerospace Applications

This report documents the goals, organization and outcomes of a Technical Interchange Meeting (TIM) on Computational Materials Approaches for Qualification by Analysis, co-organized by NASA, NIST and the FAA. The TIM was held at NASA Langley Research Center on January 15-16, 2020. Approximately 60 subject matter experts (SMEs) representing 8 aerospace manufacturers, 7 government organizations and 2 universities participated. Expertise of the SMEs spanned the Technology Readiness Level (TRL) scale from the low-to-mid TRL focus of government laboratories and universities to the high TRL perspective of the regulatory organizations and aerospace manufacturers. During this TIM, the future needs of the government regulators and manufacturers motivated the overall discussion and framed the input given by the participants. Hence, the key objectives of the TIM were to understand existing gaps in model-based, e.g., computational materials, processing and performance predictions for aerospace materials and components and forecast how they can be matured to support material, process and part-level qualification and certification (Q&C). The TIM focused on process-intensive metallic materials technologies, including, but not limited to, additive manufacturing. Participation was roughly evenly divided among the processing and performance tracks, suggesting that both topic areas are generally perceived as being both valuable and requiring additional investment. The output of this TIM may be used by both participating and other organizations, in part, as guidance for future national efforts on maturing computational materials capabilities for use in the Q&C of advanced metallic material systems in aerospace applications.

Edward H Glaessgen↗
NASA NTRSDate not supplied

Next Generation Thermal Protection System for Outer Planet Probes and Orbiter

Saturn Probe and Ice Giant Orbiter along with in-situ Probe Missions continue to be very high priority mis-sions. A result of the advocacy by OPAG and other Analysis Groups, the 3D Woven, Dual-Layer HEEET thermal protection system, mature at TRL 6, has closed the TPS gap for extreme environment missions. A mid density follow on to DL HEEET, developed to meet earth entry requirements, is a Single Layer vari-ant of the 3D woven TPS, which provides a mass effi-cient single layer 3D Mid-Density TPS (3MDCP) that has been baselined as the heatshield for MSR EEV. Continued development of 3MDCP will elevate it to TRL 6 by 2025. Conformal-PICA (C-PICA) devel-opment was pursued to establish a more efficient and robust alternate to PICA, and it is at TRL 4+. Our rationale for the next generation of TPS de-velopment is based on the on the missions needs of the next decade Outer Planet missions that are unique and more demanding than any other destinations. Taking advantage of the recent planned development of sev-eral materials, the next generation of TPS offers a much more mass efficient option for small, medium, and large class Outer Planet missions. Saturn Probes: While mission designers are gen-erally interested in shallow entry to maintain the g-load during entry around 50g, the heat-load for shallow entry can range between (100 kJ/cm2 – 300 kJ/cm2), two orders of magnitude higher than Venus or Sample Return missions. TPS must not only offer protection but must be mass efficient to perform reasonable sci-ence. TPS mass can quickly become 50% or more of the mass of the entire entry system. Recent analysis performed shows Saturn Probe missions could signifi-cantly benefit from the single layer HEEET (SL-HEEET/3MDCP). A DL-HEEET based heatshield mass could be ~ (40% -50%), SL-HEEET can provide additional (30% - 50%) mass savings. At the same time, C-PICA can provide (30% - 50%) mass savings on the backshell. The combined mass savings can be significant enough to carry an additional probe, if de-sired. Ice Giant Aerocapture Missions: Aerocapture mission architectures can provide significant ad-vantage over traditional propulsive insertion missions in multiple ways. 1) Reduced trip time ~ (4- 6) years (30% -40%), 2) Enables placing the orbiter, probe, and lander, all together and 3) Allows for greater science mass (probes and landers) due to mass efficiency. The delivery of a probe from orbit makes it easier and eliminates mission design constraints by HEEET for direct entry [4] of probes and allows for more targeted in-situ science once the Orbiter is able to collect data. In the past 20 years, progress made in GN&C for lift-guided entry missions (MSL, Orion EFT1, Mars 2020) and the expertise in blunt body aerodynamics at large scale (~ 5m) has led the EDL community to conclude that aerocapture is a “go do” engineering activity. An aerocapture mission that will deplete the excess energy of a fast arrival mission will require a mass efficient TPS that can handle extreme heat-load, ~ (100 kJ/cm2 – 500 kJ/cm2). Hence TPS, feasibility as well as mass efficiency requires assessment. Utilizing the recent developments, a comprehensive, bounding analysis was done to establish the potential for SOA (HEEET) system as well as emerging new TPS such as SL-HEEET and C-PICA. In this proposed poster, we will outline the process by which we establish bounding aerocapture trajecto-ries for hyperbolic excess velocities ranging from 27 km/s to 35 km/s, for low L/D (~ 0.4) configurations and determine conservative/bounding estimate of aer-othermal environment by using a combination of CFD simulations and stagnation point heating estimates [7]. This engineering approach allows us to assess the TPS need vs. TPS capability and determine the applicability of existing TPS. Once an applicable suite of TPS is determined, the TPS thickness and mass are comput-ed. We show that the TPS mass fraction can be as low as 5% to as high as 20%, depending on the use of ad-vanced TPS, while HEEET is sufficient but will require 50% of the entry mass.

E Venkatapathy↗
NASA NTRSDate not supplied

Robust and Mass Efficient Thermal Protection Systems for Future Venus Missions

Several international Venus missions, both orbiter, and in-situ probe are in the works after a prolonged absence. Future missions are expected to be long-duration lander missions and balloon missions to investigate the cloud layer. The entry mass for these in-situ missions will be significantly higher, and the aeroshell will be larger, than previous missions. Two cardinal requirements govern the selection and use of thermal protection systems (TPS), namely, robustness to assure mission safety during entry and mass efficiency so that the useful mass for science is maximized. One cannot trade mission safety for mass when it comes to TPS. The robustness of the TPS is a paramount requirement as it is a single point of system failure. At the same time, TPS mass is carried for the entire mission duration prior to entry, and any excess TPS mass is at the cost of science payload. Future missions will benefit enormously from TPS options with mass and performance benefits far beyond advanced carbon-carbon (ACC) that is currently proposed for the DAVINCI mission. NASA STMD, and SMD jointly invested in the development of 3-D woven thermal protection systems in the last decade, and as a result, the Heat-shield for Extreme Entry Environment Technology (HEEET) TPS has been matured to TRL 6. It is ready for future mission use including Venus. The HEEET project focused on developing a broad technology base, applicable for missions not only to Venus, but also to Saturn, the Ice Giants, and higher speed sample return missions such as Mars Sample Return requiring extreme robustness due to bio-hazards. The dual-layer HEEET (DL-HEEET) is proven to be extremely robust. DL-HEEET performed well in arc jet and laser testing conditions where heritage carbon-phenolic failed. HEEET also proved to be more mass efficient compared to Carbon-Phenolic. Recently, the principal author was invited to present his perspectives to the 2023-2032 Planetary Sciences Decadal Venus subcommittee on the current state of TPS for future Venus missions [1]. In addition to pointing out successful TPS and other developments in the last decade in support of entry systems, the principal author made two findings: 1) importance of and need to sustain TPS capabilities that took nearly a decade to develop and 2) the opportunity to further optimize TPS mass without sacrificing robustness to further enable future in-situ missions. This talk will focus on the above two recommendations and provide the rationale for them. The intent of the talk is to seek advocacy from the VEXAG community for the criticality of TPS sustainment and the benefits of 3D woven TPS optimization. TPS Capability Sustainment: HEEET development was necessitated by the atrophy of heritage carbon-phenolic. Atrophy has impacted other TPS materials as well. Avcoat, the TPS that successfully allowed astronauts to explore the Moon in the 1960s and 70s, took nearly a decade and $10’sM to recover in support of Orion/Artemis. One of the most used TPS materials, SLA 561V, had to be recovered prior to its use on Mars Pathfinder. Phenolic Impregnated Carbon Ablator (PICA), the TPS that replaced SLA 561V on MSL, has undergone multiple replacements and recovery due to constituent rayon changes. Recently, NASA invested in a domestic rayon replacement program and also invested in FMI to consolidate PICA capability for NASA missions, as a result of FMI’s decision to discontinue commercial FiberForm, which is needed for PICA. TPS used for planetary missions are unique and have no other commercial or defense use. In addition, low mission cadence is also a driver for TPS atrophy. Hence, NASA, as the steward, must take steps to ensure and sustain TPS capabilities. In this talk, we outline steps NASA can take to keep abreast of emerging risks and target risk mitigation steps to ensure TPS capability sustainment for Venus and other extreme environment missions. Next Generation of Mass efficient and Robust TPS: NASA invested in an alternate TPS to PICA based on felt-technology called Conformal-PICA which has the potential to save 30% - 50% mass over PICA. The development was discontinued at a Technology Readiness Level (TRL) of ~ 5. 3D Mid-Density Carbon Phenolic (3MDCP) is a single layer variant of HEEET, based on the insulating layer only (SL-HEEET) and currently baselined for the Mars Sample Return Mission Earth Entry System due to its mass efficiency (30% more mass efficient than DL- HEEET). Currently, SL-HEEET is limited to aeroshell diameters of < 1.3m. The SL-HEEET was compared to DL-HEEET in the recent ADVENT flag-ship class mission study in support of the Planetary Science Decadal. SL HEEET was the recommended TPS based on 30% mass savings for both the balloon and lander missions. Given C-PICA and SL-HEEET have superior reliability and mass efficiency, advocacy from VEXAG is sought for completing their further development to TRL 5/6 in this decade so as to enable TPS readiness for future missions.

Thermal Protection Systems↗
NASA NTRSDate not supplied

Development of a Thermal Vacuum Chamber for Lunar Surface Simulation

Simulating the lunar surface environment is a critical component in advancing the Technology Readiness Level (TRL) of newly developed lunar surface exploration systems and understanding regolith dynamics. Developing systems with the ability to simulate the lunar surface environment presents unique challenges associated with both the hardware requirements of the system and the uniqueness of the lunar environment. Additionally, many thermal vacuum testing systems are large and expensive to maintain and operate, increasing the cost for the end users. Many low TRL prototypes are small-scale and low-cost, making it unfeasible and unnecessary to utilize larger scale testing facilities. A small-scale, modular, and easily reconfigurable chamber with a quick testing turnaround time could allow rapid TRL advancement with minimal costs and time commitments relative to many of the current testing facilities. Here, we introduce a thermal vacuum chamber system that will support experimentation and rapid testing and maturation of small-scale hardware systems and subsystems for lunar surface exploration.

Thermal Vacuum Chamber↗
NASA NTRS2024

Hybrid Thermally Efficient Core (HyTEC) HyTEC Phase 1 – Combustor Final Report

The objective of the HyTEC – Combustor Technology project is to develop technology for a compact, low emissions, rich-burn combustor that maintains a high-level of durability. To accomplish this, the combustor will incorporate CMC liners and a CMC Dome to enable improved air utilization for mixing and NOx reduction. In particular, three technology areas were matured: metallic swirler attachment to the CMC dome, Nickel diffusion in CMC, and EBC advancements with improved durability. The maturation of these technology areas is critical to successfully designing a combustor for an engine intended to demonstrate increased thermal efficiency with integrated high-power density-core engine technologies. As part of the project, multiple swirler attachment architectures were matured through the design process, manufactured, and tested in simulated engine conditions. These architectures utilized two different approaches to attachment: a pressure loaded approach and a clamped approach. Each approach required testing at simulated engine conditions to mature the technology. These tests were developed to ensure that the probable failure mechanisms for each architecture were assessed relative to its long-term durability. After successfully testing both approaches, the pressure loaded approach was chosen based on its simplicity and its similarity to the attachment approach used on lean burn combustors. Finally, this successful testing resulted in the swirler attachment technology achieving TRL 4 status. Another aspect of concern with the metallic swirler attachment is the potential for nickel diffusion into the CMC to cause durability shortfalls. A series of tests were completed to understand both the amount of nickel that is expected to diffuse into the CMC and the effect this diffusion will have on mechanical properties of the CMC. After successfully completing this testing, the reduction in local mechanical properties was found to not have an impact on the overall durability of the CMC Dome. To further improve the overall durability, three different EBC coating architectures were also studied. One of these coatings being the current state of the art for CMC liners with the other two being proposed improvements. Coupons were produced using each coating and subjected to a series of laboratory testing. One coating was found to have an overall superior durability in laboratory testing and survived rig testing without indication of degradation. This successful testing resulted in the EBC technology achieving TRL 5 status. Overall, a design approach was successfully down-selected that achieves both the TPM requirements and the need for TRL 4+ status. This design incorporates a pressure loaded swirler attachment with an improved EBC coating (Architecture 1). This combination allows the continued use of a CMC dome and results in an improved durability for the proposed compact core combustor.

CMC↗
NASA NTRSDate not supplied

ISRU Pilot Excavator (IPEx) Technology Readiness Level 5 Design Overview

This paper details the mechanical and mechatronic design of the Technology Readiness Level (TRL)-5 In-Situ Resource Utilization (ISRU) Pilot Excavator (IPEx). IPEx is a robotic excavator designed for a technology demonstration of regolith mining in the lunar south -pole region. The novel design uses pairs of counter-acting excavation tools called bucket drums, that dig at the same time in opposing directions to reduce the reaction force needed, thereby enabling mining with a small, low-mass, robotic system. IPEx builds on the prior work of the Regolith Advanced Surface Systems Operations Robot (RASSOR), which is the TRL-4 implementation of this concept. The TRL-5 IPEx subsystems that are discussed in this paper include: Regolith Delivery Subsystem (RDS), Mobility Subsystem (MS), Cameras and Dust Mitigation Subsystem (CDMS), and Thermal Control Subsystem (TCS). Each subsystem is described in detail with rationale for design selections. Dust tolerance is a key feature for IPEx and this paper details a thermal control system with an actuated radiator cover and phase change material as well as camera modules with removable electrodynamic dust shields (EDS). Additional components such as actuators, wheels, and bucket drums are discussed in detail. Due to their complexity, the avionics and software subsystems will be discussed in a separate publication.

RASSOR↗
NASA NTRS2026

NASA Hybrid Thermally Efficient Core (HyTEC) Phase 2 Q3’GFY25 to Q2’GFY26 Unlimited Rights Annual Report for CLIN 1 and 2

The HyTEC Phase 2 Project focuses on the development and demonstration of high-power density, small engine core gas turbine engine technologies. The focus of HyTEC on these small engine core technologies will provide direct benefits to the next single-aisle class aircraft in terms of thermal efficiency, as well as integrate with other technologies, through increased hybridization, as they mature to provide substantial continuous fuel burn reductions during the aircraft lifecycle. These advances will strengthen the US position in the commercial aviation engine market and enable cost benefits in commercial aviation. In Government fiscal year 26 the HyTEC Project was closed and contract moved to the Subsonic Vehicle Technologies and Tools Project. GE Aerospace is executing two major work plans as part of the HyTEC Phase 2 contract. The first is Contract Line-Item Number (CLIN) 001 and its objective is to mature High Pressure Turbine (HPT) aerodynamics technology to TRL 5 through multiple rigs so that it may be incorporated into a complete TRL 6 evaluation and unlock learnings for the engine core demonstration as applicable. CLIN 002, the second work plan, is to design, procure and test a cost shared portion of technologies on the TRL6 Compact Core test vehicle. The compact core demonstrated through HyTEC Phase 2 along with the HPT rig TRL advancement directly matures the technologies being developed as part of the CFM RISE demonstrator program which is targeting a more than 20% fuel burn reduction at the engine level versus today’s state of the art. The RISE program’s objective is to advance both the novel Open Fan architecture, compact core technologies and advanced systems such as hybrid electric needed to achieve the desired fuel burn reduction to TRL6 ahead of a new product introduction in the single-aisle class aircraft in the 2030s timeframe.

gas turbine engine↗
DOE OSTI2025

Large Area Airborne Contamination Monitoring

The individual components of a Large Area Airborne Contamination Monitoring (LAACM) system have been demonstrated as functional and the concept of a low-cost, rapidly deployable, and rapidly expandable system has been validated. All components of a previously demonstrated system were re-evaluated. Based on this evaluation, recommendations were made for air collection equipment, scintillating material was synthesized and optimized for application, and improved images were acquired. A preliminary concept of operations was developed, outlining the need for an automated scintillation application system and software for data acquisition and analysis. The work conducted under this program in FY24 has raised the Technology Readiness Level (TRL) of the LAACM from TRL 3 to TRL 4.

Whiteside, Tad S. [Savannah River National Laborat↗
DOE OSTI2026

Separation of Hydrogen Using Pd/Ag Membranes: Experimental and Modeling Results with Potential Application to Direct Internal Recycle

Implementation of fusion energy requires processing the deuterium-tritium (D-T) mixture used to fuel the reaction, and separation of hydrogen isotopes from other gases is imperative. Specifically, the separation of hydrogen isotopes from helium is a matter of importance to the fusion fuel cycle community. Initial testing with a palladium-silver (Pd-Ag) membrane indicates that even moderate vacuum (~100 torr permeate pressure) can provide a high degree of separation (>90%) at a high ratio of H 2 to He. Given the presence of He in many fusion systems, a high technology readiness level (TRL) for Q 2 /He (where Q represents any isotope of hydrogen) separations is needed. This study demonstrates the efficacy of H 2 removal from He via permeation and potential applications for direct internal recycle. Modeling will accompany the experimental campaign to generate a predictive capability and quantify the separation performance. Modeling from previous hydrogen permeation studies has demonstrated that the typical Sieverts’ law fails to predict the measured permeation rates at high hydrogen fluxes. Existing models are being refined to integrate the effects of surface phenomena into permeation predictions, which have been expanded to account for mixtures with large ranges of Q 2 concentrations. These data will improve the TRL of permeators as a separation technology for the fusion fuel cycle.

08 HYDROGEN↗
DOE OSTI2024

DISH-STARS™ Commercialization (Abstract)

The goal of this project is to aggressively support the near-term commercialization of a new technology platform – based on the integration of solar concentrators and micro- and meso-channel process technology (MMPT) – that was evaluated and identified as a strong candidate for near-term commercialization at EERE’s inaugural Lab-Corps program during early FY2016. Known as STARS, for Solar Thermochemical Advanced Reactor System, or Dish-STARS™ when paired with parabolic dish concentrators, STARS is a promising energy-related technology developed at the Pacific Northwest National Laboratory (PNNL) that efficiently converts solar energy into chemical energy. Combined with economies through hardware mass production, the efficiency of Dish-STARS™ provides a near-term opportunity for the production of renewable electricity, fuels and chemicals. This proposed CRADA project supports the commercialization of Dish-STARS™ in these important ways: The project will support the cooperative development of Dish-STARS™ by the DOE national laboratory and private partners, including the startup company, STARS Corporation, that is being established by the PNNL Lab-Corps team that evaluated STARS on behalf of EERE. The project will provide important transition funding at the time that the previous DOE SunShot project, which has supported Dish-STARS™ development from Technology Readiness Level 3 (TRL 3) to TRL 6, is scheduled to end.

14 SOLAR ENERGY↗
DOE OSTI2024

Achieving Unprecedented CO 2 Utilization InCO 2 Concrete™: System Design, Product Development and Process Demonstration

Anthropogenic sources of carbon dioxide are generated from a number of sources, but the key among these are ordinary Portland cement (OPC) production and combustion of fossil fuels. Cement production is the largest global CO 2 source from the mineral decomposition of carbonates. This is due to the clinkering process whereby limestone (mainly consisting of CaCO 3 ) is decomposed into CaO and CO 2 , and combined with silica rich clays at high temperatures to form clinkers (i.e. the four key minerals that comprise cement). The high temperature range of 1400 – 1550°C required for this process accounts for up to 60% of the generated CO 2 from cement production. Combination of the limestone decomposition and thermal requirements of the clinkering process causes cement production to contribute 8-9% of annual global CO 2 emissions. Combustion of fossil fuels (coal, oil and gas) was shown to contribute a much larger portion of global CO 2 emissions. As of 2018, combustion of fossil fuels accounted for 65% of global CO 2 , where 41% was derived from stationary sources for electricity and heat generation and the other 24% was related to transport. To reduce these contributions, key steps forward in CO 2 utilization technologies are required. Therefore, a CO 2 mineralization technology (CO 2 mineralization concrete) to reduce the OPC content in concrete, while utilizing flue gas emissions from fossil fuel combustion has been developed to address both areas simultaneously. This Reversa™ technology utilizes low-carbon cementation agents produced by in situ CO 2 mineralization (“mineral carbonation reactions”) to offer a promising alternative to OPC. CO 2 mineralization relies upon the reaction of dissolved CO 2 with inorganic alkaline reactants to precipitate mineral carbonates (e.g., CaCO 3 ), which bind proximate particles and achieve cementation. Herein, a concrete green body, which is composed of a mixture of binder, water, and mineral aggregates, is exposed to CO 2 borne in industrial flue gas streams. This manner of CO 2 mineralization allows the production of construction components that feature equivalent engineering attributes as their OPC-based counterparts while featuring a much smaller embodied carbon intensity (eCI). The purpose of this project is to demonstrate the feasibility of the Reversa process evolving from a TRL-3 technology at the bench-scale up to TRL-6 technology at the pilot-scale. The reliability of the Reversa technology was tested to prove the effective production of three standard industrial concrete products selected during the course of the project. The results detailed herein will demonstrate the evolution of this technology to the industrial scale. The culmination of this work resulted in 9 production runs completed at the National Carbon Capture Center (NCCC), Wilsonville, AL, using natural gas (NG) flue gas as the CO 2 source. Over the course of the production runs at NCCC, the CO 2 utilization as a function of time, 24-h CO 2 uptake, electricity usage, and 28-d net area compressive strength recorded for each run. Collection of this data will be used to determine the success of the demonstration goals: (1) achieving in excess of 0.2gCO 2 /g reactant , (2) achieving greater than 50% reduction in global warming potential compared to standard produced units, and (3) ensuring compliance of carbonated concrete with industry standard specifications.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗