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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Structural Testing of a 6m Hypersonic Inflatable Aerodynamic Decelerator System

NASA is developing low ballistic coefficient technologies to support the Nations long-term goal of landing humans on Mars. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current and future launch vehicle fairing limitations. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) are being developed to circumvent this limitation and are now considered a leading technology to enable landing of heavy payloads on Mars. At the beginning of 2014, a 6m diameter HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify its structural performance under flight-relevant loads. The inflatable structure was constructed into a 60 degree sphere-cone configuration using nine inflatable torus segments composed of fiber-reinforced thin films. The inflatable tori were joined together using adhesives and high-strength textile woven structural straps. These straps help distribute the load throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials that are designed to protect the inflatable structure from heat loads that would be seen in flight during atmospheric entry. A custom test fixture was constructed to perform the static load test series. The fixture consisted of a round structural tub with enough height and width to allow for displacement of the HIAD test article as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The rigid centerbody of the HIAD was mounted to a pedestal in the center of the structural tub. Using an impermeable membrane draped over the HIAD test article, an airtight seal was created with the top rim of the static load tub. This seal allowed partial vacuum to be pulled beneath the HIAD resulting in a uniform static pressure load applied to the outer surface. Using this technique, the test article was subjected to loads of up to 50,000lbs. During the test series an extensive amount of instrumentation was used to provide a rich data set, including deflected shape, structural strap loads, torus cord loads, inflation pressures, and applied static load. In this paper the 2014 6m HIAD static load test series will be discussed in detail, including the design of the 6m HIAD test article, the test setup, and test execution. Analysis results will be described supporting the conclusions that were drawn from the test series..

Hypersonic↗
NASA NTRS2012

Flowfield Analysis of a Small Entry Probe (SPRITE) Tested in an Arc Jet

A novel concept of small size (diameter less than 15 inches) entry probes named SPRITE (Small Probe Re-entry Investigation for TPS Engineering) has been developed at NASA Ames Research Center (ARC). These flight probes have on-board data acquisition systems that have also been developed in parallel at NASA ARC by Greg Swanson1. Flight probes of this size facilitate testing over a wide range of conditions in arc jets available at NASA ARC, thereby fulfilling a 'test what you fly' paradigm. As indicated by the acronym, these probes, with suitably tailored trajectories, are primarily meant to be robotic flight test beds for TPS materials, although the design is flexible enough to accommodate additional objectives of flight-testing other vehicle subsystems. A first step towards establishing the feasibility of the SPRITE concept is to arc-jet test fully instrumented models at flight scale. In a follow-on to the Large-Scale Article Tests (LSAT2) performed in the 60 MW Interaction Heating Facility (IHF) in late 2008/early 2009, a full-scale model of Deep Space-2 (DS23) made of red oak was tested in the 20 MW Aerodynamic Heating Facility (AHF). There were no issues with mass capture by the diffuser for blunt bodies of roughly 15 inches diameter tested in the 18-inch nozzle of the AHF. Building on this initial success, two identical test articles - SPRITE-T1-1 and SPRITE-T1-2 (T1 indicating the choice of back shell geometry) - were fabricated, and one of them, SPRITE-T1-1, was tested in the AHF recently. Both these test articles, 14 inches in diameter, have a 45deg sphere-cone (like DS2) made of PICA bonded on to a 1/8th inch thick aluminum shell using RTV. The aft portion of the test article is a conical frustum (15deg cone angle) with LI-2200 bonded on to the aluminum shell. Each model is fully instrumented with: (a) thermocouples imbedded in plugs in the heat shield, (b) thermocouples bonded to the aluminum substructure; the thermocouples are distributed over the entire shell, and (c) a few strain gages. Data from some of the thermocouples and gages are acquired by the on-board data acquisition system (DAS), while data from the others are routed to the facility-provided DAS, thereby enabling a cross check on the in situ measurement capability. as inputs to v2.6.1 of the in-house materials thermal response code, FIAT

atmospheric entry↗
NASA NTRS2015

Flight Tests of a Supersonic Natural Laminar Flow Airfoil

A flight-test campaign of a supersonic natural laminar flow airfoil has been recently completed. The test surface was an 80-inch (203 cm) chord and 40-inch (102 cm) span article mounted on the centerline store location of an F-15B airplane (McDonnell Douglas Corporation, now The Boeing Company, Chicago, Illinois). The test article was designed with a leading edge sweep of effectively 0 deg to minimize boundary layer crossflow. The test article surface was coated with an insulating material to avoid significant heat transfer to and from the test article structure to maintain a quasi-adiabatic wall. An aircraft-mounted infrared camera system was used to determine boundary layer transition and the extent of laminar flow. The tests were flown up to Mach 2.0 and chord Reynolds numbers in excess of 30 million. The objectives of the tests were to determine the extent of laminar flow at high Reynolds numbers and to determine the sensitivity of the flow to disturbances. Both discrete (trip dots) and 2-D disturbances (forward-facing steps) were tested. A series of oblique shocks, of yet unknown origin, appeared on the surface, which generated sufficient crossflow to affect transition. Despite the unwanted crossflow, the airfoil performed well. The results indicate the sensitivity of the flow to the disturbances, which can translate into manufacturing tolerances, were similar to that of subsonic natural laminar flow wings.

supersonic boundary layer↗
NASA NTRS2016

T-Cap Pull-Off and Bending Behavior for Stitched Structure

The Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) is a structural concept that was developed by The Boeing Company to address the complex structural design aspects associated with a pressurized hybrid wing body aircraft configuration. An important design feature required for assembly is the integrally stitched T-cap, which provides connectivity of the corner (orthogonal) joint between adjacent panels. A series of tests were conducted on T-cap test articles, with and without a rod stiffener penetrating the T-cap web, under tension (pull-off) and bending loads. Three designs were tested, including the baseline design used in largescale test articles. The baseline had only the manufacturing stitch row adjacent to the fillet at the base of the T-cap web. Two new designs added stitching rows to the T-cap web at either 0.5- or 1.0-inch spacing along the height of the web. Testing was conducted at NASA Langley Research Center to determine the behavior of the T-cap region resulting from the applied loading. Results show that stitching arrests the initial delamination failures so that the maximum strength capability exceeds the load at which the initial delaminations develop. However, it was seen that the added web stitching had very little effect on the initial delamination failure load, but actually decreased the initial delamination failure load for tension loading of test articles without a stiffener passing through the web. Additionally, the added web stitching only increased the maximum load capability by between 1% and 12.5%. The presence of the stiffener, however, did increase the initial and maximum loads for both tension and bending loading as compared to the stringerless baseline design. Based on the results of the few samples tested, the additional stitching in the T-cap web showed little advantage over the baseline design in terms of structural failure at the T-cap web/skin junction for the current test articles.

Lovejoy, Andrew E.↗
NASA NTRS2016

Tension and Bending Testing of an Integral T-Cap for Stitched Composite Airframe Joints

The Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) is a structural concept that was developed by The Boeing Company to address the complex structural design aspects associated with a pressurized hybrid wing body aircraft configuration. An important design feature required for assembly is the integrally stitched T-cap, which provides connectivity of the corner (orthogonal) joint between adjacent panels. A series of tests were conducted on T-cap test articles, with and without a rod stiffener penetrating the T-cap web, under tension (pull-off) and bending loads. Three designs were tested, including the baseline design used in large-scale test articles. The baseline had only the manufacturing stitch row adjacent to the fillet at the base of the T-cap web. Two new designs added stitching rows to the T-cap web at either 0.5- or 1.0-inch spacing along the height of the web. Testing was conducted at NASA Langley Research Center to determine the behavior of the T-cap region resulting from the applied loading. Results show that stitching arrests the initial delamination failures so that the maximum strength capability exceeds the load at which the initial delaminations develop. However, it was seen that the added web stitching had very little effect on the initial delamination failure load, but actually decreased the initial delamination failure load for tension loading of test articles without a stiffener passing through the web. Additionally, the added web stitching only increased the maximum load capability by between 1% and 12.5%. The presence of the stiffener, however, did increase the initial and maximum loads for both tension and bending loading as compared to the stringerless baseline design. Based on the results of the few samples tested, the additional stitching in the T-cap web showed little advantage over the baseline design in terms of structural failure at the T-cap web/skin junction for the current test articles.

Lovejoy, Andrew E.↗
NASA NTRS2016

Design and Fabrication of Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged Combustion Thrust Chamber Injectors

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. On the current project, several configurations of new main injectors were considered for the thrust chamber assembly of the integrated test article. All the injector elements were of the gas-centered swirl coaxial type, similar to those used on the Russian oxidizer-rich staged-combustion rocket engines. In such elements, oxidizer-rich combustion products from the preburner/turbine exhaust flow through a straight tube, and fuel exiting from the combustion chamber and nozzle regenerative cooling circuits is injected near the exit of the oxidizer tube through tangentially oriented orifices that impart a swirl motion such that the fuel flows along the wall of the oxidizer tube in a thin film. In some elements there is an orifice at the inlet to the oxidizer tube, and in some elements there is a sleeve or "shield" inside the oxidizer tube where the fuel enters. In the current project, several variations of element geometries were created, including element size (i.e., number of elements or pattern density), the distance from the exit of the sleeve to the injector face, the width of the gap between the oxidizer tube inner wall and the outer wall of the sleeve, and excluding the sleeve entirely. This paper discusses the design rationale for each of these element variations, including hydraulic, structural, thermal, combustion performance, and combustion stability considerations. This paper also discusses the fabrication and assembly of the injector components, including the injector body/interpropellant plate, the additive manufactured GRCop-84 faceplate, and the pieces that make up the injector elements including the oxidizer tube, an inlet to the oxidizer tube, and a facenut that includes the fuel tangential inlets and forms the initial recessed volume where oxidizer and fuel first interact. Hot-fire test results of these main injector designs in an integrated test article that includes an oxidizer-rich preburner are described in companion papers at this JANNAF meeting.

Garcia, C. P.↗
NASA NTRS2017

NASA's Space Launch System: Progress Report

NASA and its commercial industry team achieved significant progress in 2016 in manufacturing and testing of the Block 1 vehicle for the first launch of the Space Launch System (SLS). Test and flight article hardware for the liquid hydrogen fuel tank as well as the engine section for the core stage were completed at Michoud Assembly Facility (MAF) in New Orleans. Test stands neared completion at Marshall Space Flight Center for the propellant tanks, engine section, intertank and payload section. Stennis Space Center completed major structural renovations on the B2 test stand, where the core stage "green run" test program will be conducted. The SLS team completed a hotfire test series at Stennis to successfully demonstrate the ability of the RS-25 engine to operate under SLS environments and performance conditions. The team also test fired the second qualification five-segment solid rocket motor and cast the first six motor segments for the first SLS mission. The Interim Cryogenic Propulsion Stage (ICPS) test article was delivered to Marshall for structural tests, and work is nearly finished on the flight stage. Flight software testing completed at Marshall included power quality and command and data handling. In 2017, that work continues. SLS completed Preliminary Design Review (PDR) on the Exploration Upper Stage (EUS), a powerful, human-rated spacecraft that will propel explorers to cis-lunar space. In 2017, hardware will continue to be integrated at MAF for core stage structural test articles and the first two operational flights. RS-25 hotfire testing will continue to explore engine performance, as well as test flight-like software and four new Engine Controller Units (ECUs) for the first mission. Production of development components for a more affordable RS-25 design is underway. Core stage structural test articles have begun arriving at Marshall. While engineering challenges typical of a new development are possible, SLS is working toward launch readiness in late 2018. This paper will discuss these and other technical and programmatic successes and challenges over the past year and provide a preview of work ahead before first flight

Cook, Jerry↗
NASA NTRS2018

Performance, Stability and Compatibility of Oxygen/RP-1 Multi-Element Oxidizer-Rich Staged-Combustion Injectors

In 2015 and 2016, the National Aeronautics and Space Administration Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/RP-1 propellant multi-element oxidizer-rich staged-combustion test article. The main objective was to provide thrust chamber combustion stability data as part of the Combustion Stability Tool Development program, although demonstration of performance and compatibility of oxidizer-rich main injectors was also important. Funding was provided by the Air Force Space and Missile Systems Center. Five configurations of main injectors were designed and fabricated, using conventional gas-centered swirl coaxial injector element designs generally similar to those used in oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines. Variations of element features included element size, recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. Ablative combustion chambers were fabricated based on hardware previously used at the NASA MSFC for testing at similar size and pressure. Existing oxygen/RP-1 oxidizer-rich subscale preburner injectors and hot gas ducts from a previous NASA-funded program were modified for use to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Testing of the resulting integrated test article - which included the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was conducted at Test Stand 116 at the East Test Area of the NASA MSFC. The test article was well instrumented with static and dynamic pressure, temperature, and vibration sensors. This paper presents and discusses all the hot-fire test results of the integrated test article thrust chamber. Eighteen successful hot-fire tests of the integrated rig were conducted. Testing was accomplished with all five of the injector element concepts. Main combustion chamber pressures ranged from 710 to 2350 psia, and main combustion chamber mixture ratios ranged from 2.47 to 2.87. A chamber barrier fuel film coolant of about 2% to 4% of the total fuel flow was used for most tests. Characteristic exhaust velocity efficiency excluding the influence of the fuel film cooling ranged from 91% to 98% of theoretical. All tests of the thrust chamber exhibited stable combustion, even down to 40% of nominal operating pressures. Compatibility of the injector face and combustion chamber walls was acceptable. This paper is a follow-on to publication of preliminary test data presented at the 2016 JANNAF Liquid Propulsion Subcommittee meeting.

Hulka, J. R.↗
NASA NTRSDate not supplied

TPSAS-NF1676L-31522-DND

Currently, there are limitations in launch capability regarding requisite payload volume and mass for long duration extra-terrestrial missions. In-space additive manufacturing may provide a solution to this challenge and is applicable for articles ranging from large system and subsystem architectures to small consumables and replacement articles. For small articles, additive manufacturing is becoming more feasible through advances in instrumentation and novel manufacturing strategies capable of generating complex shapes. Many approaches, however, utilize feedstock materials that, once consumed, must be replenished before additional architectures can be generated. Thus, although promising, a major challenge with additive manufacturing is the required starting materials to support on-demand article generation. This presentation describes research toward reversibly assembling materials that can revert to the starting material state for subsequent fabrication. Polyimides were synthesized with thermally reversible Diels-Alder reaction functionalities, furan or maleimide moieties. These materials were characterized utilizing typical spectroscopic, thermal, and mechanical techniques. Amine hardener-enriched epoxy microparticles were synthesized for use as a mechanically robust substrate. Integration of these technologies into a single system and application toward generation of secondary structures on a launch vehicle as a means of in-situ resource utilization will be described.

Christopher J. Wohl↗
NASA NTRSDate not supplied

Dynamic Characterization of NASA Armstrong Test Facility Mechanical Vibration Facility

In preparation for Sierra Nevada Corporation’s (SNC) Dream Chaser spacecraft vibration test campaign at NASA Armstrong Test Facility (ATF) in Sandusky, Ohio, a dynamic characterization of the Mechanical Vibration Facility (MVF) is currently ongoing. The Mechanical Vibration Facility is comprised of an 18 ft diameter annulus table that is driven with sixteen hydraulic vertical actuator assemblies and four hydraulic horizontal actuator assemblies. During a test campaign this allows for single axis vibration testing in the vertical axis and in each of the two orthogonal horizontal axes without the need for reconfiguring the test article. The dynamic characterization of the facility was defined to follow a building-block approach requiring three vibration test configurations: (1) MVF bare table (2) MVF table and head expander (3) MVF table, head expander, and a dummy test article. Following the completion of the facility dynamic characterization, a facility finite element model (FEM) will be test-verified allowing for more accurate pretest analysis used for determining response limits and abort levels. To date two of the three test configurations have been completed. Test results have been used to verify the facility FEM that includes both the 18 ft diameter annulus table and the aluminum head expander that fills in the center opening of the annulus to provide a continuous flat mounting surface. Both the MVF table and the head expander models were previously correlated using test results from free-free boundary condition modal tests so model updates to the facility FEM focused on the uncorrelated vertical actuator assemblies and horizontal actuator assemblies. The accuracy of the test-verified facility FEM will be further determined based on comparison of analytical and test results from the vibration test of the dummy test article. Efforts will also be made to determine fixed base modes of the test article attached to the flexible vibration table using fixed base correction techniques. This paper will discuss test and analytical results from the MVF dynamic characterization test sequence and corresponding model updating effort.

Emma L Pierson↗
NASA NTRS2022

Friction Stir Welded Aluminum-Lithium (Al-Li) 2050 Blanks for Spin Forming

Spin forming is a metalworking process used to produce axially-symmetric parts by the application of lateral pressure from a forming tool to a circular blank mounted on a revolving, heated mandrel. This process enables production of cones, domes, or caps that can incorporate features such as flat bulkheads. Friction stir welded (FSW) Al-Li 2195 blanks have been spin formed for single-piece, launch vehicle domes up to 5 m in diameter. Al-Li 2050 plate has the advantage of availability up to 165 mm thick,whereas 2195 plate is traditionally only available up to 57 mm thick. Increased plate thickness enables greater design flexibility for machining of taller structural elements, such as stiffeners. However, 2050 is only available in plate widths up to ~3.9 m. This study demonstrated the capability of making custom 2050 blanks for spin forming large (>7.8 m in diameter), single-piece, light-weight flight articles (e.g., lunar habitats and landers) by FSW two (or more) plates together. Spin forming of welded 2050 blanks has not been attempted before this investigation. The objective was to evaluate the mechanical properties of 2050 parent material and FSW regions after spin forming and subsequent T6 heat treatment. A 3 m diameter x 38 mm thick 2050 blank was fabricated from two 1.5 m wide pieces by FSW and subsequently spin formed into a cone article. Additionally, three 1.2 m diameter x 8.1 mm thick 2050 blanks were produced by FSW and spin formed into subscale cone articles with varying strain input. Mechanical properties are reported for spin formed 2050-T6 parent material and FSW regions on the articles. Abnormal grain growth (AGG) in the FSW during heat treatment was expected based on prior experience with 2195. Implementation of an AGG suppression strategy during heat treatment reduced AGG in the FSW regions and improved ductility.

Spin forming↗
NASA NTRSDate not supplied

Computational Icing Analysis on NASA’s SIDRM Geometry to Investigate Collection Efficiency

Computational icing analysis results were compared to experimental icing tunnel data including aerothermal (e.g. dry air) and supercooled water droplet rime-ice conditions from tests conducted in early 2022 at the NASA Icing Research Tunnel (IRT). The Simulated Intercompressor Duct Research Model (SIDRM) test article was used in this study, and its geometry represents the inter-compressor duct region of a turbofan engine. The test article’s purpose is to study the physics of supercooled water icing and ice crystal icing. This study compared three different icing codes: FENSAP-ICE (Eulerian approach), LEWICE3D (Lagrangian approach), and GlennICE (Lagrangian approach). All three icing codes were conducted on SIDRM’s complex body flow-field and compared to different experimental supercooled water rime runs. The test article instrumentation (pressure taps, thermocouples, etc.) and 3D laser scans of final ice shapes were used to compare against the different icing code simulations. The overall objectives are to understand how the icing codes handle capturing collection efficiency on the complex test article’s unheated surfaces. In the aerothermal cases, pressure tap readings matched the CFD results, but dry air CFD underpredicted thermocouple readings. Collection efficiency results from all three icing codes matched well together on the main body leading edge, main body slope, and the strut leading edges. All three icing codes underpredicted the final ice shape using a single-shot constant ice density approach, with more difficulty coming from the strut leading edge ice shape due to the swept wing like flow field. NASA’s overall goal for this effort is to develop computational icing tools to assist in the design and certification of engines for flight in icing conditions.

Supercooled icing, SIDRM, aerodynamics, computatio↗
NASA NTRSDate not supplied

CDISC Remote Design Method to Simulate Aircraft Interference Effects for the CATNLF Flight Test

The Crossflow Attenuated Natural Laminar Flow (CATNLF) flight test is an experimental evaluation of a laminar flow design approach that will be conducted on a small transonic wing-like model suspended from the Centerline Instrumented Pylon (CLIP) underneath an F-15 vehicle. The upcoming flight test is anticipated to provide experimental data that may be useful for transition prediction code calibrations or laminar flow computational studies. For the computational results to be reliable, relevant aerodynamics must be included in the simulations. It has been shown that the F-15 vehicle produces a notable aerodynamic influence on the CATNLF test article, and that removing the F-15 vehicle from the simulation significantly alters the laminar flow behavior. Because the F-15 vehicle is not a publicly-releasable geometry, if the CATNLF flight test dataset is to be widely used by the community, an alternative method to ensure that the CATNLF test article flight configuration aerodynamics is properly modeled is required. The objective of this research is to use the CDISC aerodynamic design tool to perform a remote design of a surface that would replicate the aerodynamic influence on the CATNLF test article. This new design approach, referred to as the Equivalent Loading via Interference Surface Effects (ELISE), is presented in this paper, including the details on the design setup and results. The ELISE design successfully reproduced the F-15 influence on the CATNLF test article, with predicted transition location within an average of 4% chord of the full flight configuration. An off-design assessment is included for small perturbations in Mach and sideslip angle that suggest the ELISE design can cover a reasonable range of expected flight conditions with the same geometry. This paper is intended to explain the design approach, as well as show results that can be expected when utilizing the ELISE design method.

CDISC↗
NASA NTRSDate not supplied

CDISC Remote Design Method to Simulate Aircraft Interference Effects for the CATNLF Flight Test

The Crossflow Attenuated Natural Laminar Flow (CATNLF)flight test is an experimental evaluation of a laminar flow design approach that will be conducted on a small transonic wing-like model suspended from the Centerline Instrumented Pylon (CLIP) underneath an F-15 vehicle. The upcoming flight testis anticipated to provide experimental data that may be useful for transition prediction code calibrations or laminar flow computational studies. For the computational results to be reliable, relevant aerodynamics must be included in the simulations. It has been shown that the F-15 vehicle produces a notable aerodynamic influence on the CATNLF test article, and that removing the F-15 vehicle from the simulation significantly alters the laminar flow behavior. Because the F-15 vehicle is not a publicly-releasable geometry, if the CATNLF flight test dataset is to be widely used by the community, an alternative method to ensure that the CATNLF test article flight configuration aerodynamics is properly modeled is required. The objective of this research is to use the CDISC aerodynamic design tool to perform a remote design of a surface that would replicate the aerodynamic influence on the CATNLF test article. This new design approach, referred to as the Equivalent Loading via Interference Surface Effects (ELISE), is presented in this paper, including the details on the design setup and results. The ELISE design successfully reproduced theF-15 influence on the CATNLF test article, with predicted transition location within an average of 4% chord of the full flight configuration. An off-design assessment is included for small perturbations in Mach and sideslip angle that suggest the ELISE design can cover a reasonable range of expected flight conditions with the same geometry. This paper is intended to explain the design approach, as well as show results that can be expected when utilizing the ELISE design method.

CDISC↗
NASA NTRS2024

Exploration Extravehicular Mobility Unit (xEMU) Chamber B Thermal Vacuum “Suit 2” Pressure Garment System Hardware and Test Design

NASA’s Exploration Extravehicular Mobility Unit (xEMU) is the government reference next-generation space suit design and is engineered to protect astronauts from extreme lunar environmental temperatures. To evaluate the xEMU hardware thermal requirements, the xEMU Testing Team invented, designed, and executed a dual-suit, uncrewed thermal vacuum (TVAC) test at Johnson Space Center’s (JSC) Chamber B. This paper details the test hardware design and test methodology for the “Suit 2” Pressure Garment System (PGS) test article. The uncrewed “Suit 2” PGS test article consisted of a full PGS assembly with simulated Portable Life Support System (PLSS) functionality provided by test equipment, including a ventilation loop and two distinct thermal control loops. This paper will discuss in depth the test hardware design, including internal suit thermal boundary simulation, sensor quantity and placement, test support equipment rigs for gas flow, water flow, and power. Custom hardware designed to add additional penetrations to the suit or provide additional interfaces for sensors will also be discussed. This paper will also address the assembly and integration sequence for the test article. In addition to test hardware design, this paper will provide insights into the test methodology for this test article, including a discussion of thermal steady-state testing, simultaneous dual suit thermal vacuum testing, and hazard mitigation and controls.

xEMU↗
NASA NTRSDate not supplied

Advanced Manufacturing of the SLS Block 1B Payload Adapter: Engineering Development Unit

The Payload Adapter (PLA) for the Space Launch System Block 1B configuration (SLS B1B) is a structural frustum that interfaces with the Exploration Upper Stage (EUS) and the Universal Stage Adapter (USA) or the cargo fairing at the aft end and interfaces with the primary or co-manifested payload at the forward end. The primary structure is composed of a single piece aluminum forward ring, a segmented aluminum aft ring, eight carbon composite sandwich panels, and composite bonded longitudinal joints. The PLA is designed and fabricated in-house at MSFC. Development and flight article builds are joint efforts between the Materials & Processes Laboratory (EM), the Spacecraft & Vehicle Systems Department (EV), the Space Systems Department (ES), and the Test Laboratory (ET). The large-scale manufacturing efforts associated with this task include the fabrication of pathfinder panels; the fabrication and assembly of a full-scale Manufacturing Demonstration Article (MDA); the fabrication, assembly, and test of an Engineering Demonstration Unit (EDU); the fabrication, assembly, and test of a qualification Article; and the fabrication, assembly, and proof test of subsequent flight articles. The PLA team recently completed assembly of the EDU. The final assembly included the instillation of the co-manifested secondary payload, NEST, and the instillation and tensioning of the Payload Separation System (PSS). This poster will detail the assembly process which includes: (1) composite panel to ring assembly, (2) adhesive bonding for the longitudinal joints, and (3) instillation of the NEST and PSS. The PLA EDU is prepared to undergo modal and static load testing.

Monique Wallace↗
DOE OSTI2024

Rotating cylinder electrode in reactive CO 2 capture: Identifying active C species via transport, VLE models and kinetics

Here, this article explores technical challenges and potential methodologies for understanding electrochemical Reactive CO 2 Capture (RCC) mechanisms. RCC offers potential energy cost advantages by directly converting captured CO 2 into fuels and chemicals, unlike traditional carbon capture and utilization (CCU) processes that require sequential capture, concentration, and compression. However, direct conversion of captured CO 2 introduces complexity due to additional equilibrium buffer reactions, making it challenging to identify active species for reduction in electrochemical studies. This article discusses methods to integrate transport, thermodynamics, and kinetics concepts to identify active carbon sources in RCC. Vapor‐Liquid Equilibrium (VLE) and transport models are validated against experimental results obtained in a gastight rotating cylinder electrode reactor and are shown as useful tools for studying RCC in heterogeneous electrocatalysts across different capture agents, solvents, and temperatures. This article establishes an experimental framework for advancing research in electrochemical RCC.

Electrocatalysis↗
DOE OSTI2025

Environmental life cycle of fentanyl: From the cradle to an unknown grave

The lack of available information on the presence and persistence of fentanyl in the environment is a significant gap in the technical literature. Although the origins of the opioid in the environment are well-known because they follow the same pathways of other drug-related environmental contaminants, the downstream effects of fentanyl in the water supply and its retention in soil are less understood. The characterization of fentanyl and its potential degradation products in complex environmental samples such as soil is severely understudied. Very few articles are available that work to identify fentanyl and its degradation products in complex samples or name the possible hazards that may result from environmental exposure and degradation. Therefore, the objectives were to identify available articles focused on environmental fentanyl and its pathways and highlight quantifiable research or results that included specific degradation products or downstream effects. Research articles focused on fentanyl between 2000 and 2024 were identified and reviewed and then filtered using Boolean search terms for environmental parameters. Various studies have determined that trace levels of fentanyl can be found in a variety of environments, and additional data suggest preferential partitioning into soils from water and long-term persistence. Despite this knowledge, very little data exists on the long-term downstream effects of fentanyl or its analogs. As the chronic effects from low-level fentanyl exposure are currently unknown, this lack of insight brings to the forefront the need for further research to improve our understanding of fentanyl persistence, degradation, and toxicity within the environment.

54 ENVIRONMENTAL SCIENCES↗