Results for “220600 -- Nuclear Reactor Technology-- Research, Test & Experimental Reactors” — Page 4

NASA NTRS2015

Development of Li-Metal Battery Cell Chemistries at NASA Glenn Research Center

State-of-the-Art lithium-ion battery technology is limited by specific energy and thus not sufficiently advanced to support the energy storage necessary for aerospace needs, such as all-electric aircraft and many deep space NASA exploration missions. In response to this technological gap, our research team at NASA Glenn Research Center has been active in formulating concepts and developing testing hardware and components for Li-metal battery cell chemistries. Lithium metal anodes combined with advanced cathode materials could provide up to five times the specific energy versus state-of-the-art lithium-ion cells (1000 Whkg versus 200 Whkg). Although Lithium metal anodes offer very high theoretical capacity, they have not been shown to successfully operate reversibly.

battery↗

NASA NTRS1970

The Entry Plasma Sheath and Its Effects on Space Vehicle Electromagnetic Systems, Volume 1

This symposium is the fourth in a series on the plasma sheath. The first three were held in Boston, Massachusetts, and were sponsored by the Air Force Cambridge Research Laboratories. The papers included in the symposium cover theoretical and experimental results of research and flight tests of many different specific aspects of the general problem of plasma sheath degradation of reentry vehicle electromagnetic systems. Only the newer advances and more recent developments in the field are included in the Fourth Plasma Sheath Symposium, with no particular attempt being made to review the general background and history of the problem or to cover those aspects discussed in the previous symposia. Flight data from both NASA and DOD programs are included.

Source record↗

NASA NTRSDate not supplied

Experimental and Numerical Investigation of the NASA High Efficiency Centrifugal Compressor Vaned Stage Geometry and Aerodynamic Performance

Since its inception in the early 2010s, the NASA High Efficiency Centrifugal Compressor (HECC) has been enigmatic for the propulsion research community: experimental data and numerical simulations of the stage have generally not aligned in their quantifications of performance metrics. Typically, the fault for these disagreements is assigned to the numerical simulations as the simulations are models, and models are inherently incomplete representations of the experiment. This “incompleteness” may manifest in assumptions regarding roughness or heat transfer, simplifications of the flow path (i.e., neglecting bleed flows), or the oft-scapegoated turbulence model. In the case of HECC, recent work showed unexpected discrepancies between the intended impeller geometry defined in the design report (termed Design-Intent) and the manufactured impeller used in the experimental campaigns (termed As-Manufactured). That work used numerical simulations to establish that the geometric differences between the Design-Intent and As-Manufactured impellers were significant enough to result in drastically different performance predictions for the HECC vaneless diffuser configuration. The Design-Intent impeller simulations over predicted the performance relative to the experiment, whereas the As-Manufactured simulations better represented the experimental data, both in terms of one-dimensional performance metrics and spanwise flow profiles. This effort expands on that work by examining in detail the geometry and aerodynamic performance of the HECC vaned diffuser configuration. Further differences between the Design-Intent and As-Manufactured geometries have been discovered in the vaned diffuser and exit guide vanes, and these differences are documented herein. The summations of the geometric differences for all of the components were used to create two numerical models of HECC vaned diffuser configuration: the Design-Intent simulations which are generated from the original geometry definitions given in the design report and the As-Manufactured simulations which are the best available representation of the manufactured compressor hardware used in the experimental test campaigns. In congruence with the earlier vaneless diffuser work, the numerical predictions of the Design-Intent choked mass flow rate, total pressure ratio, and efficiency were notably greater than that of the As-Manufactured simulations. To increase confidence in the experimental dataset, measurements from a recent test campaign conducted in 2024 are used to validate the original experimental data acquired from 2012 to 2014 with good repeatability overall, especially considering the passage of time and differences in the data acquisition systems between the test campaigns. Both numerical simulations were then extensively evaluated against the experimental data. The As-Manufactured simulations provided better estimates of the stage performance than the Design-Intent cases in terms of most performance metrics. Nonetheless, more detailed results still show opportunities for improvement. Despite a more accurate representation of the physical hardware, characterization of the impeller work input remains a challenge even for rigorously developed numerical models.

vaned diffuser↗

NASA NTRSDate not supplied

Experimental and Numerical Investigation of the NASA High Efficiency Centrifugal Compressor Vaned Stage Geometry and Aerodynamic Performance

Since its inception in the early 2010s, the NASA High Efficiency Centrifugal Compressor (HECC) has been enigmatic for the propulsion research community: experimental data and numerical simulations of the stage have generally not aligned in their quantifications of performance metrics. Typically, the fault for these disagreements is assigned to the numerical simulations as the simulations are models, and models are inherently incomplete representations of the experiment. This “incompleteness” may manifest in assumptions regarding roughness or heat transfer, simplifications of the flow path (i.e., neglecting bleed flows), or the oft-scapegoated turbulence model. In the case of HECC, recent work showed unexpected discrepancies between the intended impeller geometry defined in the design report (termed Design-Intent) and the manufactured impeller used in the experimental campaigns (termed As-Manufactured). That work used numerical simulations to establish that the geometric differences between the Design-Intent and As-Manufactured impellers were significant enough to result in drastically different performance predictions for the HECC vaneless diffuser configuration. The Design-Intent impeller simulations over predicted the performance relative to the experiment, whereas the As-Manufactured simulations better represented the experimental data, both in terms of one-dimensional performance metrics and spanwise flow profiles. This effort expands on that work by examining in detail the geometry and aerodynamic performance of the HECC vaned diffuser configuration. Further differences between the Design-Intent and As-Manufactured geometries have been discovered in the vaned diffuser and exit guide vanes, and these differences are documented herein. The summations of the geometric differences for all of the components were used to create two numerical models of HECC vaned diffuser configuration: the Design-Intent simulations which are generated from the original geometry definitions given in the design report and the As-Manufactured simulations which are the best available representation of the manufactured compressor hardware used in the experimental test campaigns. In congruence with the earlier vaneless diffuser work, the numerical predictions of the Design-Intent choked mass flow rate, total pressure ratio, and efficiency were notably greater than that of the As-Manufactured simulations. To increase confidence in the experimental dataset, measurements from a recent test campaign conducted in 2024 are used to validate the original experimental data acquired from 2012 to 2014 with good repeatability overall, especially considering the passage of time and differences in the data acquisition systems between the test campaigns. Both numerical simulations were then extensively evaluated against the experimental data. The As-Manufactured simulations provided better estimates of the stage performance than the Design-Intent cases in terms of most performance metrics. Nonetheless, more detailed results still show opportunities for improvement. Despite a more accurate representation of the physical hardware, characterization of the impeller work input remains a challenge even for rigorously developed numerical models.

centrifugal compressor↗

NASA NTRS2024

Wave Rotor Enhanced Nuclear (WREN) Propulsion: NASA Innovative Advanced Concepts (NIAC) - Phase I Final Report

Nuclear Thermal Propulsion (NTP) is identified as one of the preferred propulsion technologies for manned missions throughout the solar system (NASA MSFC).[1, 2] The state-ofthe-art NTP cycle is based on a solid core Nuclear Engine for Rocket Vehicle Application (NERVA)[3] class technology (Fig. 1) that is envisioned to provide a specific impulse of 900 seconds doubling chemical rocket performance (450 seconds). Even with this impressive increase, the NTP NERVA designs still have issues providing adequate initial to final mass fractions for high ΔV missions.[4] Nuclear Electric Propulsion (NEP) can provide extremely high Isp (2,000 to over 10,000 seconds) but with only low thrust and limits on mass to power ratios. The need for an electric power source also adds the issue of heat rejection in space where thermal energy conversion is at best 30-40% under ideal conditions. NASA Space Technology Mission Directorate (STMD) has recently expressed interest in finding advanced nuclear propulsion technology through the NASA Go:Thrust RFI.[5, 6] A novel Wave Rotor (WR) topping cycle has been proposed for our NASA NIAC concept. It promises to deliver similar thrust as NERVA class NTP propulsion, but with Isp in the 1,200-2,000 second range. Coupled with an NEP cycle, the duty cycle Isp can further be increased (1,800-4,000 seconds) with minimal addition of dry mass. This bimodal design enables fast transit trajectories for manned missions to Mars and revolutionizes the deep space exploration of our solar system.

Nuclear Thermal Propulsion↗

NASA NTRSDate not supplied

Thermal Cycle and Capacitance Testing of Mezzo Industries Phase Change Material Heat Capacitor

A phase change material heat capacitor prototype, designed and built under a NASA Small Business Innovation Research grant by Mezzo Technologies, was tested through coolant thermal cycling within the expected flow rates and temperature range of the Orion spacecraft’s propylene glycol-water mixture coolant system. Testing was performed with the phase change material contained under sealed conditions after a degassing procedure and exposed to atmosphere without prior degassing. The useful heat storage capacitance was measured and compared to Orion spacecraft requirements and theoretical N-Pentadecane storage capacity.

Propylene Glycol Water↗

NASA NTRS2024

AC and DC Fault Management for Megawatt Electrified Aircraft Electrical Powertrains - Task 2: Power Quality Filtering Using Nanocrystalline Soft Magnetic Inductor

The NASA RTAPS program on AC and DC Fault Management for Megawatt Electrified Power Train is a multi-year joint project with Pratt & Whitney (P&W), Collins Aerospace (CA), and RTX Technology Research Center (RTRC). This research program focuses on the high-voltage distribution issues that present a significant technological obstacle in the adoption of Electrified Aviation Propulsion (EAP) systems. One challenge to the adoption of high-voltage distribution systems with power electronic converters is the need for filter elements to limit the generation and propagation of noise, protect the cable systems from premature aging and prevent against excessive heating within subcomponents due to high-frequency induced currents. While increased distribution voltages aide in reducing the cable mass for a fixed power system, the associated mass with the filtering elements for power electronic converter can grow with increasing distribution voltages – thereby mitigating any benefit associated with increasing the distribution system voltage. To enable high-voltage distribution systems with high system specific power densities, new magnetic materials must be developed. Therefore, the second task of the NASA RTAPS program is associated with the design and application of advanced soft magnetic components for Megawatt class electric propulsion systems, specifically the motor drive system. This report covers the collaborative work between NASA Glenn Research Center (GRC), RTRC, P&W and CA in the development of three types of magnetic components over the span of the three-year program. These critical magnetic components are the DC side EMI filter, which limits the propagation of harmful electromagnetic noise to the rise of the distribution system, and the AC side damping with the dv/dt filter, which limits the fast rise time of the power electronic converter output voltage to limit the degradation on the cable/motor insulation systems. Each of these components are investigated from component level design and are optimized at the system level with a combined modelling and testing effort. In the final experimental evaluation of the NASA developed soft magnetic material with a dv/dt filter, a commercial-off-the-shelf (COTS) magnetic core and the GRC magnetic core are optimized and loaded at 320Arms to evaluate their difference in performance. After a run time of 30 minutes in a MW-class motor driver at RTRC, the NASA GRC cores were found to not only offer a lower temperature rise of nearly 25°𝐶, but also a reduction in measured core loss of 25% (12.75W to 9.5W).

Elecrified Aircraft Propulsion↗

NASA NTRS1995

Air-Breathing Aerospace Plane Development Essential: Hypersonic Propulsion Flight Tests

Hypersonic airbreathing propulsion utilizing scramjets can change transatmospheric accelerators for low earth-to-orbit and return transportation. The value and limitation of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. It is proposed that near full-scale hypersonic propulsion flight tests are essential for developing computational design technology so that it can be used for designing this system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests.

Unmeel B Mehta↗

NASA NTRS1997

Analysis of Impact Induced Damage and its Effect on Structural Integrity of Space Flight Composite Overwrapped Pressure Vessels

The objective of this research work has been to provide analytical background and support to the ongoing experimental program at NASA, White Sands Test Facility, involving testing composite overwrapped pressure vessels (COPV) for impact damage and cyclic pressurization. Preliminary theoretical basis, including the governing equations for a shallow shell subjected to internal pressure, has been established. Effects of the Griffith type cracks on the structural integrity of the cylindrical vessel were evaluated by methods of Fracture Mechanics. The results indicate that the effective mass of the pressure vessel is an important factor influencing the response to impact events. We also have found that the material properties of the target, contained in the constitutive equations of the composite attached to the Aluminum liner, dominate the impact event in the low velocity range, the material properties become less important, while the target mass distribution and the impactor mass become more significant as the velocity of the impactor increases. Therefore, at high-velocity impact it is not only the kinetic energy of the impactor but also its mass which has a significant effect on the dynamics of the event, and consequently on the induced damage. This work also suggests a methodology for an assessment of the rate of loading effects on the degradation of the material toughness associated with a high-velocity impact where the rate effects become significant. To model the rate dependence of the material response a viscoelastic-plastic constitutive equations were assumed, and on this basis predictions are made regarding the rate dependent material resistance curve. Other dynamic phenomena associated with the impact event have been treated in the framework of the Computational Mechanics using the courtesy of Prof. P. Guebelle and his graduate student at University of Illinois at Urbana-Champaign who have an access to a super-fast computer located on their campus. Finally, the guidelines for a follow-up research program are provided in the body of this report. They address three major areas: theoretical research, numerical studies, and further experimental work.

Michael P Wnuk↗

NASA NTRS2012

Analysis of Screen Channel LAD Bubble Point Tests in Liquid Methane at Elevated Temperature

This paper examines the effect of varying the liquid temperature and pressure on the bubble point pressure for screen channel Liquid Acquisition Devices in cryogenic liquid methane using gaseous helium across a wide range of elevated pressures and temperatures. Testing of a 325 x 2300 Dutch Twill screen sample was conducted in the Cryogenic Components Lab 7 facility at the NASA Glenn Research Center in Cleveland, Ohio. Test conditions ranged from 105 to 160K and 0.0965 – 1.78 MPa. Bubble point is shown to be a strong function of the liquid temperature and a weak function of the amount of subcooling at the LAD screen. The model predicts well for saturated liquid but under predicts the subcooled data.

Reaction Control System↗

NASA NTRS1966

Research Achievements Review, Volume II

Materials research - diffusion bonding methods, stress corrosion tests, nondestructive testing, material designs for electronic equipment, and nonmetallic material development.

Materials Science↗

NASA NTRSDate not supplied

Characterization Testing of a 20 Kelvin Cryocooler for Space Applications

Future NASA space exploration missions will require long-duration storage and liquefaction of cryogenic liquids, enabled by active cooling provided by cryocoolers. Recent gap analyses of Lunar and Mars transportation systems have identified 20 K-class cryocoolers as a critical enabling technology for chemical and nuclear thermal propulsion architectures using liquid hydrogen propellant. To address this technology gap, NASA has undergone the development of a high-efficiency, high-capacity 20 K cryocooler via an SBIR partnership with Creare. While the January 2025 testing demonstrated functionality and compliance with contractual requirements, the objective of the NASA-led characterization effort was to generate a dataset for supporting future mission designs across a broader operating envelope, including off-nominal conditions. The cryocooler demonstrated strong performance, achieving a peak coefficient of performance of 17.68% relative to Carnot efficiency and a maximum lift capacity of 24.4 W at 21 K. Overall, the results confirm that the 20 W 20 K cryocooler provides a flexible range of capabilities to enable zero boil-off storage of liquid hydrogen for future Lunar and Mars missions. The data collected provide a strong foundation for model validation and future system design efforts.

Cryocooler↗

NASA NTRSDate not supplied

Characterization Testing of a 20 Kelvin Cryocooler for Space Applications

Future NASA space exploration missions will require long-duration storage and liquefaction of cryogenic liquids, enabled by active cooling provided by cryocoolers. Recent gap analyses of Lunar and Mars transportation systems have identified 20 K-class cryocoolers as a critical enabling technology for chemical and nuclear thermal propulsion architectures using liquid hydrogen propellant. To address this technology gap, NASA has undergone the development of a high-efficiency, high-capacity 20 K cryocooler via an SBIR partnership with Creare. While the January 2025 testing demonstrated functionality and compliance with contractual requirements, the objective of the NASA-led characterization effort was to generate a dataset for supporting future mission designs across a broader operating envelope, including off-nominal conditions. The cryocooler demonstrated strong performance, achieving a peak coefficient of performance of 17.68% relative to Carnot efficiency and a maximum lift capacity of 24.4 W at 21 K. Overall, the results confirm that the 20 W 20 K cryocooler provides a flexible range of capabilities to enable zero boil-off storage of liquid hydrogen for future Lunar and Mars missions. The data collected provide a strong foundation for model validation and future system design efforts.

Zero boil-off↗

NASA NTRS1985

Life cycle test results of a bipolar nickel hydrogen battery

A history is given of low Earth orbit (LEO) laboratory test data on a 6.5 ampere-hour bipolar nickel hydrogen battery designed and built at the NASA Lewis Research Center. The bipolar concept is a means of achieving the goal of producing an acceptable battery, of higher energy density, able to withstand the demands of low-Earth-orbit regimes. Over 4100 LEO cycles were established on a ten cell battery. It seems that any perturbation on normal cycling effects the cells performance. Explanations and theories of the battery's behavior are varied and widespread among those closely associated with it. Deep discharging does provide a reconditioning effect and further experimentation is planned in this area. The battery watt-hour efficiency is about 75 percent and the time averaged, discharge voltage is about 1.26 volts for all cells at both the C/4 and LEO rate. Since a significant portion of the electrode capacity has degraded, the LEO cycle discharges are approaching depths of 90 to 100 percent of the high rate capacity. Therefore, the low end-of-discharge voltages occur precipitously after the knee of the discharge curve and is more an indication of electrode capacity and is a lesser indicator of overall cell performance.

Cataldo, R. L.↗

NASA NTRS1990

NASA Systems Autonomy Demonstration Project: Advanced Automation Demonstration of Space Station Freedom Thermal Control System

The NASA Systems Autonomy Demonstration Project (SADP) was initiated in response to Congressional interest in Space station automation technology demonstration. The SADP is a joint cooperative effort between Ames Research Center (ARC) and Johnson Space Center (JSC) to demonstrate advanced automation technology feasibility using the Space Station Freedom Thermal Control System (TCS) test bed. A model-based expert system and its operator interface were developed by knowledge engineers, AI researchers, and human factors researchers at ARC working with the domain experts and system integration engineers at JSC. Its target application is a prototype heat acquisition and transport subsystem of a space station TCS. The demonstration is scheduled to be conducted at JSC in August, 1989. The demonstration will consist of a detailed test of the ability of the Thermal Expert System to conduct real time normal operations (start-up, set point changes, shut-down) and to conduct fault detection, isolation, and recovery (FDIR) on the test article. The FDIR will be conducted by injecting ten component level failures that will manifest themselves as seven different system level faults. Here, the SADP goals, are described as well as the Thermal Control Expert System that has been developed for demonstration.

Jeffrey Dominick↗

NASA NTRS1997

Future Aerospace Technology in the Service of the Alliance: Sustained Hypersonic Flight - Volume 3

Sustained Hypersonic Flight capability is envisioned for several future military and civilian applications, such as long range immediate reaction reconnaissance, high speed interception of air targets, long range precision strike against hardened or time critical targets, and access to space. The inherent reduction in time-to-target and low vulnerability will permit new operational tactics. The symposium outlined mission needs and operational scenarios for hypersonic vehicles. Addressed were: Technological issues and challenges in external hypersonic aerodynamics and design, propulsion and engine/airframe integration, military ramjet applications, overall system design including structures and materials development needs, and test facilities. Ground based test facilities cannot provide full simulation at Mach numbers above 5. Therefore, the use of Computational Fluid Dynamics design tools is essential. For routine use of the computational tools in analysis, design and optimization, it is necessary to reduce the time for the entire computational process by two orders of magnitude. Hypersonic boundary layer transition remains a critical design issue because of the important impact on engine drag and on heating, which can affect the choice of materials and thermal protection systems. In addition, reliable prediction of scramjet net thrust is an absolute must in resolving hypersonic air breathing vehicle design issues. Due to current facility and computational shortfalls, the development of future hypersonic flight systems requires research flight tests in the technology areas of boundary layer transition and air-breathing propulsion engine performance. For sustained hypersonic flight beyond Mach 6, the supersonic combustion ramjet (scramjet) engine is the only choice for the near future. Only this air-breathing concept offers a significant promise of large reductions in required propellant fractions, increased payload fractions, and reduced size vehicles, together with a foreseeable technological feasibility. Airframe/engine integration, combustor design and thermal management are the predominant engineering tasks. Fuels, hydrogen or hydrocarbon, must be matched to the operational needs of military or civil use. Experience in existing ramjet propelled missiles capable of speeds up to Mach 4 can support the development effort. The potential mission and cost benefits of sustained hypersonic flight to both military and civil applications are tremendous. From the budget point of view, the possibility of sharing development costs between military and civil programs offers a specific advantage.

Combustion chambers↗

NASA NTRSDate not supplied

Experimental Studies of Open Rotor Installation Effects

Open rotor propulsion technologies offer an opportunity for reducing fuel burn. At the same time, great challenges arise from this radically different engine architecture in terms of aircraft system integration. The aeroacoustic effects of integration are one of those key challenges and, furthermore, can offer the possibility for total aircraft system noise reduction. The inter-relationship of the aerodynamic and aeroacoustic system integration effects is particularly important to enable future application. An extensive model scale test campaign was conducted to investigate a broad range of these open rotor installation effects for both a conventional and an unconventional airframe. The experimental campaign was conducted in the Boeing Low Speed Aeroacoustic Facility with specially designed modifications for efficient positioning of the open rotor rig relative to the airframe. The airframe was traversed remotely relative to the fixed open rotor rig providing for the investigation of a large number of installation positions. Eight positions around the main wing of the conventional airframe and eleven positions above the hybrid wing body airframe were documented. Other parameters investigated in the test campaign included forward flight simulation Mach number, angle of attack, rotor speed, and airframe control surface deployment. In addition, both airframes were modified for alternate configurations. The conventional airframe was configured with both a T and a U tail while the unconventional hybrid wing body airframe was configured with different vertical control surfaces. An extensive instrumentation package was deployed. The acoustics were documented with a fixed array of far field microphones, a traversing array of in-flow microphones, an out-of-flow traversing phased array, and unsteady pressure transducers mounted flush on the surface of the airframes. In addition, mean flow surveys were measured with an articulating arm traversing system. The flow field surveys were particularly valuable in documenting flow distortion effects for the various installation positions and those created by angle of attack. This presentation will report the key results obtained for open rotor installation effects and discuss future prospects with the perspective of these results.

Michael J Czech↗

NASA NTRS1990

Experimental Investigations of Low-Energy (4 to 40 eV) Collisions of O - ( 2 P) Ions and O( 3 P) Atoms With Surfaces

Using a newly-developed, magnetically confined source, low-energy, ground state oxygen negative ions and neutral atoms are generated. The energy range is variable, and atom and neutrals have been generated at energies varying from 2 eV to 40 eV and higher. It was found that the interaction of these low-energy species with a solid magnesium fluoride target leads to optical emissions in the (at least) visible and infrared regions of the spectrum. Researchers describe y details of the photodetachment source, and present spectra of the neutral and ion glows in the wavelength range 250 to 850 nm (for O - ) and 600 to 850 nm (for O), and discuss the variability of the emissions for incident energies between 4 and 40 eV.

A Chutjian↗