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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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At least 991 records · Page 55
All sourcesNASA NTRSDOE OSTI
NASA NTRS2023

High Temperature / High Frequency Temperature Measurement: Achieving MHz Response at Temperatures > 1000 Deg-F

High frequency temperature measurements are not possible with current SOA temperature measurement devices at temperatures higher than ~700 deg-F. Temperature measurements above ~1000 deg-F are typical for entry vehicle thermal protection systems (TPS), rocket engines, ram/scramjet engines, etc. A new technology capable of achieving ~1200 deg-F temperature measurements with MHz response has been achieved. Investigations are underway to extend the concept to even higher temperatures. A U.S. patent application is about to be submitted.

High Frequency↗
NASA NTRSDate not supplied

LES Analysis of Loss Anomaly in a Turbine Blade with Large Trailing Edge Radius

There is great interest in the usage of Ceramic Matrix Composites (CMC) as a turbine blade material. CMCs can tolerate higher temperatures than alloyed metallic blades, resulting in better thermal efficiency of gas turbine engines. However, depending on the manufacturing process of the CMC, blades may have a larger trailing edge thickness. The design space therefore needs to be updated due to the resulting flow physics. Recently, experimental results acquired at NASA Glenn Transonic Turbine Blade Cascade Rig showed that a loss measure generally increased with increasing trailing edge thickness. For one case however, the losses initially decreased with increasing Reynolds number but rose and peaked at Re ~1.24x10 6 and then subsequently dropped to the original pre-peak trend. A possible cause of this intriguing phenomenon was speculated to be transonic vortex shedding, which is the mechanism of vortex shedding promoted by reflected shed pressure waves at the trailing edge at relatively high Reynolds numbers and higher Mach numbers. A Reynolds-averaged Navier-Stokes (RANS) analysis or a low-resolution large eddy simulation (LES) does not reproduce this apparent anomaly and thus it is worth performing a wall-resolved LES (the total mesh count of ~290 million cells) to investigate the aerodynamics of the CMC blade in the context of transonic vortex shedding. Our numerical results at Re=1,246,350 show that the pressure waves generated by the vortex shedding in the wake travel upstream and significantly influence the transition and separation on the suction side thus enhancing the said vortex shedding in the wake. This feedback seems not to hold under a low-Re condition (Re=621,900). The Reynolds number dependence was also examined by numerical perturbation of the pressure waves in the wake and by examining how such perturbation attenuates or endures. It is confirmed that the perturbation of the pressure waves is quickly damped.

LES↗
NASA NTRSDate not supplied

Highlights of the Test of the QUiet Electric Engine (QUEEN V1) in the NASA GRC Acoustical Testing Laboratory (ATL)

This presentation for the NASA Acoustics Technical Working Group meeting highlights preliminary acoustic and thermal test results from the Quiet Electric Engine (QUEEN V1) experiment recently performed in the NASA Glenn Research Center Acoustical Testing Laboratory. The QUEEN V1 features a single Commercial-Off-the-Shelf fan with an Outlet Guide Vane. Four inlet duct acoustic liners were tested. Electronic Speed Controller (ESC) temperatures were recorded for two different cooling methods: air cooling and a cold plate.

Acoustics↗
DOE OSTI2026

High energy neutrinos from pulsar-powered optical transients: LFBOTs as potential origin of the KM3NeT event KM3-230213A

Recently, the KM3NeT Collaboration reported the detection of an ultra-high energy ($\sim 220$ PeV) neutrino event, KM3-230213A. In this work, we perform a detailed investigation into whether this event could originate from the diffuse neutrino flux produced by a class of pulsar-powered optical transients. In particular, we consider populations of ordinary supernovae (SNe), super-luminous supernovae (SLSNe), and luminous fast blue optical transients (LFBOTs) with a newly formed magnetar as the central engine. We discuss both the thermal electromagnetic and non-thermal neutrino emission from such sources. We scan the parameter space of the dipolar magnetic field strength and the initial spin period to determine characteristic optical emission properties and lightcurve timescales of these transients. Additionally, our scan identifies which classes of these transients can reproduce the required diffuse flux level and neutrino energies. Combining our results, we conclude that a diffuse neutrino flux from a population of LFBOTs can explain the KM3NeT event. Therefore, pulsar-powered optical transients may serve as promising sources for the current and upcoming high-energy and ultra-high energy neutrino telescopes.

Mukhopadhyay, Mainak [Fermilab; Chicago U., KICP] ↗
NASA NTRS1986

Evaluation of a Stirling engine heater bypass with the NASA Lewis nodal-analysis performance code

In support of the U.S. Department of Energy's Stirling Engine Highway Vehicle Systems program, the NASA Lewis Research Center investigated whether bypassing the P-40 Stirling engine heater during regenerative cooling would improve engine performance. The Lewis nodal-analysis Stirling engine computer simulation was used for this investigation. Results for the heater-bypass concept showed no significant improvement in the indicated thermal efficiency for the P-40 Stirling engine operating at full-power and part-power conditions. Optimizing the heater tube length produced a small increase in the indicated thermal efficiency with the heater-bypass concept.

Sullivan, T. J.↗
DOE OSTI2024

Catalytic enhancement in the performance of the microscopic two-stroke heat engine

Here we consider a model of heat engine operating in the microscopic regime: the two-stroke engine. It produces work and exchanges heat in two discrete strokes that are separated in time. The working body of the engine consists of two d-level systems initialized in thermal states at two distinct temperatures. Additionally, an auxiliary nonequilibrium system called catalyst may be incorporated with the working body of the engine, provided the state of the catalyst remains unchanged after the completion of a thermodynamic cycle. This ensures that the work produced by the engine arises solely from the temperature difference. Upon establishing the rigorous thermodynamic framework, we characterize twofold improvement stemming from the inclusion of a catalyst. Firstly, we prove that in the noncatalytic scenario, the optimal efficiency of the two-stroke heat engine with a working body composed of two-level systems is given by the Otto efficiency, which can be surpassed by incorporating a catalyst with the working body. Secondly, we show that incorporating a catalyst allows the engine to operate in frequency and temperature regimes that are not accessible for noncatalytic two-stroke engines. We conclude with a general conjecture about the advantage brought by a catalyst: including the catalyst with the working body always allows to improve efficiency over the noncatalytic scenario for any microscopic two-stroke heat engines. We prove this conjecture for two-stroke engines where the working body is composed of two d-level systems initialized in thermal states at two distinct temperatures, as long as the final joint state leading to optimal efficiency in the noncatalytic scenario is not a product state, or at least one of the d-level system is not thermal.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
NASA NTRS1988

T55-L-712 turbine engine compressor housing refurbishment-plasma spray project

A study was conducted to assess the feasibility of reclaiming T55-L-712 turbine engine compressor housings with an 88 wt percent aluminum to 12 wt percent silicon alloy applied by a plasma spray process. Tensile strength testing was conducted on as-sprayed and thermally cycled test specimens which were plasma sprayed with 0.020 to 0.100 in. coating thicknesses. Satisfactory tensile strength values were observed in the as-sprayed tensile specimens. There was essentially no decrease in tensile strength after thermally cycling the tensile specimens. Furthermore, compressor housings were plasma sprayed and thermally cycled in a 150-hr engine test and a 200-hr actual flight test during which the turbine engine was operated at a variety of loads, speeds and torques. The plasma sprayed coating system showed no evidence of degradation or delamination from the compressor housings. As a result of these tests, a procedure was designed and developed for the application of an aluminum-silicon alloy in order to reclaim T55-L-712 turbine engine compressor housings.

Leissler, George W.↗
NASA NTRS1995

RTE: A computer code for Rocket Thermal Evaluation

The numerical model for a rocket thermal analysis code (RTE) is discussed. RTE is a comprehensive thermal analysis code for thermal analysis of regeneratively cooled rocket engines. The input to the code consists of the composition of fuel/oxidant mixture and flow rates, chamber pressure, coolant temperature and pressure. dimensions of the engine, materials and the number of nodes in different parts of the engine. The code allows for temperature variation in axial, radial and circumferential directions. By implementing an iterative scheme, it provides nodal temperature distribution, rates of heat transfer, hot gas and coolant thermal and transport properties. The fuel/oxidant mixture ratio can be varied along the thrust chamber. This feature allows the user to incorporate a non-equilibrium model or an energy release model for the hot-gas-side. The user has the option of bypassing the hot-gas-side calculations and directly inputting the gas-side fluxes. This feature is used to link RTE to a boundary layer module for the hot-gas-side heat flux calculations.

Naraghi, Mohammad H. N.↗
NASA NTRS1980

Similarity tests of turbine vanes, effects of ceramic thermal barrier coatings

The role of material thermal conductivity was analyzed for its effect on the thermal performance of air-cooled gas turbine components coated with a ceramic thermal barrier material when tested at reduced temperatures and pressures. It is shown that the thermal performance can be evaluated reliably at reduced gas and coolant conditions; however, thermal conductivity corrections are required for the data at reduced conditions. Corrections for a ceramic thermal barrier coated vane are significantly different than for an uncoated vane. Comparison of uncorrected test data, therefore, would show erroneously that the thermal barrier coating was ineffective. When thermal conductivity corrections are applied to the test data these data are then shown to be representative of engine data and also show that the thermal barrier coating increases the vane cooling effectiveness by 12.5 percent.

Gladden, H. J.↗
NASA NTRS1980

Similarity tests of turbine vanes - Effects of ceramic thermal barrier coatings

The role of material thermal conductivity was analyzed for its effect on the thermal performance of air-cooled gas turbine components coated with a ceramic thermal barrier material when tested at reduced temperatures and pressures. This study shows that the thermal performance can be evaluated reliably at reduced gas and coolant conditions. However, thermal conductivity corrections are required for the data at reduced conditions. These corrections for a ceramic thermal barrier coated vane are significantly different than the corrections for an uncoated vane. Comparison of uncorrected test data, therefore, would show erroneously that the thermal barrier coating was ineffective. When thermal conductivity corrections are applied to the test data these data are then shown to be representative of engine data and also show that the thermal barrier coating increases the vane cooling effectiveness by 12.5 precent.

Gladden, H. J.↗
NASA NTRS1989

Probabilistic modeling for simulation of aerodynamic uncertainties in propulsion systems

The numerical simulation of the probabilistic aerothermodynamic response of propulsion system components to randomness in their environment was explored. The reusable rocket engine turbopumps were selected as an example because of the severe cryogenic environment in which they operate. The thermal and combustion instabilities, coupled with the engine thrust requirements from start up to shut down, lead to randomness in the flow variables and uncertainties in the aerodynamic loading. The probabilistic modeling of the turbopumps aerodynamic response was accomplished using the panel method coupled with fast probability integration methods. The aerodynamic response in the form of probabilistic rotor blades and splitter loading were predicted and the results presented for specified flow coefficient and rotor preswirl variance. Possible future applications of the aerothermodynamic probabilistic modeling in engine transient simulation, condition monitoring and engine life prediction are briefly discussed.

Hamed, Awatef↗
NASA NTRS1990

Probabilistic modeling for simulation of aerodynamic uncertainties in propulsion systems

The numerical simulation of the probabilistic aerothermodynamic response of propulsion system components to randomness in their environment was explored. The reusable rocket engine turbopumps were selected as an example because of the severe cryogenic environment in which they operate. The thermal and combustion instabilities, coupled with the engine thrust requirements from start up to shut down, lead to randomness in the flow variables and uncertainties in the aerosynamic loading. The probabilistic modeling of the turbopumps' aerodynamic response was accomplished using the panel method coupled with fast probability integration methods. The aerodynamic response in the form of probabilistic rotor blades and splitter loading were predicted and the results presented for specified flow coefficient and rotor preswirl variance. Possible future applications of the aerothermodynamic probabilistic modeling in engine transient simulation, condition monitoring, and engine life prediction are briefly discussed.

Hamed, A.↗
NASA NTRS1995

The failure analysis, redesign, and final preparation of the Brilliant Eyes Thermal Storage Unit for flight testing

This paper describes the engineering thought process behind the failure analysis, redesign, and rework of the flight hardware for the Brilliant Eyes Thermal Storage Unit (BETSU) experiment. This experiment was designed to study the zero-g performance of 2-methylpentane as a suitable phase change material. This hydrocarbon served as the cryogenic storage medium for the BETSU experiment which was flown 04 Mar 94 on board Shuttle STS-62. Ground testing had indicated satisfactory performance of the BETSU at the 120 Kelvin design temperature. However, questions remained as to the micro-gravity performance of this unit; potential deviations in ground (1 g) versus space flight (0 g) performance, and how the unit would operate in a realistic space environment undergoing cyclical operation. The preparations and rework performed on the BETSU unit, which failed initial flight qualification, give insight and lessons learned to successfully develop and qualify a space flight experiment.

Lamkin, T.↗
NASA NTRSDate not supplied

Vacuum Operation of Consolidated Heat Pipe (CHP) for Fission Surface Power

Consolidated Heat Pipe (CHP) is a new technology that enables direct thermal power delivery to the hot-end of a Stirling engine using a heat pipe--a two phase passive heat transfer device. CHP was developed after the efforts of the Kilopower Using Stirling TechnologY (KRUSTY) test where a heat pipe was used to deliver thermal power from a fission-basedreactor to a Stirling engine to produce 1 kWe of useable electrical power in 2018. Large thermal losses were noted during the KRUSTY test where a temperature drop of 145 °C was measured between the heat pipe’s condenser and the engine’s hot-end. The Consolidated Heat Pipe was designed to address and mitigate this temperature loss. CHP was designed, built and tested at the Glenn Research Center (GRC). The initial test was performed in ambient air conditions, and the results were presented in “Consolidation of a Sodium Heat Pipe and Stirling Engine for Fission Surface Power” at the Thermal Fluids Analysis Workshop in 2023. Researchers at GRC have tested this technology again in a vacuum environment in 2024. Results show that the heat pipe and the hot-end of the Stirling engine are isothermal with a minimal temperature differential of approximately 2.5 °C in varying operational states. The Consolidated Heat Pipe technology has proven to be an efficient way of delivering thermal power directly to Stirling engines.

Greeta J Thaikattil↗
NASA NTRS1991

Thermal analysis of conceptual designs for GPHS/FPSE power systems of 250 We and 500 We

Thermal analyses were performed for two distinct configurations of a proposed space nuclear power system which combines General Purpose Heat Source (GPHS) modules with the state of the art Free-Piston Stirling Engines (FPSEs). The two configurations correspond to systems with power levels of 250 and 500 W(sub e). The 250 W(sub e) GPHS/FPSE power system utilizes four GPHS modules and one FPSE, and the 500 W(sub e) contains eight GPHS modules and two FPSEs. The configurations of the systems and the bases for selecting the configurations are described. Brief introductory sections are included to describe the GPHS modules and free piston Stirling engines. The primary focus of the thermal analyses is on the temperature of the iridium fuel clad within the GPHS modules. A design goal temperature of 1573 K was selected as the upper limit for the fuel clad during normal operating conditions. The basis for selecting this temperature limit is discussed in detail. Results obtained from thermal analysis of the 250 W(sub e) GPHS/FPSE power system indicate fuel clad temperatures which slightly exceed the design goal temperature of 1573 K. The results are considered favorable due to the numerous conservative assumptions used in developing the thermal model and performing the thermal analysis. To demonstrate the effects of the conservatism, a brief sensitivity analysis is performed in which a few of the key system parameters are varied to determine their effect on the fuel clad temperatures. It is concluded that thermal analysis of a more detailed thermal model would be expected to yield fuel clad temperatures below the design foal temperature limiy 1573 K.

Mccomas, Thomas J.↗
NASA NTRS1976

Development of an inflatable radiator system

Conceptual designs of an inflatable radiator system developed for supplying short duration supplementary cooling of space vehicles are described along with parametric trade studies, materials evaluation/selection studies, thermal and structural analyses, and numerous element tests. Fabrication techniques developed in constructing the engineering models and performance data from the model thermal vacuum tests are included. Application of these data to refining the designs of the flight articles and to constructing a full scale prototype radiator is discussed.

Leach, J. W.↗
NASA NTRS1977

Development of a plasma sprayed ceramic gas path seal for high pressure turbine applications

The plasma sprayed graded layered yittria stabilized zirconia (ZrO2)/metal(CoCrAlY) seal system for gas turbine blade tip applications up to 1589 K (2400 F) seal temperatures was studied. Abradability, erosion, and thermal fatigue characteristics of the graded layered system were evaluated by rig tests. Satisfactory abradability and erosion resistance was demonstrated. Encouraging thermal fatigue tolerance was shown. Initial properties for the plasma sprayed materials in the graded, layered seal system was obtained, and thermal stress analyses were performed. Sprayed residual stresses were determined. Thermal stability of the sprayed layer materials was evaluated at estimated maximum operating temperatures in each layer. Anisotropic behavior in the layer thickness direction was demonstrated by all layers. Residual stresses and thermal stability effects were not included in the analyses. Analytical results correlated reasonably well with results of the thermal fatigue tests. Analytical application of the seal system to a typical gas turbine engine application predicted performance similar to rig specimen thermal fatigue performance. A model for predicting crack propagation in the sprayed ZrO2/CoCrAlY seal system was proposed, and recommendations for improving thermal fatigue resistance were made. Seal system layer thicknesses were analytically optimized to minimize thermal stresses in the abradability specimen during thermal fatigue testing. Rig tests on the optimized seal configuration demonstrated some improvement in thermal fatigue characteristics.

Shiembob, L. T.↗
NASA NTRS1988

Vacuum application of thermal barrier plasma coatings

Coatings are presently applied to Space Shuttle Main Engine (SSME) turbine blades for protection against the harsh environment realized in the engine during lift off-to-orbit. High performance nickel, chromium, aluminum, and yttrium (NiCrAlY) alloy coatings, which are applied by atmospheric plasma spraying, crack and spall off because of the severe thermal shock experienced during start-up and shut-down of the engine. Ceramic coatings of yttria stabilized zirconia (ZrO2-Y2O3) were applied initially as a thermal barrier over coating to the NiCrAlY but were removed because of even greater spalling. Utilizing a vacuum plasma spraying process, bond coatings of NiCrAlY were applied in a low pressure atmosphere of argon/helium, producing significantly improved coating-to-blade bonding. The improved coatings showed no spalling after 40 MSFC burner rig thermal shock cycles, cycling between 1700 and -423 F. The current atmospheric plasma NiCrAlY coatings spalled during 25 test cycles. Subsequently, a process was developed for applying a durable thermal barrier coating of ZrO2-Y2O3 to the turbine blades of first stage high-pressure fuel turbopumps utilizing the vacuum plasma process. The improved thermal barrier coating has successfully passed 40 burner rig thermal shock cycles without spalling. Hot firing in an SSME turbine engine is scheduled for the blades. Tooling was installed in preparation for vacuum plasma spray coating other SSME hardware, e.g., the titanium main fuel valve housing (MFVH) and the fuel turbopump nozzle/stator.

Holmes, R. R.↗