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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 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Calorimeter probes for measuring high thermal flux

The paper describes expendable, slug-type calorimeter probes developed for measuring high heat-flux levels of 10-30 kW/sq cm in electric-arc jet facilities. The probes are constructed with thin tungsten caps mounted on Teflon bodies; the temperature of the back surface of the tungsten cap is measured, and its rate of change gives the steady-state, absorbed heat flux as the calorimeter probe heats to destruction when inserted into the arc jet. It is concluded that the simple construction of these probes allows them to be expendable and heated to destruction to obtain a measurable temperature slope at high heating rates.

Russell, L. D.↗
NASA NTRS1979

Calorimeter probes for measuring high thermal flux

Expendable, slug-type calorimeter probes were developed for measuring high heat-flux levels of 10-30 kW/sq cm in electric-arc jet facilities. The probes were constructed with thin tungsten caps mounted on Teflon bodies. The temperature of the back surface of the tungsten cap is measured, and its time rate of change gives the steady-state absorbed heat flux as the calorimeter probe heats to destruction when inserted into the arc jet. Design, construction, test, and performance data are presented.

Russell, L. D.↗
NASA NTRS1997

Transpiring Cooling of a Scram-Jet Engine Combustion Chamber

The peak cold-wall heating rate generated in a combustion chamber of a scram-jet engine can exceed 2000 Btu/sq ft sec (approx. 2344 W/sq cm). Therefore, a very effective heat dissipation mechanism is required to sustain such a high heating load. This research focused on the transpiration cooling mechanism that appears to be a promising approach to remove a large amount of heat from the engine wall. The transpiration cooling mechanism has two aspects. First, initial computations suggest that there is a reduction, as much as 75%, in the heat flux incident on the combustion chamber wall due to the transpirant modifying the combustor boundary layer. Secondly, the heat reaching the combustor wall is removed from the structure in a very effective manner by the transpirant. It is the second of these two mechanisms that is investigated experimentally in the subject paper. A transpiration cooling experiment using a radiant heating method, that provided a heat flux as high as 200 Btu/sq ft sec ( approx. 234 W/sq cm) on the surface of a specimen, was performed. The experiment utilized an arc-lamp facility (60-kW radiant power output) to provide a uniform heat flux to a test specimen. For safety reasons, helium gas was used as the transpirant in the experiments. The specimens were 1.9-cm diameter sintered, powdered-stainless-steel tubes of various porosities and a 2.54cm square tube with perforated multi-layered walls. A 15-cm portion of each specimen was heated. The cooling effectivenes and efficiencies by transpiration for each specimen were obtained using the experimental results. During the testing, various test specimens displayed a choking phenomenon in which the transpirant flow was limited as the heat flux was increased. The paper includes a preliminary analysis of the transpiration cooling mechanism and a scaling conversion study that translates the results from helium tests into the case when a hydrogen medium is used.

Choi, Sang H.↗
NASA NTRS2019

Common Probe Design Study and Follow-On Activities

The Common Probe Study was funded by the NASA's Planetary Science Division in the Science Mission Directorate in 2018 to investigate the feasibility of a common aeroshell design for atmospheric probe missions at Venus, Jupiter, Saturn, Uranus, and Neptune. The study involved 4 NASA Centers: Ames Research Center, Goddard Space Flight Center, Langley Research Center, and the Jet Propulsion Laboratory. The common aeroshell design that was studied was a 400 kg, 1.5 m diameter, 45-degree sphere cone shape with a high density heatshield material (Heatshield for Extreme Entry Environments Technology, or HEEET) and a parachute system to extract the descent vehicle. This size of aeroshell could accommodate a descent vehicle of 0.75 m diameter, which could encompass both Tier 1 and Tier 2 science instruments at each of the 5 destinations. Study methodology: First, a notional payload of instruments for each destination was defined based on the top priority measurements indicated by the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified that bounded the EFPA range.Next, 3-DoF simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Conical ribbon parachutes were sized based on heatshield separation dynamics. Aero-heating correlations were used to generate stagnation point convective and radiative heat flux profiles. High fidelity thermal response models for various TPS materials were used to size stagnation point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts.Based on these analyses, we have found that the common design is applicable for atmospheric probe missions for 4 out of the 5 destinations. Because of the unique gravity well for Jupiter, the entry environments are more severe resulting in heat loads an order of magnitude higher than for the other destinations.The next step is to determine what follow-on activities NASA should engage in. A questionnaire for the atmospheric probe community has been developed, with a focus on what size of aeroshell should be further analyzed (smaller or same diameter), and what incentives would make using such an aeroshell, if assembled and available, desirable to mission proposers.Preliminary results from this questionnaire will be presented.

Hwang, Helen H.↗
NASA NTRS1986

Theoretical study of the electron temperature in the high-latitude ionosphere for solar maximum and winter conditions

The T(e) variation in the high-latitude ionosphere at altitudes between 120 and 800 km has been modeled for solar maximum, winter solstice, and strong magnetic activity conditions. The calculated electron temperatures are consistent with the plasma densities and ion temperatures computed from a time-dependent ionospheric model. Heating rates for both solar EUV and auroral precipitation were included. In general, the predicted UT variation of the electron temperature that results from the displacement between the magnetic and geographic poles is only a few hundred degrees. However, in sunlit trough regions, T(e) hot spots develop, and these hot spots show a marked UT variation, by as much as 2500 K. The dominant parameter controlling the T(e) variation above 200 km is the magnetospheric heat flux into the ionosphere, which is essentially unknown. For realistic values of the magnetospheric heat flux, the maximum electron temperature ranges from 5000 to 10,000 K at 800 km. A magnetospheric heat flux is particularly effective in enhancing trough electron temperatures. In general, the electron heat flux at high altitudes is uniquely related to the electron temperature and gradient, except on auroral field lines where thermoelectric heat flow is important.

Schunk, R. W.↗
NASA NTRS1986

Measurement of heat-transfer coefficients in shock wave-turbulent boundary layer interaction regions with a multi-layered thin film heat transfer gauge

A thin film heat transfer gauge is applied to the measurement of heat transfer coefficients in the interaction regions of incident shock waves and fully developed turbulent boundary layers. It was developed to measure heat flux with high spatial resolution and fast response for wind tunnels with long flow duration. To measure the heat transfer coefficients in the interaction region in detail, experiments were performed under the conditions of Mach number = 4, total pressure = 1.2 MPa, 0.59 to approximately 0.65. Reynolds number = 1.3 to approximately 1.5 x 10 to the 7th power and incident shock angles from 17.8 to 22.8 degrees. The results show that the heat transfer coefficient changes complicatedly in the interaction region. At the beginning the interaction region, the heat transfer coefficient decreases at first, reaches its minimum value at the point where the pressure begins to increase, and then increases sharply. When the boundary layer begins to separate, even a small separation bubble causes significant changes in the heat transfer coefficient, while the pressure does not show any changes which suggests that the boundary layer begins to separate.

Hayashi, M.↗
NASA NTRS1981

Satellite estimation of the surface energy balance, moisture availability and thermal inertia

A method for inferring the distribution of surface heat and evaporative fluxes and the ground moisture availability and thermal inertia (ground conductive capacity) is used to analyze two urbanized areas, Los Angeles and St. Louis. The technique employs infrared satellite temperature measurements in conjunction with a one-dimensional boundary-layer model. Results show that there is a marked reduction of evaporation and moisture availability and a corresponding elevation of sensible heat flux over urbanized areas and over cropped areas with low vegetative cover. Conversely, low heat flux and high evaporation characterize vegetated and, especially, forested areas. Warm urban centers appear directly related to a reduction in vegetation, which normally allows for a greater fraction of available radiant energy to be converted into latent heat flux. The distribution of thermal inertia was surprisingly ill-defined and its variation between urban and rural areas was quite small. Thus, the increased heat storage within the urban fabric, which has been proposed as the underlying cause of the nocturnal heat island, may be caused mainly by enhanced daytime surface heating which occurs because of surface dryness, rather than by large spatial variations in the conductivity of the surface.

Carlson, T. N.↗
NASA NTRS2008

Geological Mapping of the North Polar Region of Venus (V-1 Snegurochka Planitia): Significant Problems and Comparisons to the Earth's Archean

The geological features, structures, thermal conditions, interpreted processes, and outstanding questions related to both the Earth's Archean and Venus share many similarities and we are using a problem-oriented approach to Venus mapping, guided by perspectives from the Archean record of the Earth, to gain new insight into both. The Earth's preserved and well-documented Archean record provides important insight into high heat-flux tectonic and magmatic environments and structures and Venus reveals the current configuration and recent geological record of analogous high-temperature environments unmodified by subsequent several billion years of segmentation and overprinting, as on Earth. We have problems on which progress might be made through comparison. Here we present the major goals of the geological mapping of the V-1 Snegurochka Planitia Quadrangle, and themes that could provide important insights into both planets:

Head, James W.↗
NASA NTRS2001

Thermal Conductivity of Ceramic Thermal Barrier and Environmental Barrier Coating Materials

Thermal barrier and environmental barrier coatings (TBC's and EBC's) have been developed to protect metallic and Si-based ceramic components in gas turbine engines from high temperature attack. Zirconia-yttria based oxides and (Ba,Sr)Al2Si2O8(BSAS)/mullite based silicates have been used as the coating materials. In this study, thermal conductivity values of zirconia-yttria- and BSAS/mullite-based coating materials were determined at high temperatures using a steady-state laser heat flux technique. During the laser conductivity test, the specimen surface was heated by delivering uniformly distributed heat flux from a high power laser. One-dimensional steady-state heating was achieved by using thin disk specimen configuration (25.4 mm diam and 2 to 4 mm thickness) and the appropriate backside air-cooling. The temperature gradient across the specimen thickness was carefully measured by two surface and backside pyrometers. The thermal conductivity values were thus determined as a function of temperature based on the 1-D heat transfer equation. The radiation heat loss and laser absorption corrections of the materials were considered in the conductivity measurements. The effects of specimen porosity and sintering on measured conductivity values were also evaluated.

Zhu, Dong-Ming↗
NASA NTRS2008

Geological Mapping of Fortuna Tessera (V-2): Venus and Earth's Archean Process Comparisons

The geological features, structures, thermal conditions, interpreted processes, and outstanding questions related to both the Earth's Archean and Venus share many similarities and we are using a problem-oriented approach to Venus mapping, guided by insight from the Archean record of the Earth, to gain new insight into the evolution of Venus and Earth's Archean. The Earth's preserved and well-documented Archean record provides important insight into high heat-flux tectonic and magmatic environments and structures and the surface of Venus reveals the current configuration and recent geological record of analogous high-temperature environments unmodified by subsequent several billion years of segmentation and overprinting, as on Earth. Elsewhere we have addressed the nature of the Earth's Archean, the similarities to and differences from Venus, and the specific Venus and Earth-Archean problems on which progress might be made through comparison. Here we present the major goals of the Venus-Archean comparison and show how preliminary mapping of the geology of the V-2 Fortuna Tessera quadrangle is providing insight on these problems. We have identified five key themes and questions common to both the Archean and Venus, the assessment of which could provide important new insights into the history and processes of both planets.

Head, James W.↗
NASA NTRSDate not supplied

Separation and Transition on the ROTEX-T Cone-Flare

As part of NATO STO AVT-346 “Predicting Hypersonic Boundary-Layer Transition on Complex Geometries,” coordinated experimental and computational studies were conducted on the ROTEX-T, a cone-flare geometry used in a successful flight-test experiment. At the as-flown conditions, a separation bubble existed at the compression corner. Separation, reattachment, and the multifaceted linear instability paths leading this bubble to transition to turbulence are challenging to predict, but have significant impact on surface pressure and heat flux. High-resolution background-oriented schlieren and infrared thermography measurements were made in the AFOSR–Notre Dame Large Mach-6 Quiet Tunnel at freestream unit Reynolds numbers from 5.8×10(exp 6) to 12.2×10(exp 6) m-1and nominally zero angle of attack. High-speed self-aligned focusing schlieren, infrared thermography, and focused laser differential interferometry measurements were made in the AFRL Mach-6 Ludwieg Tube from 2.2×10(exp 6) to 24.7×10(exp 6) m-1and nominally zero degrees angle of attack. The surface heat-flux and Stanton number distributions were computed. Separation and reattachment locations, as well as the flow state at each, were determined from the combination of surface and off-wall measurements. The convective and global boundary-layer instabilities of the axisymmetric laminar flow at the experimental conditions were investigated computationally. Amplification of Mack’s first and second modes were observed to have logarithmic amplification factors between 5 to 7.5 at the separation location, depending on conditions. The flow was found to be globally unstable to stationary three-dimensional disturbances concentrated in the reattachment region. Previous analysis of the ROTEX-T flight data had not assessed reattachment location or the flow state upon reattachment. Thanks to the insights gained from the coordinated, on- and off-wall ground-test measurements, these evaluations have now been made. The separation location indicated by laminar simulations is consistently numerically predicted to be upstream of the experimentally observed location for a transitional separation bubble. The cause of this difference is understood to lie in the steady-state and axisymmetric assumptions made by both solvers employed to compute the basic states analyzed, as flow topology considerations assert that unsteady two-dimensional or axisymmetric separation bubbles are structurally unstable and will become three-dimensional. Computed laminar heating rates prior to separation agreed well with experiment; transitional heating rates after reattachment were between laminar and turbulent computations.

CFD Validation↗
NASA NTRSDate not supplied

Separation and Transition on the ROTEX-T Cone-Flare

As part of NATO STO AVT-346 “Predicting Hypersonic Boundary-Layer Transition on Complex Geometries,” coordinated experimental and computational studies were conducted on the ROTEX-T, a cone-flare geometry used in a successful flight-test experiment. At the as-flown conditions, a separation bubble existed at the compression corner. Separation, reattachment, and the multifaceted linear instability paths leading this bubble to transition to turbulence are challenging to predict, but have significant impact on surface pressure and heat flux. High-resolution background-oriented schlieren and infrared thermography measurements were made in the AFOSR–Notre Dame Large Mach-6 Quiet Tunnel at freestream unit Reynolds numbers from 5.8×10(exp 6) to 12.2×10(exp 6) m-1and nominally zero angle of attack. High-speed self-aligned focusing schlieren, infrared thermography, and focused laser differential interferometry measurements were made in the AFRL Mach-6 Ludwieg Tube from 2.2×10(exp 6) to 24.7×10(exp 6) m-1and nominally zero degrees angle of attack. The surface heat-flux and Stanton number distributions were computed. Separation and reattachment locations, as well as the flow state at each, were determined from the combination of surface and off-wall measurements. The convective and global boundary-layer instabilities of the axisymmetric laminar flow at the experimental conditions were investigated computationally. Amplification of Mack’s first and second modes were observed to have logarithmic amplification factors between 5 to 7.5 at the separation location, depending on conditions. The flow was found to be globally unstable to stationary three-dimensional disturbances concentrated in the reattachment region. Previous analysis of the ROTEX-T flight data had not assessed reattachment location or the flow state upon reattachment. Thanks to the insights gained from the coordinated, on- and off-wall ground-test measurements, these evaluations have now been made. The separation location indicated by laminar simulations is consistently numerically predicted to be upstream of the experimentally observed location for a transitional separation bubble. The cause of this difference is understood to lie in the steady-state and axisymmetric assumptions made by both solvers employed to compute the basic states analyzed, as flow topology considerations assert that unsteady two-dimensional or axisymmetric separation bubbles are structurally unstable and will become three-dimensional. Computed laminar heating rates prior to separation agreed well with experiment; transitional heating rates after reattachment were between laminar and turbulent computations.

CFD Validation↗
NASA NTRS2019

Analysis of Base-Heating Environment During Ground Testing of a Lunar Lander Demonstrator

A three-dimensional, unstructured-grid, pressure-based, reacting flow, computational fluid dynamics (CFD) and heat transfer methodology was employed to study the base-heating environment of a lunar lander demonstrator, during terrestrial ground testing. Two base-heating environments were investigated: lunar lander demonstrator sitting on pad, and lunar lander demonstrator hovering at a distance above ground. Two unique and quite different base-flow physics are found for these two cases. In the case of demonstrator sitting on pad, the computed heat fluxes are high and the base flow is unsteady, caused by a Coanda effect precursor that makes the fountain jet oscillate about the center of the base. For the second case, due to the higher elevation of the demonstrator and the layout of the nozzles, the Coanda effect forces the fountain jet to be attached to two of the nozzle plumes, resulting in much lower computed base heat fluxes.

Wang, T. S.↗
NASA NTRS1985

Experimental Measurements and Analysis of Heat Transfer and Gas Dynamics in a Rotating Turbine Stage

The life-cycle fatigue of a particular component of a high-pressure turbine stage is influenced by the local heat flux and the resulting metal temperature excursions experienced by this component during normal operation. Flow through a turbine stage is seldom uniform and is often influenced by non-uniform incoming gas streams that can result in locally high heat fluxes. In the case of the rotor blades passing through the nozzle guide vane wakes and passages, these local heat rates translate into local temperature fluctuations that will likely affect the fatigue life of the blade. In order to better understand the influence of the heat load on the component life, one would like to have both the local steady-state heat flux and the high-frequency component of the heat flux.

Dunn, M. G.↗
NASA NTRS2010

High efficiency tantalum-based ceramic composite structures

Tantalum-based ceramics are suitable for use in thermal protection systems. These composite structures have high efficiency surfaces (low catalytic efficiency and high emittance), thereby reducing heat flux to a spacecraft during planetary re-entry. These ceramics contain tantalum disilicide, molybdenum disilicide and borosilicate glass. The components are milled, along with a processing aid, then applied to a surface of a porous substrate, such as a fibrous silica or carbon substrate. Following application, the coating is then sintered on the substrate. The composite structure is substantially impervious to hot gas penetration and capable of surviving high heat fluxes at temperatures approaching 3000.degree. F. and above.

Stewart, David A.↗
NASA NTRS1994

Simulation of a turbulent flame in a channel

The interaction between turbulent premixed flames and channel walls is studied. Combustion is represented by a simple irreversible reaction with a large activation temperature. Feedback to the flowfield is suppressed by invoking a constant density assumption. The effect of wall distance on local and global flame structure is investigated. Quenching distances and maximum wall heat fluxes computed in laminar cases are compared to DNS results. It is found that quenching distances decrease and maximum heat fluxes increase relative to laminar flame values. It is shown that these effects are due to large coherent structures which push flame elements towards to wall. The effect of wall strain is studied in flame-wall interaction in a stagnation line flow; this is used to explain the DNS results. It is also shown that 'remarkable' flame events are produced by interaction with a horseshoe vortex: burnt gases are pushed towards the wall at high speed and induce quenching and high wall heat fluxes while fresh gases are expelled from the wall region and form finger-like structures. Effects of the wall on flame surface density are investigated, and a simple model for flame-wall interaction is proposed; its predictions compare well with the DNS results.

Bruneaux, G.↗
NASA NTRSDate not supplied

Icing Physics Studies Using the 3D SIDRM Test Article: Ice Crystal Icing Analysis

In-flight icing is an important safety issue and is a factor that affects aircraft design and performance. Newer regulations are driving a need for improvements in airframe and engine icing simulation capability. Experimental data is required for the development of icing physics models and simulation validation. This paper presents the analysis of the ice crystal icing data subset from tests conducted in 2022 at the NASA Icing Research Tunnel that studied both supercooled liquid and ice-crystal icing. The test article that was utilized replicated 3D geometrical features of an inter-compressor duct and strut region of a turbofan engine. The surfaces of the Simulated Inter-compressor Duct Research Model (SIDRM) can be heated to simulate the warm surfaces of the turbofan inter-compressor duct. The test article is instrumented with pressure taps, heaters, heat flux gauges, and thermocouples, while a 3D laser scanner, cameras, and a weighing scale to measure ice mass were utilized to characterize the icing behavior. The aim of these tests was to generate ice accretions on the SIDRM test article under well-characterized icing conditions. To that end, ice crystal cloud characterization tests were conducted that provided cloud property values at the test section. The ice crystal accretion tests investigated how different test condition parameters impacted ice accretion size, location, and characteristics (such as physical attributes). A key finding was that a sufficiently high surface heat flux at the leading edge can provide a continuous supply of liquid melt to critical accretion zones downstream, resulting in elevated ice mass accretions. Another key finding was that warmer heater settings in critical accretion zones can suppress ice accretion. These findings are consistent with icing behavior observed in full-scale engine icing tests conducted previously at the NASA Glenn Research Center. Test runs investigating the contribution of liquid melt generated at the leading edge found that the melt can be transported downstream to critical accretion regions in two forms, via runback and splash. The form of melt supplied downstream affected accretion behavior and features. Sharkteeth-shaped ice accretions were generated at certain test conditions, a physical attribute that has been observed in full-scale engine icing tests. The ice accretion data will be used to develop and validate 3D computational engine icing tools, such as GlennICE, that predictively assesses the onset and growth of ice. One of the goals of the sponsoring NASA project is to develop simulation models and tools that can assist in the design and certification of engines for flight in icing conditions in a cost‑effective way.

ice crystal icing↗
NASA NTRSDate not supplied

Icing Physics Studies Using the 3D SIDRM Test Article: Ice Crystal Icing Analysis

In-flight icing is an important safety issue and is a factor that affects aircraft design and performance. Newer regulations are driving a need for improvements in airframe and engine icing simulation capability. Experimental data is required for the development of icing physics models and simulation validation. This paper presents the analysis of the ice crystal icing data subset from tests conducted in 2022 at the NASA Icing Research Tunnel that studied both supercooled liquid and ice-crystal icing. The test article that was utilized replicated 3D geometrical features of an inter-compressor duct and strut region of a turbofan engine. The surfaces of the Simulated Inter-compressor Duct Research Model (SIDRM) can be heated to simulate the warm surfaces of the turbofan inter-compressor duct. The test article is instrumented with pressure taps, heaters, heat flux gauges, and thermocouples, while a 3D laser scanner, cameras, and a weighing scale to measure ice mass were utilized to characterize the icing behavior. The aim of these tests was to generate ice accretions on the SIDRM test article under well-characterized icing conditions. To that end, ice crystal cloud characterization tests were conducted that provided cloud property values at the test section. The ice crystal accretion tests investigated how different test condition parameters impacted ice accretion size, location, and characteristics (such as physical attributes). A key finding was that a sufficiently high surface heat flux at the leading edge can provide a continuous supply of liquid melt to critical accretion zones downstream, resulting in elevated ice mass accretions. Another key finding was that warmer heater settings in critical accretion zones can suppress ice accretion. These findings are consistent with icing behavior observed in full-scale engine icing tests conducted previously at the NASA Glenn Research Center. Test runs investigating the contribution of liquid melt generated at the leading edge found that the melt can be transported downstream to critical accretion regions in two forms, via runback and splash. The form of melt supplied downstream affected accretion behavior and features. Sharkteeth-shaped ice accretions were generated at certain test conditions, a physical attribute that has been observed in full-scale engine icing tests. The ice accretion data will be used to develop and validate 3D computational engine icing tools, such as GlennICE, that predictively assesses the onset and growth of ice. One of the goals of the sponsoring NASA project is to develop simulation models and tools that can assist in the design and certification of engines for flight in icing conditions in a cost‑effective way.

ice crystal icing↗