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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Radial and circumferential flow surveys at the inlet and exit of the Space Shuttle Main Engine High Pressure Fuel Turbine Model

The main objective of this test was to obtain detailed radial and circumferential flow surveys at the inlet and exit of the SSME High Pressure Fuel Turbine model using three-hole cobra probes, hot-film probes, and a laser velocimeter. The test was designed to meet several objectives. First, the techniques for making laser velocimeter, hot-film probe, and cobra probe measurements in turbine flows were developed and demonstrated. The ability to use the cobra probes to obtain static pressure and, therefore, velocity had to be verified; insertion techniques had to be established for the fragile hot-film probes; and a seeding method had to be established for the laser velocimetry. Once the measurement techniques were established, turbine inlet and exit velocity profiles, temperature profiles, pressure profiles, turbulence intensities, and boundary layer thicknesses were measured at the turbine design point. The blockage effect due to the model inlet and exit total pressure and total temperature rakes on the turbine performance was also studied. A small range of off-design points were run to obtain the profiles and to verify the rake blockage effects off-design. Finally, a range of different Reynolds numbers were run to study the effect of Reynolds number on the various measurements.

Hudson, S. T.↗
DOE OSTI2020

Attosecond synchronization of extreme ultraviolet high harmonics from crystals

The interaction of strong near-infrared (NIR) laser pulses with wide-bandgap dielectrics produces high harmonics in the extreme ultraviolet (XUV) wavelength range. These observations have opened up the possibility of attosecond metrology in solids, which would benefit from a precise measurement of the emission times of individual harmonics with respect to the NIR laser field. In this work, we show that, when high-harmonics are detected from the input surface of a magnesium oxide crystal, a bichromatic probing of the XUV emission shows a clear synchronization largely consistent with a semiclassical model of electron–hole recollisions in bulk solids. On the other hand, the bichromatic spectrogram of harmonics originating from the exit surface of the 200 μm-thick crystal is strongly modified, indicating the influence of laser field distortions during propagation. Our tracking of sub-cycle electron and hole re-collisions at XUV energies is relevant to the development of solid-state sources of attosecond pulses.

74 ATOMIC AND MOLECULAR PHYSICS↗
NASA NTRS1974

Two-dimensional cold-air cascade study of a film-cooled turbine stator blade. 1: Experimental results of pressure-surface film cooling tests

The effect of film coolant ejection from the pressure side of a stator blade was determined in a two-dimensional cascade. Stator exit surveys were made for each of six rows of coolant holes. Successive multirow tests were made with two, three, four, five, and six rows of coolant holes open. The results of the multirow tests are compared with the predicted multirow performance obtained by adding the single-row data. Results are presented in terms of stator primary-air efficiency as a function of coolant fraction.

Moffitt, T. P.↗
NASA NTRS1976

Flux-flow voltages during guided flux collapse from hollow superconducting cylinders

Voltages across diametrically opposite contact pairs on the outer surface of superconducting Nb tubes are found to depend dramatically upon the spatial configuration of the voltage-measuring leads relative to the positions of the moving magnetic flux lines. Experiments have been conducted to study these voltages for different wall thicknesses and for a variety of arrangements of the leads when flux in the hole and the wall of the Nb tube is made to exit or enter, completely or partially, by applying heat at a narrow strip along its length. Using the critical-state concept, a model for the change of flux and the resulting electric fields in the Nb tube on application of a heat pulse is presented. The resulting time-dependent and time-integrated voltages are calculated in excellent quantitative agreement with the experimental results. These results show that the flux-flow voltages across two contact points on a superconductor arise from the generation of an induced electric field over a chosen path in the superconductor between the contact points and from a change of magnetic flux through the surface bounded by the measuring leads and the chosen path in the superconductor.

Khanna, S. M.↗
NASA NTRS1992

Experimental and computational investigation of the NASA Low-Speed Centrifugal Compressor flow field

An experimental and computational investigation of the NASA Low-Speed Centrifugal Compressor (LSCC) flow field has been conducted using laser anemometry and Dawes' 3D viscous code. The experimental configuration consists of a backswept impeller followed by a vaneless diffuser. Measurements of the three-dimensional velocity field were acquired at several measurement planes through the compressor. The measurements describe both the throughflow and secondary velocity field along each measurement plane. In several cases the measurements provide details of the flow within the blade boundary layers. Insight into the complex flow physics within centrifugal compressors is provided by the computational analysis, and assessment of the CFD predictions is provided by comparison with the measurements. Five-hole probe and hot-wire surveys at the inlet and exit to the rotor as well as surface flow visualization along the impeller blade surfaces provide independent confirmation of the laser measurement technique.

Hathaway, M. D.↗
NASA NTRS1993

Experimental and computational investigation of the NASA low-speed centrifugal compressor flow field

An experimental and computational investigation of the NASA Lewis Research Center's low-speed centrifugal compressor (LSCC) flow field was conducted using laser anemometry and Dawes' three-dimensional viscous code. The experimental configuration consisted of a backswept impeller followed by a vaneless diffuser. Measurements of the three-dimensional velocity field were acquired at several measurement planes through the compressor. The measurements describe both the throughflow and secondary velocity field along each measurement plane. In several cases the measurements provide details of the flow within the blade boundary layers. Insight into the complex flow physics within centrifugal compressors is provided by the computational fluid dynamics analysis (CFD), and assessment of the CFD predictions is provided by comparison with the measurements. Five-hole probe and hot-wire surveys at the inlet and exit to the impeller as well as surface flow visualization along the impeller blade surfaces provided independent confirmation of the laser measurement technique. The results clearly document the development of the throughflow velocity wake that is characteristic of unshrouded centrifugal compressors.

Hathaway, Michael D.↗
NASA NTRS2001

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/Lean Combustor for High Speed Civil Transport Aircraft Engines

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of a HSCT engine cycle, Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NOx. The spatial profiles of NOx and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NOx. Based on this study, it was also concluded that (1) While NOx formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NOx exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) A RQL combustor can achieve the emissions goal of EINOX = 5 at the Supersonic Cruise operating condition for a HSCT engine.

Rosfjord, T. J.↗
NASA NTRS2001

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/ Lean Combustor for High Speed Civil Transport Aircraft Engines

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of an HSCT engine cycle. Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NO(x). The spatial profiles of NO(x) and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NO(x). Based on this study, it was also concluded that: (1) While NO(x) formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NO(x) exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) An RQL combustor can achieve the emissions goal of EINO(x) = 5 at the Supersonic Cruise operating condition for an HSCT engine.

Tacina, Robert R.↗
NASA NTRSDate not supplied

Effects of Trailing Edge Thickness and Blade Loading Distribution on the Aerodynamic Performance of Simulated CMC Turbine Blades

The aerodynamic performance of three blade sets that represent the geometric manufacturing constraints of Ceramic Matrix Composite (CMC) blades was measured experimentally in a large-scale transonic turbine blade cascade. The trailing edge thicknesses of CMC blades are anticipated to be significantly larger than those of current state-of-the-art metallic blades. The blades tested in the current study had trailing edge thicknesses of 5%, 7%, and 9% relative to the blade axial chord. The three blade sets were designed with matching throat dimensions, so the blade loading distributions were varied to retain similar overall loading levels. Data were acquired at four Reynolds numbers, covering a factor of six range. All data were acquired at the design isentropic exit Mach number of 0.74. Measurements include blade loading and five-hole probe surveys at two downstream stations. The effects of inlet turbulence intensity were also quantified. Total pressure loss data were integrated to determine overall loss levels for each of the three measured blade passages. Excellent periodicity was noted. For low inlet turbulence levels, losses were surprisingly lower for the thickest trailing edge at low Reynolds numbers, but were highest at the maximum Reynolds number. In general, losses were found to scale well with Reynolds number, although front loading was found to significantly reduce the sensitivity of loss to Reynolds number. For high inlet turbulence intensity, losses were found to scale with trailing edge thickness as expected, and the Reynolds number sensitivity was reduced for all three blade sets. Loss levels at the highest Reynolds number were comparable at low and high inlet turbulence intensity levels.

Paul W Giel↗
NASA NTRSDate not supplied

Effects of Trailing Edge Thickness and Blade Loading Distribution on the Aerodynamic Performance of Simulated CMC Turbine Blades

The aerodynamic performance of three blade sets that represent the geometric manufacturing constraints of Ceramic Matrix Composite (CMC) blades was measured experimentally in a large-scale transonic turbine blade cascade. The trailing edge thicknesses of CMC blades are anticipated to be significantly larger than those of current state-of-the-art metallic blades. The blades tested in the current study had trailing edge thicknesses of 5%, 7%, and 9% relative to the blade axial chord. The three blade sets were designed with matching throat dimensions, so the blade loading distributions were varied to retain similar overall loading levels. Data were acquired at four Reynolds numbers, covering a factor of six range. All data were acquired at the design isentropic exit Mach number of 0.74. Measurements include blade loading and five-hole probe surveys at two downstream stations. The effects of inlet turbulence intensity were also quantified. Total pressure loss data were integrated to determine overall loss levels for each of the three measured blade passages. Excellent periodicity was noted. For low inlet turbulence levels, losses were surprisingly lower for the thickest trailing edge at low Reynolds numbers, but were highest at the maximum Reynolds number. In general, losses were found to scale well with Reynolds number, although front loading was found to significantly reduce the sensitivity of loss to Reynolds number. For high inlet turbulence intensity, losses were found to scale with trailing edge thickness as expected, and the Reynolds number sensitivity was reduced for all three blade sets. Loss levels at the highest Reynolds number were comparable at low and high inlet turbulence intensity levels.

Turbine↗
NASA NTRS2005

Advanced Turbulence Modeling Concepts

The ZCET program developed at NASA Glenn Research Center is to study hydrogen/air injection concepts for aircraft gas turbine engines that meet conventional gas turbine performance levels and provide low levels of harmful NOx emissions. A CFD study for ZCET program has been successfully carried out. It uses the most recently enhanced National combustion code (NCC) to perform CFD simulations for two configurations of hydrogen fuel injectors (GRC- and Sandia-injector). The results can be used to assist experimental studies to provide quick mixing, low emission and high performance fuel injector designs. The work started with the configuration of the single-hole injector. The computational models were taken from the experimental designs. For example, the GRC single-hole injector consists of one air tube (0.78 inches long and 0.265 inches in diameter) and two hydrogen tubes (0.3 inches long and 0.0226 inches in diameter opposed at 180 degree). The hydrogen tubes are located 0.3 inches upstream from the exit of the air element (the inlet location for the combustor). To do the simulation, the single-hole injector is connected to a combustor model (8.16 inches long and 0.5 inches in diameter). The inlet conditions for air and hydrogen elements are defined according to actual experimental designs. Two crossing jets of hydrogen/air are simulated in detail in the injector. The cold flow, reacting flow, flame temperature, combustor pressure and possible flashback phenomena are studied. Two grid resolutions of the numerical model have been adopted. The first computational grid contains 0.52 million elements, the second one contains over 1.3 million elements. The CFD results have shown only about 5% difference between the two grid resolutions. Therefore, the CFD result obtained from the model of 1.3-million grid resolution can be considered as a grid independent numerical solution. Turbulence models built in NCC are consolidated and well tested. They can handle both coarse and fine grids near the wall. They can model the effect of anisotropy of turbulent stresses and the effect of swirling. The chemical reactions of Magnusson model and ILDM method were both used in this study.

Shih, Tsan-Hsing↗
NASA NTRS1995

Study of compressible flow through a rectangular-to-semiannular transition duct

Detailed flow field measurements are presented for compressible flow through a diffusing rectangular-to-semiannular transition duct. Comparisons are made with published computational results for flow through the duct. Three-dimensional velocity vectors and total pressures were measured at the exit plane of the diffuser model. The inlet flow was also measured. These measurements are made using calibrated five-hole probes. Surface oil flow visualization and surface static pressure data were also taken. The study was conducted with an inlet Mach number of 0.786. The diffuser Reynolds based on the inlet centerline velocity and the exit diameter of the diffuser was 3,200,000. Comparison of the measured data with previously published computational results are made. Data demonstrating the ability of vortex generators to reduce flow separation and circumferential distortion is also presented.

Foster, Jeffry↗
NASA NTRS1992

Investigation of three-dimensional flow field in a turbine including rotor/stator interaction. II - Three-dimensional flow field at the exit of the nozzle

The nozzle exit flowfield was measured at two axial locations with a miniature five-hole probe. Measurements were taken from hub-to-tip, blade-to-blade at 21 radial locations and at two axial locations downstream of the nozzle trailing edge to resolve the flowfield accurately including the nozzle wake, secondary flow region, horseshoe vortex and losses. All three components of the velocity, stagnation pressure, static pressure, and pitch and yaw angles have been resolved very accurately. The wake data seems to indicate that the decay of the wake is faster than the wake of an isolated nozzle row. The cause of this is attributed to the presence of the rotor downstream. A distinct vortex core has been observed near the tip. The indications are that the horseshoe vortex and the passage vortex have merged to produce a single loss core region. Roughly a third of the blade height passage near the tip and a third of the blade height near the hub is dominated by secondary flow, passage vortex and the horseshoe vortex phenomena. Only the middle third of the nozzle behaves as per design. These and other data are presented, interpreted and synthesized to understand the nozzle flowfield.

Zaccaria, M.↗
NASA NTRS1982

Flow distributions and discharge coefficient effects for jet array impingement with initial crossflow

To model the impingement cooled mid-chord region of gas turbine airfoils in cases where an initial crossflow is present, the paper presents experimentally determined flow distributions for jet arrays with ten spanwise rows of holes, which range from uniform to highly nonuniform. The jet flow after impingement is constrained to exit in a single direction along the channel formed by the jet orifice plate and the impingement surface. The streamwise distributions and crossflow velocities are presented for ratios of the initial crossflow rate to the total jet flow rate ranging from zero to unity. For crossflow to jet velocity ratios greater than a value somewhat less than unity, jet orifice discharge coefficients do not remain constant, but decrease significantly, showing a secondary dependence on z/d, where z is the channel height and d is the jet hole diameter.

Florschuetz, L. W.↗
NASA NTRS2015

Characterization of the Scale Model Acoustic Test Overpressure Environment using Computational Fluid Dynamics

The Scale Model Acoustic Test (SMAT) is a 5% scale test of the Space Launch System (SLS), which is currently being designed at Marshall Space Flight Center (MSFC). The purpose of this test is to characterize and understand a variety of acoustic phenomena that occur during the early portions of lift off, one being the overpressure environment that develops shortly after booster ignition. The pressure waves that propagate from the mobile launcher (ML) exhaust hole are defined as the ignition overpressure (IOP), while the portion of the pressure waves that exit the duct or trench are the duct overpressure (DOP). Distinguishing the IOP and DOP in scale model test data has been difficult in past experiences and in early SMAT results, due to the effects of scaling the geometry. The speed of sound of the air and combustion gas constituents is not scaled, and therefore the SMAT pressure waves propagate at approximately the same speed as occurs in full scale. However, the SMAT geometry is twenty times smaller, allowing the pressure waves to move down the exhaust hole, through the trench and duct, and impact the vehicle model much faster than occurs at full scale. The DOP waves impact portions of the vehicle at the same time as the IOP waves, making it difficult to distinguish the different waves and fully understand the data. To better understand the SMAT data, a computational fluid dynamics (CFD) analysis was performed with a fictitious geometry that isolates the IOP and DOP. The upper and lower portions of the domain were segregated to accomplish the isolation in such a way that the flow physics were not significantly altered. The Loci/CHEM CFD software program was used to perform this analysis.

Nielsen, Tanner↗
NASA NTRS1983

The performance of an annular vane swirler

In connection with the desirability of optimizing the design of a gas turbine combustion chamber, there exists a need for a more complete understanding of the fluid dynamics of the flow in such chambers. In order to satisfy this need, experimental and theoretical research is being conducted with the objective to study two-dimensional axisymmetric geometries under low speed, nonreacting, turbulent, swirling flow conditions. The flow enters the test section and proceeds into a larger chamber. Inlet swirl vanes are adjustable to a variety of vane angles. The present investigation concentrates on the time-mean flow characteristics which are generated by the upstream annular swirler. The investigation makes use of a five-hole pitot probe technique. A theoretical analysis of swirl numbers associated with several idealized exit velocity profiles is included, and values of the ratio of maximum swirl velocity to maximum axial velocity at different swirl strengths are given for each case.

Lilley, D. G.↗
DOE OSTI2024

Chaotic LLM billiards

Abstract We study null geodesics of the ten-dimensional LLM geometries. In particular, we show that there are a subset of these null geodesics that are confined to the LLM plane. The effective dynamics of these in-plane geodesics is that of a Hamiltonian system with two degrees of freedom (a phase space of dimension 4). We show that these are chaotic. In the two-coloring of the LLM plane, if they start in the empty region, they cannot penetrate the filled region and viceversa. The dynamical problem is therefore very similar to that of a billiards problem with fixed obstacles. We study to what extent LLM geometries with many droplets may be treated as an incipient black hole and draw analogies with the fuzzball proposal. We argue that for in-plane null geodesics deep in the interior of a region with a lot of droplets, in order to exit towards theAdSboundary they will need to undergo a process that resembles diffusion. This mechanism can account for signals getting lost in the putative black hole for a very long time.

Physics↗
NASA NTRS1976

Explosive Entrances

Explosive Technology, Inc. manufactured explosives first used by NASA to separate stages of the Gemini launch vehicle. When firemen need to get into a burning building or chop a hole to provide ventilation, axes can be devastatingly slow. Controlled explosives developed to separate manned upper stages of space rockets in case of mishap have been adapted to cutting emergency exits and demolishing unsafe buildings and bridges. Detonation cuts through thick steel girders or other materials more cleanly than torches or saws. This device can also cut emergency holes in airplanes and trains so surviving passengers can escape.

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