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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

The prediction of nozzle performance and heat transfer in hydrogen/oxygen rocket engines with transpiration cooling, film cooling, and high area ratios

An advanced engineering computational model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multi-species, chemically reacting and diffusing Navier-Stokes equations are modelled, finite difference approach that is tailored to be conservative in an axisymmetric coordinate system for both the inviscid and viscous terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and transpiration cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent plug-and-spool rocket engine analysis cases performed. Further, the Soret term was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration cooled rocket engine.

Kacynski, Kenneth J.↗
NASA NTRS1947

Effect of Exhaust Pressure on the Cooling Characteristics of a Liquid-Cooled Engine

Data for a liquid-cooled engine with a displacement volume of 1710 cubic inches were analyzed to determine the effect of exhaust pressure on the engine cooling characteristics. The data covered a range of exhaust pressures from 7 to 62 inches of mercury absolute, inlet-manifold pressures from 30 to 50 inches of mercury absolute, engine speeds from 1600 to 3000 rpm, and fuel-air ratios from 0.063 to 0.100. The effect of exhaust pressure on engine cooling was satisfactorily incorporated in the NACA cooling-correlation method as a variation in effective gas temperature with exhaust pressure. Large variations of cylinder-head temperature with exhaust pressure were obtained for operation at constant charge flow. At a constant charge flow of 2 pounds per second (approximately 1000 bhp) and a fuel-air ratio of 0.085, an increase in exhaust pressure from 10 to 60 inches of mercury absolute resulted in an increase of 40 F in average cylinder-head temperature. For operation at constant engine speed and inlet-manifold pressure and variable exhaust pressure (variable charge flow), however, the effect of exhaust pressure on cylinder-head temperature is small. For example, at an inlet-manifold pressure of 40 inches of mercury absolute, an engine speed of 2400 rpm.- and a fuel-air ratio of 0.085, the average cylinder-head temperature was about the same at exhaust pressures of 10 and 60 inches of,mercury absolute; a rise and a subsequent decrease of about 70 occurred between these extremes.

Doyle, Ronald B.↗
NASA NTRS1950

Preliminary analysis of problem of determining experimental performance of air-cooled turbine II : methods for determining cooling-air-flow characteristics

In the determination of the performance of an air-cooled turbine, the cooling-air-flow characteristics between the root and the tip of the blades must be evaluated. The methods, which must be verified and the unknown functions evaluated, that are expected to permit the determination of pressure, temperature, and velocity through the blade cooling-air passages from specific investigation are presented.

FLOW, COMPRESSIBLE↗
NASA NTRS1971

Analytical prediction of the heat transfer from a blood vessel near the skin surface when cooled by a symmetrical cooling strip

An analytical method was developed to estimate the amount of heat extracted from an artery running close to the skin surface which is cooled in a symmetrical fashion by a cooling strip. The results indicate that the optimum width of a cooling strip is approximately three times the depth to the centerline of the artery. The heat extracted from an artery with such a strip is about 0.9 w/m-C which is too small to affect significantly the temperature of the blood flow through a main blood vessel, such as the carotid artery. The method is applicable to veins as well.

Chato, J. C.↗
NASA NTRS1990

Self-irradiated cooling condensations - The source of the optical line emission from cooling flows

The optical filaments seen at the centers of clusters of galaxies, which cannot be explained by standard photoionization or shock models, can be straightforwardly interpreted as self-irradiated cooling condensations. Fully developed (about 10,000 K) condensations embedded in thermally unstable 10 million K gas will be bathed in a powerful EUV/soft X-ray flux emanating from the surrounding condensing regions. The present models of the condensation/irradiation process give line ratios and luminosities similar to those observed. Because it is assumed that the condensations arise from linear perturbations, which grow significantly only when the local cooling time is less than the free-fall time, condensing regions of the type modeled here should arise only in cooling flows of central pressure less than about 1 billion ergs/cu cm and should extend only over the inner regions of the cluster.

Voit, G. Mark↗
NASA NTRS1936

Performance of Air-cooled Engine Cylinders Using Blower Cooling

An investigation was made to obtain information on the minimum quantity of air and power required to cool conventional air cooled cylinders at various operating conditions when using a blower. The results of these tests show that the minimum power required for satisfactory cooling with an overall blower efficiency of 100 percent varied from 2 to 6 percent of the engine power depending on the operating conditions. The shape of the jacket had a large effect on the cylinder temperatures. Increasing the air speed over the front of the cylinder by keeping the greater part of the circumference of the cylinder covered by the jacket reduced the temperatures over the entire cylinder.

Schey, Oscar W↗
NASA NTRS1940

Correlation of Cooling Data from an Air-Cooled Cylinder and Several Multicylinder Engines

The theory of engine-cylinder cooling developed in a previous report was further substantiated by data obtained on a cylinder from a Wright r-1820-g engine. Equations are presented for the average head and barrel temperatures of this cylinder as functions of the engine and the cooling conditions. These equations are utilized to calculate the variation in cylinder temperature with altitude for level flight and climb. A method is presented for correlating average head and barrel temperatures and temperatures at individual points on the head and the barrel obtained on the test stand and in flight. The method is applied to the correlation and the comparison of data obtained on a number of service engines. Data are presented showing the variation of cylinder temperature with time when the power and the cooling pressure drop are suddenly changed.

Pinkel, Benjamin↗
NASA NTRS1958

Effect of Chord Size on Weight and Cooling Characteristics of Air-Cooled Turbine Blades

An analysis has been made to determine the effect of chord size on the weight and cooling characteristics of shell-supported, air-cooled gas-turbine blades. In uncooled turbines with solid blades, the general practice has been to design turbines with high aspect ratio (small blade chord) to achieve substantial turbine weight reduction. With air-cooled blades, this study shows that turbine blade weight is affected to a much smaller degree by the size of the blade chord.

Esgar, Jack B↗
NASA NTRS1995

Cooling Panel Optimization for the Active Cooling System of a Hypersonic Aircraft

Optimization of cooling panels for an active cooling system of a hypersonic aircraft is explored. The flow passages are of rectangular cross section with one wall heated. An analytical fin-type model for incompressible flow in smooth-wall rectangular ducts with coupled wall conduction is proposed. Based on this model, the a flow rate of coolant to each design minimum mass flow rate or coolant for a single cooling panel is obtained by satisfying hydrodynamic, thermal, and Mach number constraints. Also, the sensitivity of the optimal mass flow rate of coolant to each design variable is investigated. In addition, numerical solutions for constant property flow in rectangular ducts, with one side rib-roughened and coupled wall conduction, are obtained using a k-epsilon and wall function turbulence model, these results are compared with predictions of the analytical model.

Youn, B.↗
NASA NTRS2002

Influence of Cooling Hole Geometry and Material Conductivity on the Thermal Response of Cooled Silicon Nitride Plate

To complement the effectiveness of ceramic materials and the applicability to turbine engine applications, a parametric study using the finite element method was carried out. This study conducted thorough analyses of a thermal-barrier-coated silicon nitride (Si3N4) plate specimen with cooling channels, where its thermal conductivity was verified in an attempt to minimize the thermal stresses and reach an optimal rate of stress. The thermal stress profile was generated for specimens with circular and square cooling channels. Lower stresses were reported for a higher magnitude of thermal conductivity and in particular for the circular cooling channel arrangement. Contour plots for the stresses and the temperature are presented and discussed.

Abdul-Aziz, Ali↗
NASA NTRS1951

Calculations of Laminar Heat Transfer Around Cylinders of Arbitrary Cross Section and Transpiration-Cooled Walls with Application to Turbine Blade Cooling

An approximate method for development of flow and thermal boundary layers in laminar regime on cylinders with arbitrary cross section and transpiration-cooled walls is obtained by use of Karman's integrated momentum equation and an analogous heat-flow equation. Incompressible flow with constant property values throughout boundary layer is assumed. Shape parameters for approximated velocity and temperature profiles and functions necessary for solution of boundary-layer equations are presented as charts, reducing calculations to a minimum. The method is applied to determine local heat-transfer coefficients and surface temperature-cooled turbine blades for a given flow rate. Coolant flow distributions necessary for maintaining uniform blade temperatures are also determined.

TURBINE COOLING↗
NASA NTRS2014

Development of the Glenn-HT Computer Code to Enable Time-Filtered Navier-Stokes (TFNS) Simulations and Application to Film Cooling on a Flat Plate Through Long Cooling Tubes

Computational fluid dynamics (CFD) analysis using Reynolds-averaged Navier-Stokes (RANS) formulation for turbomachinery-related flows has enabled improved engine component designs. RANS methodology has limitations which are related to its inability to accurately describe the spectrum of flow phenomena encountered in engines. Examples of flows that are difficult to compute accurately with RANS include phenomena such as laminarturbulent transition, turbulent mixing due to mixing of streams, and separated flows. Large eddy simulation (LES) can improve accuracy but at a considerably higher cost. In recent years, hybrid schemes which take advantage of both unsteady RANS and LES have been proposed. This study investigated an alternative scheme, the time-filtered Navier-Stokes (TFNS) method applied to compressible flows. The method developed by Shih and Liu was implemented in the Glenn-HT code and applied to film cooling flows. In this report the method and its implementation is briefly described. The film effectiveness results obtained for film cooling from a row of 30 holes with a pitch of 3.0 diameters emitting air at a nominal density ratio of unity and four blowing ratios of 0.5, 1.0, 1.5 and 2.0 are shown. Flow features under those conditions are also described.

infrared↗
NASA NTRS2014

Development of the Glenn-Heat-Transfer (Glenn-HT) Computer Code to Enable Time-Filtered Navier Stokes (TFNS) Simulations and Application to Film Cooling on a Flat Plate Through Long Cooling Tubes

Computational fluid dynamics (CFD) analysis using Reynolds-averaged Navier-Stokes (RANS) formulation for turbomachinery-related flows has enabled improved engine component designs. RANS methodology has limitations that are related to its inability to accurately describe the spectrum of flow phenomena encountered in engines. Examples of flows that are difficult to compute accurately with RANS include phenomena such as laminar/turbulent transition, turbulent mixing due to mixing of streams, and separated flows. Large eddy simulation (LES) can improve accuracy but at a considerably higher cost. In recent years, hybrid schemes that take advantage of both unsteady RANS and LES have been proposed. This study investigated an alternative scheme, the time-filtered Navier-Stokes (TFNS) method applied to compressible flows. The method developed by Shih and Liu was implemented in the Glenn-Heat-Transfer (Glenn-HT) code and applied to film-cooling flows. In this report the method and its implementation is briefly described. The film effectiveness results obtained for film cooling from a row of 30deg holes with a pitch of 3.0 diameters emitting air at a nominal density ratio of unity and two blowing ratios of 0.5 and 1.0 are shown. Flow features under those conditions are also described.

heat transfer↗
NASA NTRS1975

Two dimensional cold air cascade study of a film cooled turbine stator blade. 2: Experimental results of full film cooling tests

The effect of full film cooling on the performance of a turbine stator blade was studied in a two-dimensional cascade. The blade contained 12 rows of coolant holes, 6 rows on the suction surface and 6 on the pressure surface. Separate tests were first made with coolant ejection from each of the 12 rows. Then successive tests were made with various combinations of coolant rows open, including full film cooling. The principal results are presented in terms of primary-air efficiency as a function of coolant fraction. In addition, the efficiency results of the multirow tests are compared with the multi-row efficiency predicted by adding the single-row results.

Prust, H. W., Jr.↗
NASA NTRS1987

SiO cooling instability in the envelopes of cool giant stars

In the outer atmospheres of oxygen-rich giant and supergiant stars, the formation of the silicon monoxide molecule not only represents a first step toward the production of dust but may also produce an efficient radiative cooling agent. Molecular cooling occurs as a thermal instability in many cases with an onset at 2000 to 3500 K, rapidly reducing the temperature by as much as 1000 K. This process may occur in the upper photosphere-lower chromosphere of such stars and can have profound dynamical consequences. The process occurs far from LTE, on a time scale controlled by the reaction rates for SiO formation. SiO masers may be indicative of this process. The molecule CS which may play a similar role in carbon stars and serve as the basis for carbonaceous dust production.

Muchmore, David O.↗