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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1961

Supersonic Free-Flight Measurements of Heat Transfer and Transition on a 10 degree Cone having a Low Temperature Ratio

Heat-transfer coefficients in the form of Stanton number and boundary-layer transition data were obtained from a free-flight test of a 100-inch-long 10 deg. total-angle cone with a 1/16-inch tip radius which penetrated deep into the region of infinite stability of laminar boundary layer over a range of wall-to-local-stream temperature radius and for local Mach numbers from 1.8 to 3.5. Experimental heat-transfer coefficients, obtained at Reynolds numbers up to 160 x 10(exp 6), were in general somewhat higher than theoretical values. A maximum Reynolds number of transition of only 33 x 10(exp 6) was obtained. Contrary to theoretical and some other experimental investigations, the transition Reynolds number initially increased while the wall temperature ratio increased at relatively constant Mach number. Further increases in wall temperature ratio were accompanied by a decrease in transition Reynolds number. Increasing transition Reynolds number with increasing Mach number was also indicated at a relatively constant wall temperature ratio.

Merlet, Charles F.↗
NASA NTRS1995

Energy transfer in compressible turbulence

This letter investigates the compressible energy transfer process. We extend a methodology developed originally for incompressible turbulence and use databases from numerical simulations of a weak compressible turbulence based on Eddy-Damped-Quasi-Normal-Markovian (EDQNM) closure. In order to analyze the compressible mode directly, the well known Helmholtz decomposition is used. While the compressible component has very little influence on the solenoidal part, we found that almost all of the compressible turbulence energy is received from its solenoidal counterpart. We focus on the most fundamental building block of the energy transfer process, the triadic interactions. This analysis leads us to conclude that, at low turbulent Mach number, the compressible energy transfer process is dominated by a local radiative transfer (absorption) in both inertial and energy containing ranges.

Bataille, Francoise↗
NASA NTRS1996

Preliminary Comparison of Two-Way Satellite Time and Frequency Transfer and GPS Common-View Time Transfer During the INTELSAT Field Trial

For a decade and a half Global Positioning System (GPS) common-view time transfer has greatly served the needs of primary timing laboratories for regular intercomparisons of remote atomic clocks. However, GPS as a one-way technique has natural limits and may not meet all challenges of the comparison of the coming new generation of atomic clocks. Two-way satellite time and frequency transfer (TWSTFT) is a promising technique which may successfully complement GPS. For two years, regular TWSTFT's have been performed between eight laboratories situated in both Europe and North America, using INTELSAT satellites. This has enabled an extensive direct comparison to be made between these two high performance time transfer methods. The performance of the TWSTFT and GPS common view methods are compared over a number of time-transfer links. These links use a variety of time-transfer hardware and atomic clocks and have baselines of substantially different lengths. The relative merits of the two time-transfer systems are discussed.

Davis, John A.↗
NASA NTRS1991

Retrieval of total precipitable water over high latitude regions using radiometric measurements near 90 and 183 GHz

The functional dependence of brightness temperature, at frequencies near 183 and 90 GHz, on total precipitable water in the range less than 0.5 g/cm is examined by means of radiative transfer calculations of a number of observed temperature and relative humidity profiles over both ocean and land surfaces. A wide range of land and sea surface reflectivities is assumed in the calculations. Both sensitivity and precision of total precipitable water estimates are derived from these calculations. The estimated values in the Southern Hemisphere showed a gradual decrease toward the polar region, as expected. Radiative transfer calculations were made with a number of radiosonde data sets for a wide range of surface reflectivity and temperature. The calculated results are used to estimate total precipitable water from the airborne radiometric measurements over Alaska. Reflectivities at both 92 and 193 GHz are found to be mostly in the range of 0.2-0.4. High reflectivity is attributed to scattering by the snow particles on the ground.

Wang, J. R.↗
NASA NTRS1957

Measurements of Heat Transfer and Boundary-Layer Transition on an 8-Inch-Diameter Hemisphere-Cylinder in Free Flight for a Mach Number Range of 2.00 to 3.88

Measurements of aerodynamic heat transfer have been made along the hemisphere and cylinder of a hemisphere-cylinder rocket-propelled model in free flight up to a Mach number of 3.88. The test Reynolds number based on free-stream condition and diameter of model covered a range from 2.69 x l0(exp 6) to 11.70 x 10(exp 6). Laminar, transitional, and turbulent heat-transfer coefficients were obtained. The laminar data along the body agreed with laminar theory for blunt bodies whereas the turbulent data along the cylinder were consistently lower than that predicted by the turbulent theory for a flat plate. Measurements of heat transfer at the stagnation point were, in general, lower than the theory for stagnation-point heat transfer. When the Reynolds number to the junction of the hemisphere-cylinder was greater than 6 x l0(exp 6), the transitional Reynolds number varied from 0.8 x l0(exp 6) to 3.0 x 10(exp 6); however, than 6 x l(exp 6) when the Reynolds number to the junction was less, than the transitional Reynolds number varied from 7.0 x l0(exp 6) to 24.7 x 10(exp 6).

Garland, Benjamine J.↗
NASA NTRS1960

Heat Transfer to 36.75 and 45 degree Swept Blunt Leading Edges in Free Flight at Mach Numbers from 1.70 to 2.99 and From 2.50 to 4.05

A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, measured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of leading-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on the cylindrical portion of the leading edge.

ONeal, Robert L.↗
NASA NTRS1955

Analysis of turbulent heat transfer, mass transfer, and friction in smooth tubes at high Prandtl and Schmidt numbers

The expression for eddy diffusivity from a previous analysis was modified in order to account for the effect of kinematic viscosity on the turbulence in the region close to a wall. By using the modified expression, good agreement was obtained between predicted and experimental results for heat and mass transfer at Prandtl and Schmidt numbers between 0.5 and 3000. The effects of length-to-diameter ratio and of variable viscosity were also investigated for a wide range of Prandtl numbers.

Deissler, Robert G↗
NASA NTRS1982

Jet array impingement flow distributions and heat transfer characteristics: Effects of initial crossflow and nonuniform array geometry

Flow distributions and heat transfer characteristics for two-dimensional arrays of circular air jets impinging on a surface parallel to the jet orifice plate were determined. The configurations considered were intended to model those of interest in current and contemplated gas turbine airfoil midchord cooling applications. The geometry of the airfoil applications considered dictates that all of the jet flow, after impingement, exit in the chordwise (i.e., streamwise) direction toward the trailing edge. Experimental results for the effect of an initial crossflow on both flow distributions and heat transfer characteristics for a number of the prior uniform array geometries. The effects of nonuniform array geometries on flow distributions and heat transfer characteristics for noninitial crossflow configurations are discussed.

Florschuetz, L. W.↗
NASA NTRS1982

Full-coverage film cooling

Transpiration cooling is treated and then full coverage discrete hole injection for three injection orientations. Spacings with pitch to diameter ratios of 5 and 10 are discussed. The array is staggered, with the transverse pitch and the streamwise pitch the same. Results are presented in terms of the Stanton number using the heat transfer coefficient defined in terms of the difference between the wall temperature and the free stream temperature. Two values of Stanton number are provided for each situation: one with the injectant at wall temperature, and the other with the injectant at free stream temperature. These two values are equivalent to knowing the heat transfer coefficient and the adiabatic effectiveness. The heat transfer coefficient thus defined is used with the actual wall temperature to and the actual gas temperature to calculate the heat load. The principle of superposition thus invoked is valid exactly when the governing equations are linear.

Moffat, R. J.↗
NASA NTRS1960

Comparison of Turbulent Heat-Transfer Results for Uniform Wall Heat Flux and Uniform Wall Temperature

The purpose of this note is to examine in a more precise way how the Nusselt numbers for turbulent heat transfer in both the fully developed and thermal entrance regions of a circular tube are affected by two different wall boundary conditions. The comparisons are made for: (a) Uniform wall temperature (UWT); and (b) uniform wall heat flux (UHF). Several papers which have been concerned with the turbulent thermal entrance region problem are given. 1 Although these analyses have all utilized an eigenvalue formulation for the thermal entrance region there were differences in the choices of eddy diffusivity expressions, velocity distributions, and methods for carrying out the numerical solutions. These differences were also found in the fully developed analyses. Hence when making a comparison of the analytical results for uniform wall temperature and uniform wall heat flux, it was not known if differences in the Nusselt numbers could be wholly attributed to the difference in wall boundary conditions, since all the analytical results were not obtained in a consistent way. To have results which could be directly compared, computations were carried out for the uniform wall temperature case, using the same eddy diffusivity, velocity distribution, and digital computer program employed for uniform wall heat flux. In addition, the previous work was extended to a lower Reynolds number range so that comparisons could be made over a wide range of both Reynolds and Prandtl numbers.

Siegel, R.↗
NASA NTRS1985

Experimental aerodynamic heating to simulated space shuttle tiles in laminar and turbulent boundary layers with variable flow angles at a nominal Mach number of 7

The heat transfer to simulated shuttle thermal protection system tiles was investigated experimentally by using a highly instrumented metallic thin wall tile arranged with other metal tiles in a staggered tile array. Cold wall heating rate data for laminar and turbulent flow were obtained in the Langley 8 foot high Temperature Tunnel at a nominal Mach number of 7, a nominal total temperature of 3300R, a free stream unit Reynolds number from 3.4 x 10 sup 5 to 2.2 10 sup 6 per foot, and a free stream dynamic pressure from 2.1 to 9.0 psia. Experimental data are presented to illustrate the effects of flow angularity and gap width on both local peak heating and overall heating loads. For the conditions of the present study, the results show that localized and total heating are sensitive to changes in flow angle only for the test conditions of turbulent boundary layer flow with high kinetic energy and that a flow angle from 30 deg to 50 deg will minimize the local heating.

Avery, D. E.↗
NASA NTRS1958

Heat-Transfer and Pressure Measurements from a Flight Test of the Third 1/18-Scale Model of the Titan Intercontinental Ballistic Missile up to a Mach Number of 3.86 and Reynolds Number per Foot of 23.5 x 10(exp 6) and a Comparison with Heat Transfer

Heat-transfer and pressure measurements were obtained from a flight test of a 1/18-scale model of the Titan intercontinental ballistic missile up to a Mach number of 3.86 and Reynolds number per foot of 23.5 x 10(exp 6) and are compared with the data of two previously tested 1/18-scale models. Boundary-layer transition was observed on the nose of the model. Van Driest's theory predicted heat-transfer coefficients reasonably well for the fully laminar flow but predictions made by Van Driest's theory for turbulent flow were considerably higher than the measurements when the skin was being heated. Comparison with the flight test of two similar models shows fair repeatability of the measurements for fully laminar or turbulent flow.

Graham, John B., Jr.↗
NASA NTRS1974

Pressure and heat flux results from the space shuttle/external fuel tank interaction test at Mach numbers 16 and 19

Heat transfer rates and pressures were measured on a 0.0175-scale model of the space shuttle external tank (ET), model MCR0200. Tests were conducted with the ET model separately and while mated with a 0.0175-scale model of the orbiter, model 21-OT (Grumman). The tests were conducted in the AEDC-VKF Hypervelocity Wind Tunnel (F) at Mach numbers 16 and 19. The primary data consisted of the interaction heating rates experienced by the ET while mated with the orbiter in the flight configuration. Data were taken for a range of Reynolds numbers from 50,000 to 65,000 under laminar flow conditions.

Brewer, E. B.↗