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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 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

High Reynolds number test of a NACA 651-213, a equals 0.5 airfoil at transonic speeds

Wind-Tunnel tests were conducted in the Lockheed-Georgia Company's compressible flow facility to determine the transonic two-dimensional aerodynamic characteristics of a NACA 65 sub 1-213 a = 0.50 airfoil. The results are correlated with data obtained in the NASA-Langley 8-foot transonic pressure tunnel and the NAE high Reynolds number 15x60-inch two-dimensional test facility. The tests were conducted over a Mach number range from 0.60 to 0.80 and an angle of attack range from -1 deg to 8 deg. Reynolds numbers, based on the airfoil chord, were varied.

Burdges, K. P.↗
NASA NTRS1996

Reynolds-Averaged Navier-Stokes Studies of Low Reynolds Number Effects on the Losses in a Low Pressure Turbine

Experimental data from jet-engine tests have indicated that unsteady blade-row interaction effects can have a significant impact on the efficiency of low-pressure turbine stages. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Preliminary studies indicate that Reynolds number effects may contribute to the lower efficiencies at cruise conditions. In the current study, numerical experiments have been performed to quantify the Reynolds number dependence of unsteady wake/separation bubble interaction on the performance of a low-pressure turbine.

Dorney, Daniel J.↗
NASA NTRS1984

Comparative wind tunnel tests at high Reynolds numbers of NACA 64 621 airfoils with two aileron configurations

An experimental program to measure the aerodynamic characteristics of the NACA 64-621 airfoil when equipped with plain ailerons of 0.38 chord and 0.30 chord and with 0.38 chord balanced aileron has been conducted in a pressurized 6 x 12-inch High Reynolds Number Wind Tunnel. Surface pressures were measured and integrated to yield lift and pressure drag coefficients for angles of attack from -3 deg to +42 deg, and for selected aileron deflections from 0 to -90 deg at nominal Mach and Reynolds numbers of 0.25 and 5 x l0 exp 6, respectively. When resolved into thrust coefficient for wind turbine aerodynamic control applications, the data indicated the anticipated decrease in thrust coefficient with negative aileron deflection at low angles of attack; however, as angle of attack increased, thrust coefficients eventually became positive. All aileron configurations, even at -90 deg deflections, showed this trend. Hinge moments for each configuration complete the data set.

Gregorek, G. M.↗
NASA NTRS1995

Comparative wind tunnel test at high Reynolds numbers of NACA 64 621 airfoils with two aileron configurations

An experimental program to measure the aerodynamic characteristics of the NACA 64-621 airfoil when equipped with plain ailerons of 0.38 chord and 0.30 chord and with 0.38 chord balanced aileron has been conducted in the pressurized O.S.U. 6 x 12 ft High Reynolds Number Wind Tunnel. Surface pressures were measured and integrated to yield lift and pressure drag coefficients for angles of attack from -3 to +42 deg and for selected aileron deflections from 0 to -90 deg at nominal Mach and Reynolds numbers of 0.25 and 5 x 10(exp 6). When resolved into thrust coefficient for wind turbine aerodynamic control applications, the data indicated the anticipated decrease in thrust coefficient with negative aileron deflection at low angles of attack; however, as angle of attack increased, thrust coefficients eventually became positive. All aileron configurations, even at -90 deg deflections showed this trend. Hinge moments for each configuration complete the data set.

Gregorek, G. M.↗
NASA NTRS1974

Flight Reynolds number effects on a fighter-type, circular-arc-19 deg conic boattail nozzle at subsonic speeds

A circular-arc - conic boattail nozzle, typical of those used on a twin engine fighter, was tested on an underwing nacelle mounted on an F-106B aircraft. The boattail had a radius ratio r/r sub c of 0.41 and a terminal boattail angle of approximately 19 deg. The gas generator was a J85-GE-13 turbojet engine. The effects of Reynolds number and angle of attack on boattail pressure drag and boattail pressure profiles were investigated. Increasing Reynolds number resulted in reduced boattail drag at both Mach numbers of 0.6 and 0.9.

Chamberlin, R.↗
NASA NTRS1979

On the laminar separation, transition, and turbulent reattachment of low Reynolds number flows near the leading edge of airfoils

The laminar separation, transition, and turbulent reattachment near the leading edge of a cylindrical nose-constant thickness airfoil model were investigated using a low turbulence, low speed smoke wind tunnel. The locations of separation, transition, and reattachment were obtained from smoke flow photographs and surface oil flow techniques for chord Reynolds numbers from about 150,000 to 470,000. These visual data combined with static pressure distributions and lift and drag measurements delineate the effects of angle of attack, flap deflection angle, and chord Reynolds number on the separation bubble characteristics. The data concerning the length of the laminar and turbulent portions of the bubble agree with the empirical prediction methods for short bubbles.

Arena, A. V.↗
NASA NTRS1979

Trailing-edge flows at high Reynolds number

An investigation of trailing-edge flows at high Reynolds number and subsonic Mach numbers is presented. Symmetric and asymmetric trailing-edge flows are studied, each flow having pressure gradient regions upstream of the trailing edge similar to an airfoil. Measurements include model surface pressures, mean velocity, turbulent shear stress, and turbulent kinetic energy profiles in the trailing-edge and near-wake regions. Comparisons of the symmetric data with numerical solutions of boundary layer as well as Navier-Stokes equations employing two different turbulence models show increasing effects on viscous interactions as the Mach number increases. Both turbulence models yielded solutions of the mean flow of comparable quality. The experimental results of the asymmetric case are discussed.

Viswanath, P. R.↗
NASA NTRS1981

High Reynolds number research - 1980

The fundamental aerodynamic questions for which high Reynolds number experimental capability is required were examined. Potential experiments which maximize the research returns from the use of the National Transonic Facility (NTF) were outlined. Calibration plans were reviewed and the following topics were discussed: fluid dynamics; high lit; configuration aerodynamics; aeroelasticity and unsteady aerodynamics; wind tunnel/flight correlation; space vehicles; and theoretical aerodynamics

Mckinney, L. W.↗
NASA NTRS2006

Low Reynolds Number Aerodynamic Characteristics of Several Airplane Configurations Designed to Fly in the Mars Atmosphere at Subsonic Speeds

A 1/4-scale wind tunnel model of an airplane configuration developed for short duration flight at subsonic speeds in the Martian atmosphere has been tested in the Langley Research Center Transonic Dynamics Tunnel. The tunnel was pumped down to extremely low pressures to represent Martian Mach/Reynolds number conditions. Aerodynamic data were obtained and upper and lower surface wind pressures were measured at one spanwise station on some configurations. Three unswept wings of the same planform but different airfoil sections were tested. Horizontal tail incidence was varied as was the deflection of plain and split trailing-edge flaps. One unswept wing configuration was tested with the lower part of the fuselage removed and the vertical/horizontal tail assembly inverted and mounted from beneath the fuselage. A sweptback wing was also tested. Tests were conducted at Mach numbers from 0.50 to 0.90. Wing chord Reynolds number was varied from 40,000 to 100,000 and angles of attack and sideslip were varied from -10deg to 20deg and -10deg to 10deg, respectively.

Re, Richard J.↗
NASA NTRS2008

Active Flow Control at Low Reynolds Numbers on a NACA 0015 Airfoil

Results from a low Reynolds number wind tunnel experiment on a NACA 0015 airfoil with a 30% chord trailing edge flap tested at deflection angles of 0, 20, and 40 are presented and discussed. Zero net mass flux periodic excitation was applied at the ap shoulder to control flow separation for flap deflections larger than 0. The primary objective of the experiment was to compare force and moment data obtained from integrating surface pressures to data obtained from a 5-component strain-gage balance in preparation for additional three-dimensional testing of the model. To achieve this objective, active flow control is applied at an angle of attack of 6 where published results indicate that oscillatory momentum coefficients exceeding 1% are required to delay separation. Periodic excitation with an oscillatory momentum coefficient of 1.5% and a reduced frequency of 0.71 caused a significant delay of separation on the airfoil with a flap deflection of 20. Higher momentum coefficients at the same reduced frequency were required to achieve a similar level of flow attachment on the airfoil with a flap deflection of 40. There was a favorable comparison between the balance and integrated pressure force and moment results.

Melton, LaTunia Pack↗
NASA NTRS2018

Direct Numerical Simulation of Receptivity to Roughness in a Swept-Wing Boundary Layer at High Reynolds Numbers

Direct numerical simulations (DNS) are performed to examine the receptivity to roughness in a spatially developing three-dimensional boundary layer over an in finite-swept natural-laminar-flow wing at a free stream Mach number of 0:75 and a chord Reynolds number of approximately 25 million based on the long, swept chord. Stationary cross ow disturbances are excited by applying either critically spaced discrete cylinders of micron size or naturally occurring distributed roughness in the leading-edge region. The DNS data show that the spanwise spectral content of the excited cross ow disturbances is highly dependent upon the shape of roughness elements, and the initial growth of the cross ow structures is a nonlinear function of the element height. The linear growth rate of the excited cross ow disturbances predicted by DNS shows good agreement with linear parabolized stability equations. The receptivity study lays the foundation for investigating the stabilization of the naturally most unstable steady cross ow mode by using spanwise periodic DREs.

Nicholson, Gary↗
NASA NTRS2019

Stabilization of a Swept-Wing Boundary Layer by Discrete Roughness Elements at High Reynolds Numbers

Direct numerical simulations (DNS) are performed to study potential stabilizing ef- fect of spanwise periodic discrete roughness elements (DREs) on cross ow instabilities in a spatially developing three-dimensional boundary layer over an in nite-swept natural- laminar- ow wing at a freestream Mach number of 0:75 and a chord Reynolds number of approximately 25 million. In the DNS, both the spanwise periodic DREs and distributed roughness in the leading-edge region are implemented to simulate a typical experimen- tal scenario in which multiple steady cross ow modes including the most unstable mode (i.e., the \target" mode) emerge because of the presence of naturally distributed surface roughness in the leading edge region and spanwise periodic control cylinders of subcritical wavelength are used to force small-wavelength disturbances (i.e., the control mode) for damping the target mode. The DNS results show that the e ectiveness of DRE control is sensitive to roughness diameter, height, and chordwise placement. For the DRE parame- ters considered in this study, the stabilizing e ect on the target mode is small within the computational domain that ended at about 35% of the chord.

Nicholson, Gary↗
NASA NTRS1979

Reynolds number, scale and frequency content effects on F-15 inlet instantaneous distortion

An inlet instantaneous distortion study program sponsored by NASA was recently completed using an F-15 fighter aircraft. Peak distortion data from subscale inlet model wind tunnel tests are shown to be representative of full-scale flight test peak distortion. The effects on peak distortion are investigated for engine presence, Reynolds number, scale and frequency content. Data are presented which show that: (1) the effect of engine presence on total pressure recovery, peak distortion, and turbulence is small but favorable, (2) increasing the Reynolds number increases total pressure recovery, decreases peak distortion, and decreases turbulence, and (3) increasing the filter cutoff frequency increases the calculated values of both peak distortion and turbulence.

Stevens, C. H.↗
NASA NTRS1984

The application of cryogenics to high Reynolds number testing in wind tunnels. I - Evolution, theory, and advantages

During the time which has passed since the construction of the first wind tunnel in 1870, wind tunnels have been developed to a high degree of sophistication. However, their development has consistently failed to keep pace with the demands placed on them. One of the more serious problems to be found with existing transonic wind tunnels is their inability to test subscale aircraft models at Reynolds numbers sufficiently near full-scale values to ensure the validity of using the wind tunnel data to predict flight characteristics. The Reynolds number capability of a wind tunnel may be increased by a number of different approaches. However, the best solution in terms of model, balance, and model support loads, as well as in terms of capital and operating cost appears to be related to the reduction of the temperature of the test gas to cryogenic temperatures. The present paper has the objective to review the evolution of the cryogenic wind tunnel concept and to describe its more important advantages.

Kilgore, R. A.↗
NASA NTRS2000

Review of Skin Friction Measurements Including Recent High-Reynolds Number Results from NASA Langley NTF

This paper reviews flat plate skin friction data from early correlations of drag on plates in water to measurements in the cryogenic environment of The NASA Langley National Transonic Facility (NTF) in late 1996. The flat plate (zero pressure gradient with negligible surface curvature) incompressible skin friction at high Reynolds numbers is emphasized in this paper, due to its importance in assessing the accuracy of measurements, and as being important to the aerodynamics of large scale vehicles. A correlation of zero pressure gradient skin friction data minimizing extraneous effects between tests is often used as the first step in the calculation of skin friction in complex flows. Early data compiled by Schoenherr for a range of momentum thickness Reynolds numbers, R(sub Theta) from 860 to 370,000 contained large scatter, but has proved surprisingly accurate in its correlated form. Subsequent measurements in wind tunnels under more carefully controlled conditions have provided inputs to this database, usually to a maximum R(sub Theta) of about 40,000. Data on a large axisymmetric model in the NASA Langley National Transonic Facility extends the upper limit in incompressible R(sub Theta) to 619,800 using the van Driest transformation. Previous data, test techniques, and error sources ar discussed, and the NTF data will be discussed in detail. The NTF Preston tube and Clauser inferred data accuracy is estimated to be within -2 percent of a power-law curve fit, and falls above the Spalding theory by 1 percent at R(sub Theta) of about 600,000.

Watson, Ralph D.↗