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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 NTRS1960

The Influence of Low Wall Temperature on Boundary-Layer Transition and Local Heat Transfer on 2-Inch-Diameter Hemispheres at a Mach Number of 4.95 and a Reynolds Number per Foot of 73.2 x 10(exp 6)

Measurements of the location of boundary-layer transition and the local heat transfer have been made on 2-inch-diameter hemispheres in the Langley gas dynamics laboratory at a Mach number of 4.95, a Reynolds number per foot of 73.2 x 10(exp 6), and a stagnation temperature of approximately 400 F. The transient-heating thin-skin calorimeter technique was used, and the initial values of the wall-to-stream stagnation- temperature ratios were 0.16 (cold-model tests) and 0.65 (hot-model test). During two of the four cold tests, the boundary-layer flow changed from turbulent to laminar over large regions of the hemisphere as the model heated. On the basis of a detailed consideration of the magnitude of roughness possibly present during these two cold tests, it appears that this destabilizing effect of low wall temperatures (cooling) was not caused by roughness as a dominant influence. This idea of a decrease in boundary-layer stability with cooling has been previously suggested. (See, for example, NASA Memorandum 10-8-58E.) For the laminar data obtained during the early part of the hot test, the correlation of the local-heating data with laminar theory was excellent.

Cooper, Morton↗
NASA NTRS1991

Analysis of Fluctuating Static Pressure Measurements in a Large High Reynolds Number Transonic Cryogenic Wind Tunnel

Dynamic measurements of fluctuating static pressure levels were made using flush mounted high frequency response pressure transducers at eleven locations in the circuit of the National Transonic Facility (NTF) over the complete operating range of this wind tunnel. Measurements were made at test section Mach numbers from 0.2 to 1.2, at pressure from 1 to 8.6 atmospheres and at temperatures from ambient to -250 F, resulting in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made independently at variable Mach number, variable Reynolds number, and variable drivepower, each time keeping the other two variables constant thus allowing for the first time, a distinct separation of these three important variables. A description of the NTF emphasizing its flow quality features, details on the calibration of the instrumentation, results of measurements with the test section slots covered, downstream choke, effects of liquid nitrogen injection and gaseous nitrogen venting, comparisons between air and nitrogen, isolation of the effects of Mach number, Reynolds number, and fan drive power, and identification of the sources of significant flow disturbances is included. The results indicate that primary sources of flow disturbance in the NTF may be edge-tones generated by test section sidewall re-entry flaps and the venting of nitrogen gas from the return leg of the tunnel circuit between turns 3 and 4 in the cryogenic mode of operation. The tests to isolate the effects of Mach number, Reynolds number, and drive power indicate that Mach number effects predominate. A comparison with other transonic wind tunnels shows that the NTF has low levels of test section fluctuating static pressure especially in the high subsonic Mach number range from 0.7 to 0.9.

Igoe, William B.↗
NASA NTRS1979

Flowfield and acoustic properties of a Mach number 0.9 jet at a low Reynolds number

An anechoic vacuum chamber test facility is used to assess the flow field and acoustic properties of a jet with a low Reynolds number of about 3600 and a Mach number of 0.9. A conventional Pitot pressure probe is used to monitor the mean flow field, a hot-wire probe is employed to detect flow fluctuations, and the acoustic field is probed by condenser microphones. The jet is operated at the low Reynolds number in order to examine the behavior of the large-scale 'orderly' structure with particular emphasis on its noise-generating capabilities. It is shown that a nonlinear mechanism involving the large-scale flow instability is responsible for a dominant portion of the noise generated from the jet.

Stromberg, J. L.↗
NASA NTRS2005

Off-Design Reynolds Number Effects for a Supersonic Transport

A high Reynolds number wind tunnel test was conducted to assess Reynolds number effects on the aerodynamic performance characteristics of a realistic, second-generation supersonic transport concept. The tests included longitudinal studies at transonic and low-speed, high-lift conditions across a range of chord Reynolds numbers (8 million to 120 million). Results presented focus on Reynolds number and static aeroelastic sensitivities at Mach 0.30 and 0.90 for a configuration without a tail. Static aeroelastic effects, which mask Reynolds number effects, were observed. Reynolds number effects were generally small and the drag data followed established trends of skin friction as a function of Reynolds number. A more nose-down pitching moment was produced as Reynolds number increased because of an outward movement of the inboard leading-edge separation at constant angles of attack. This study extends the existing Reynolds number database for supersonic transports operating at off-design conditions.

Owens, Lewis R.↗
NASA NTRS1954

Performance and boundary-layer data from 12 degree and 23 degree conical diffusers of area ratio 2.0 at Mach numbers up to choking and Reynolds numbers up to 7.5 x 10(6)

For each of two inlet-boundary-layer thicknesses, performance and boundary-layer characteristics have been determined for a 12 degree, 10-inch-inlet-diameter diffuser, a 12 degree, 21-inch-inlet-diameter diffuser, and a 23 degree, 21-inch-inlet-diameter diffuser. The investigation covered an inlet Mach number range from about 0.10 to coking. The corresponding inlet Reynolds number, based on inlet diameter, varied from about 0.5 x 10(6) to 7.5 x 10(6).

Little, B H , Jr↗
NASA NTRS1989

Investigation of CAST-10-2/DOA 2 airfoil in NAE high Reynolds number two-dimensional test facility

A common airfoil model with the CAST 10-2/DOA-2 profile and 228 mm (9 inches) chord length was tested. The tests performed in NAE covered the Mach numbers from 0.3 to 0.8 and Reynolds numbers from 10 to 30 million. The model was tested with transition free and with transition fixed at 5 percent chord for both the upper and the lower surfaces. The data obtained were analyzed for the effects of Reynolds number, transition fixing and Mach number. The role of the boundary layer on the displacement effect, the interaction with the shock wave and the trailing edge separation are examined. The results are summarized as follows: (1) the airfoil performance depends strongly on Reynolds number and transition fixing; (2) with transition fixed, the aerodynamic quantities such as lift, pitching moment and drag show a monotonic variation with Reynolds number; (3) with transition free, the aerodynamic quantities vary less regularly with Reynolds number and a slight parametric dependency is shown. The weak dependency is due to the compensatory effect of the forward shift of the transition position and the thinning of the turbulent boundary layer as Reynolds number increases; (4) the shock Mach number and the shock position are weakly dependent on Reynolds number; and (5) the long extent of the laminar boundary layer at transonic speeds reduces the drag appreciably at low Reynolds numbers. The drag bucket around the design Mach number can be observed below Reynolds number 15 million.

Chan, Y. Y.↗
NASA NTRS1986

High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

Johnson, Charles B.↗
NASA NTRS1956

Comparison of experimental and theoretical normal-force distributions (including Reynolds number effects) on an ogive-cylinder body at Mach number 1.98

Effects of Reynolds number and angle of attack on the pressure distribution and normal-force characteristics of a body of revolution consisting of a fineness ratio 3 ogival nose tangent to a cylindrical afterbody 7 diameters long have been determined. The test Mach number was 1.98 and the angle-of-attack range from 0 degree to 20 degrees. The Reynolds numbers, based on body diameter, were 0.15 x 10(6) and 0.45 x 10(6). The experimental results are compared with theory.

Perkins, Edward W↗
NASA NTRS2007

The Role of Separation Bubbles on the Aerodynamic Characteristics of Airfoils, Including Stall and Post-Stall, at Low Reynolds Numbers

Airfoils at high Reynolds numbers, in general, have small separation bubbles that are usually confined to the leading edge. Since the Reynolds number is large, the turbulence model for the transition region between the laminar and turbulent flow is not important. Furthermore, the onset of transition occurs either at separation or prior to separation and can be predicted satisfactorily by empirical correlations when the incident angle is small and can be assumed to correspond to laminar separation when the correlations do not apply, i.e., at high incidence angles.

Chen, Hsun H.↗
NASA NTRS1976

Effect of Reynolds number on the aerodynamic stability and control characteristics of a 55 deg clipped-delta-wing orbiter configuration at supersonic Mach numbers

A wind-tunnel investigation has been conducted at Mach numbers from 1.60 to 4.63 for Reynolds numbers varying from approximately 0.8 million to 10.0 million (based on body length) to determine the Reynolds number effects on the static longitudinal and lateral aerodynamic stability and control characteristics of two scale models of the Grumman H-33 space shuttle orbiter. The results indicate that for the Reynolds number range investigated, there are no Reynolds number effects on the longitudinal or lateral aerodynamic characteristics of either model with zero control deflections and only negligible effects with control deflections. The roll control is constant for both models with the possible exception of a slight increase with increases in Reynolds number up to 2.0 million at the lower Mach numbers. This very small Reynolds number effect is much more apparent for the smaller model and disappears with increasing Mach and Reynolds numbers.

Blair, A. B., Jr.↗
NASA NTRS1966

Grid turbulence at large Reynolds numbers.

Grid turbulence for large Reynolds numbers, noting consistency of power spectra values with Kolmogoroff scaling, decay law and Reynolds number effect on turbulence

GRID↗
NASA NTRS1960

Effect of Reynolds Number on the Force and Pressure Distribution Characteristics of a Two-Dimensional Lifting Circular Cylinder

A two-dimensional lifting circular cylinder has been tested over a Mach number range from 0.011 to 0.32 and a Reynolds number range from 135,000 to 1,580,000 to determine the force and pressure distribution characteristics. Two flaps having chords of 0.37 and 6 percent of the cylinder diameter, respectively, and attached normal to the surface were used to generate lift. A third configuration which had 6-percent flaps 1800 apart was also investigated. All flaps were tested through a range of angular positions. The investigation also included tests of a plain cylinder without flaps. The lift coefficient showed a wide variation with Reynolds number for the 6-percent flap mounted on the bottom surface at the 50-percent-diameter station, varying from a low of about 0.2 at a Reynolds number of 165,000 to a high of 1.54 at a Reynolds number of 350,000 and then decreasing almost linearly to a value of 1.0 at a Reynolds number of 1,580,000. The pressure distribution showed that the loss of lift with Reynolds number above the critical was the result of the separation point moving forward on the upper surface. Pressure distributions on a plain cylinder also showed similar trends with respect to the separation point. The variation of drag coefficient with Reynolds number was in direct contrast to the lift coefficient with the minimum drag coefficient of 0.6 occurring at a Reynolds number of 360,000. At this point the lift-drag ratios were a maximum at a value of 2.54. Tests of a flap with a chord of 0.0037 diameter gave a lift coefficient of 0.85 at a Reynolds number of 520,000 with the same lift-drag ratio as the larger flap but the position of the flap for maximum lift was considerably farther forward than on the larger flap. Tests of two 6-percent flaps spaced 180 deg apart showed a change in the sign of the lift developed for angular positions of the flap greater than 132 deg at subcriti- cal Reynolds numbers. These results may find use in application to air- craft using forebody strakes. The drag coefficient developed by the flaps when normal to the relative airstream was approximately equal to that developed by a flat plate in a similar attitude.

Lockwood, Vernard E.↗
NASA NTRS1939

Preliminary Investigation of Certain Laminar-Flow Airfoils for Application at High Speeds and Reynolds Numbers

In order to extend the useful range of Reynolds numbers of airfoils designed to take advantage of the extensive laminar boundary layers possible in an air stream of low turbulence, tests were made of the NACA 2412-34 and 1412-34 sections in the NACA low-turbulence tunnel. Although the possible extent of the laminar boundary layer on these airfoils is not so great as for specially designed laminar-flow airfoils, it is greater than that for conventional airfoils, and is sufficiently extensive so that at Reynolds numbers above 11,000,000 the laminar region is expected to be limited by the permissible 'Reynolds number run' and not by laminar separation as is the case with conventional airfoils. Drag measurements by the wake-survey method and pressure-distribution measurements were made at several lift coefficients through a range of Reynolds numbers up to 11,400,000. The drag scale-effect curve for the NACA 1412-34 is extrapolated to a Reynolds number of 30,000,000 on the basis of theoretical calculations of the skin friction. Comparable skin-friction calculations were made for the NACA 23012. The results indicate that, for certain applications at moderate values of the Reynolds number, the NACA 1412-34 and 2412-34 airfoils offer some advantages over such conventional airfoils as the NACA 23012. The possibility of maintaining a more extensive laminar boundary layer on these airfoils should result in a small drag reduction, and the absence of pressure peaks allows higher speeds to be reached before the compressibility burble is encountered. At lower Reynold numbers, below about 10,000,000, these airfoils have higher drags than airfoils designed to operate with very extensive laminar boundary layers.

Jacobs, E.N.↗
NASA NTRS1995

Reynold-Number Effects on Near-Wall Turbulence

The Reynolds stress budget in a full developed turbulent channel flow for three Reynolds numbers (Re = 180,395,590) are used to investigate the near wall scaling of various turbulence quantities. We find that as the Reynolds number increases, the extent of the region where the production of the kinetic energy is equal to the dissipation increases. At the highest Reynolds number the region of equilibrium extends from y+ - 120 to y+ = 240. As the Reynolds number increases, we find that wall scaling collapses the budgets for the streamwise fluctuating component, but the budgets for the other two components show Reynolds number dependency.

Mansour, N. N.↗
NASA NTRS1980

Flow field and acoustic properties of a Mach number 0.9 jet at a low Reynolds number

An experimental study of the flow field and acoustic properties of a low Reynolds number M = 0.9 jet has been performed in a low pressure anechoic test chamber. The mean flow field was surveyed with a conventional pitot pressure probe and flow fluctuations were detected with a normal hot wire probe. Also, condenser microphone measurements were made in the acoustic field. The flow fluctuations within the jet were found to be dominated initially by a relatively discrete, large-scale, wave-like instability centered around a Strouhal number of 0.44. The axial wavelength of this instability was determined to be 1.45 jet diameters and its azimuthal character includes the n = 0 and n = plus or minus 1 modes. The growth of this instability coupled with its nonlinear breakdown are major contributors to the termination of the potential core region of the jet. The acoustic field of the jet, in contrast to the flow field, has a broad frequency spectrum with a peak amplitude near a Strouhal number of St = 0.2. The results indicate that a nonlinear mechanism involving the large-scale flow instability is responsible for a dominant portion of the noise generated from this jet.

Stromberg, J. L.↗