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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 NTRS1986

Effect of Reynolds Number and Mach Number on flow angularity probe sensitivity

Preliminary calibrations were performed on nine flow angularity probes in the Langley 7- by 10-Foot High-Speed Tunnel (7 x 10 HST) and the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). These probes will be used in surveying the test section flows of the National Transonic Facility (NTF). The probes used in this study have a pyramid head with five pressure orifices. The calibrations consisted of both isolated probe measurements and rake-mounted multiprobe measurements that covered a range of subsonic Mach numbers up to 0.90 and Reynolds numbers per foot up to 40 X 10 to the 6th power. The preliminary calibration in the 7 x 10 HST included testing the probes both individually and in a rake. The 0.3-m TCT calibration tested two probes singly at varying Reynolds numbers. The results from these tests include Mach number, Reynolds number, and rake-mounting effects. The results of these tests showed probe sensitivity to be slightly affected by Mach number. At Reynolds numbers per foot above 10 x 10 to the 6th power, the probe did not exhibit a Reynolds number sensitivity.

Smith, L. A.↗
NASA NTRS1961

Experimental Investigation of a Hypersonic Glider Configuration at a Mach Number of 6 and at Full-Scale Reynolds Numbers

The aerodynamic characteristics of a hypersonic glider configuration, consisting of a slender ogive cylinder with three highly swept wings, spaced 120 apart, with the wing chord equal to the body length, were investigated experimentally at a Mach number of 6 and at Reynolds numbers from 6 to 16 million. The objectives were to evaluate the theoretical procedures which had been used to estimate the performance of the glider, and also to evaluate the characteristics of the glider itself. A principal question concerned the viscous drag at full-scale Reynolds number, there being a large difference between the total drags for laminar and turbulent boundary layers. It was found that the procedures which had been applied for estimating minimum drag, drag due to lift, lift curve slope, and center of pressure were generally accurate within 10 percent. An important exception was the non-linear contribution to the lift coefficient which had been represented by a Newtonian term. Experimentally, the lift curve was nearly linear within the angle-of-attack range up to 10 deg. This error affected the estimated lift-drag ratio. The minimum drag measurements indicated that substantial amounts of turbulent boundary layer were present on all models tested, over a range of surface roughness from 5 microinches maximum to 200 microinches maximum. In fact, the minimum drag coefficients were nearly independent of the surface smoothness and fell between the estimated values for turbulent and laminar boundary layers, but closer to the turbulent value. At the highest test Reynolds numbers and at large angles of attack, there was some indication that the skin friction of the rough models was being increased by the surface roughness. At full-scale Reynolds number, the maximum lift-drag ratio with a leading edge of practical diameter (from the standpoint of leading-edge heating) was 4.0. The configuration was statically and dynamically stable in pitch and yaw, and the center of pressure was less than 2-percent length ahead of the centroid of plan-form area.

Seiff, Alvin↗
NASA NTRS1959

Lift, Drag, and Pitching Moments of an Arrow Wing Having 80 Degree of Sweepback at Mach Numbers from 2.48 to 3.51 and Reynolds Numbers up to 11.0 Million

Measurements were made of the lift, drag, and pitching moments on an arrow wing (taper ratio of zero) having an aspect ratio of 1.4 and a leading-edge sweepback of 80 (degrees). The wing was designed to have a subsonic leading-edge and a Clark-Y airfoil with a thickness ratio of 12 percent of the chord perpendicular to the wing leading edge. The wing was tested both with and without the wing tips bent upward in an attempt to alleviate possible flow separation in the vicinity of the wing tips. Small jets of air were used to fix transition near the wing leading edge. Force results are presented for Mach numbers of 2.48, 2.75, 3.04, 3.28, and 3.51 at Reynolds numbers of 3.5 and 9.0 million and for a Mach number of 3.04 at a Reynolds number of 11.0 million. The measured aerodynamic characteristics are compared with those estimated by linear theory. The maximum lift-drag ratio measured was much less than that predicted. This difference is attributed to lack of full leading-edge thrust and to the experimental lift-curve slope being about 20 percent below the theoretical value.

Hopkins, Edward J.↗
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↗