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At least 37 records · Page 2

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NASA NTRS1972

Results of a study of Mach number and Reynolds number effects on the lee side vortex flow field characteristics of an ogive-cylinder-frustum-cylinder at angles of attack to 25 degrees, volume 1

An experimental program was conducted to survey the lee side vortex flow field about an ogive-cylinder-frustum-cylinder at angles of attack to 25 degrees for two Reynolds numbers at Mach number 0.8, and one Reynolds number at Mach number 1.96. The data were obtained using miniature 5-port conical pressure probes calibrated for angle of attack and roll angle over a Mach number range of 0.6 to 3.0. The results are presented here as local flow field properties and circulation strengths for various body stations.

Foley, J. E.↗

NASA NTRS1978

Flight-measured pressure characteristics of aft-facing steps in high Reynolds number flow at Mach numbers of 2.20, 2.50, and 2.80 and comparison with other data

The YF-12 airplane was studied to determine the pressure characteristics associated with an aft-facing step in high Reynolds number flow for nominal Mach numbers of 2.20, 2.50, and 2.80. Base pressure coefficients were obtained for three step heights. The surface static pressures ahead of and behind the step were measured for the no-step condition and for each of the step heights. A boundary layer rake was used to determine the local boundary layer conditions. The Reynolds number based on the length of flow ahead of the step was approximately 10 to the 8th power and the ratios of momentum thickness to step height ranged from 0.2 to 1.0. Base pressure coefficients were compared with other available data at similar Mach numbers and at ratios of momentum thickness to step height near 1.0. In addition, the data were compared with base pressure coefficients calculated by a semiempirical prediction method. The base pressure ratios are shown to be a function of Reynolds number based on momentum thickness. Profiles of the surface pressures ahead of and behind the step and the local boundary layer conditions are also presented.

Powers, S. G.↗

NASA NTRS1980

A fan pressure ratio correlation in terms of Mach number and Reynolds number for the Langley 0.3 meter transonic cryogenic tunnel

Calibration data for the two dimensional test section of the Langley 0.3 Meter Transonic Cryogenic Tunnel were used to develop a Mach number-Reynolds number correlation for the fan pressure ratio in terms of test section conditions. Well established engineering relationships combined to form an equation which is functionally analogous to the correlation. A geometric loss coefficient which is independent of Reynolds number or Mach number was determined. Present and anticipated uses of this concept include improvement of tunnel control schemes, comparison of efficiencies for operationally similar wind tunnels, prediction of tunnel test conditions and associated energy usage, and determination of Reynolds number scaling laws for similar fluid flow systems.

Lawing, P. L.↗

NASA NTRS1989

The effects of Reynolds number and Richardson number on the structure of a vertical co-flowing buoyant jet

The behavior of a vertical jet of helium issuing into a coflow of air at a fixed exit velocity ratio of 2.0 has been studied for various Reynolds numbers and Richardson numbers. It is found that the transition to turbulence is very sudden and that the point of transition moves closer to the jet exit as either the Reynolds number or the Richardson number increases. Under most of the conditions considered, the flow exhibits a strong periodic longitudinal instability whose wavelength increases with Richardson number.

Subbarao, E. R.↗

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 NTRS1961

Heat Transfer and Boundary-Layer Transition on a Highly Polished Hemisphere-Cone in Free Flight at Mach Numbers Up to 3.14 and Reynolds Numbers Up to 24 x 10(exp 6)

A highly polished hemisphere-cone having a ratio of nose radius to base radius of 0.74 and a half-angle of 14.5 was flight tested at Mach numbers up to 4.70. Temperature and pressure data were obtained at Mach numbers up to 3.14 and a free-stream Reynolds number of 24 x 10(exp 6) based on body diameter. The nose of the model had a surface roughness of 2 to 5 microinches as measured with an interferometer. The measured Stanton numbers were in good agreement with theory. Transition Reynolds numbers based on the laminar boundary-layer momentum thickness at transition ranged from 2,190 to 794. Comparison with results from previous tests of blunt shapes having a surface roughness of 20 to 40 microinches showed that the high degree of polish was instrumental in delaying the transition from laminar to turbulent flow.

Buglia, James J.↗

NASA NTRS1959

Measurements of Local Heat Transfer and Pressure on Six 2-Inch-Diameter Blunt Bodies at a Mach Number of 4.95 and at Reynolds Numbers Per Foot up to 81 x 10(exp 6)

Measurements of the local heat transfer and pressure distribution have been made on six 2-inch-diameter, blunt, axially symmetric bodies in the Langley gas dynamics laboratory at a Mach number of 4.95 and at Reynolds numbers per foot up to 81 x 10(exp 6). During the investigation laminar flow was observed over a hemispherical-nosed body having a surface finish from 10 to 20 microinches at the highest test Reynolds number per foot (for this configuration) of 77.4 x 10(exp 6). Though it was repeatedly possible to measure completely laminar flow at this Reynolds number for the hemisphere, it was not possible to observe completely laminar flow on the flat-nosed body for similar conditions. The significance of this phenomenon is obscured by the observation that the effects of particle impacts on the surface in causing roughness were more pronounced on the flat-nosed body. For engineering purposes, a method developed by M. Richard Dennison while employed by Lockheed Aircraft Corporation appears to be a reasonable procedure for estimating turbulent heat transfer provided transition occurs at a forward location on the body. For rearward-transition locations, the method is much poorer for the hemispherical nose than for the flat nose. The pressures measured on the hemisphere agreed very well with those of the modified Newtonian theory, whereas the pressures on all other bodies, except on the flat-nosed body, were bracketed by modified Newtonian theory both with and without centrifugal forces. For the hemisphere, the stagnation-point velocity gradient agreed very well with Newtonian theory. The stagnation-point velocity gradient for the flat- nosed model was 0.31 of the value for the hemispherical-nosed model. If a Newtonian type of flow is assumed, the ratio 0.31 will be independent of Much number and real-gas effects.

Cooper, Morton↗

NASA NTRS1959

Investigation at Mach Numbers of 0.20 to 3.50 of a Blended Diamond Wing and Body Combination of Sonic Design but with Low Wave-Drag Increase with Increasing Mach Number

A diamond wing and body combination was designed to have an area distribution which would result in near optimum zero-lift wave-drag coefficients at a Mach number of 1.00, and decreasing wave-drag coefficient with increasing Mach number up to near sonic leading-edge conditions for the wing. The airfoil section were computed by varying their shape along with the body radii (blending process) to match the selected area distribution and the given plan form. The exposed wing section had an average maximum thickness of about 3 percent of the local chords, and the maximum thickness of the center-line chord was 5.49 percent. The wing had an aspect ratio of 2 and a leading-edge sweep of 45 deg. Test data were obtained throughout the Mach number range from 0.20 to 3.50 at Reynolds numbers based on the mean aerodynamic chord of roughly 6,000,000 to 9,000,000. The zero-lift wave-drag coefficients of the diamond model satisfied the design objectives and were equal to the low values for the Mach number 1.00 equivalent body up to the limit of the transonic tests. From the peak drag coefficient near M = 1.00 there was a gradual decrease in wave-drag coefficient up to M = 1.20. Above sonic leading-edge conditions of the wing there was a rise in the wave-drag coefficient which was attributed in part to the body contouring as well as to the wing geometry. The diamond model had good lift characteristics, in spite of the prediction from low-aspect-ratio theory that the rear half of the diamond wing would carry little lift. The experimental lift-curve slope obtained at supersonic speeds were equal to or greater than the values predicted by linear theory. Similarly the other basic aerodynamic parameters, aerodynamic center position, and maximum lift-drag ratios were satisfactorily predicted at supersonic speeds.

Holdaway, George H.↗

NASA NTRS1971

Sharp flat plate heat transfer in helium at Mach numbers of 22.8 to 86.8 and in corner flow with air at Mach number of 19

Surface heat transfer rates were measured on a sharp flat plate at zero angle of attack in a hypersonic shock tunnel. The density and leading edge Knudsen number were varied to span the continuum to near free molecule regimes. The strong interaction parameter varied from 11 to 16,000 with Knudsen numbers from 0.56 to 17.1 respectively. Local heat transfer rates in the corner flow region produced by the intersection of two perpendicular flat plates with sharp leading edges were determined for various flow densities. The strength of the shock wave from the vertical plate was varied by adjusting the angle of attack from 0 to 5 deg. The unit Reynolds number varied from 1,000 to 17,200 and the Knudsen numbers from 1.6 to 27. The strong interaction parameter varied from 14 to 500.

Nagamatsu, H. T.↗

NASA NTRS1976

Force test of a 0.88 percent scale 142-inch diameter solid rocket booster (MSFC model number 461) in the NASA/MSFC high Reynolds number wind tunnel (SA13F)

The results are presented of a force test of a .88 percent scale model of the 142 inch solid rocket booster without protuberances, conducted in the MSFC high Reynolds number wind tunnel. The objective of this test was to obtain aerodynamic force data over a large range of Reynolds numbers. The test was conducted over a Mach number range from 0.4 to 3.5. Reynolds numbers based on model diameter (1.25 inches) ranged from .75 million to 13.5 million. The angle of attack range was from 35 to 145 degrees.

Johnson, J. D.↗

NASA NTRS1976

An investigation to determine the pressure distribution on the 0.0137 scale solid rocket booster forebody (MSFC model 467) at angles of attack at or near 90 deg and high Reynolds numbers in the MSFC High Reynolds Number Wind Tunnel (SA29F)

An aerodynamic investigation was conducted in the MSFC High Reynolds Number Wind Tunnel to determine the pressure distribution over the foresection of the current 146 inch diameter shuttle SRB. The test model consisted of a 0.0137 scale version of the SRB nose cone and a forward portion of the cylindrical body which was approximately 2.7 calibers in length. The pressure distributions are plotted as a function of longitudinal station ratioed to body diameter and circumferential location for each angle of attack and Mach number. A Reynolds number variation study was made for Mach numbers of 0.4 and 0.6 at an angle of attack of 270 deg and roll angle of 180 deg.

Ramsey, P. E.↗

NASA NTRS1979

Effect of Mach number and Reynolds number on a normal shock-wave/turbulent boundary-layer interaction

Experimental values of wall pressure and skin friction have been obtained for Mach numbers from 1.32 to 1.48 and for Reynolds numbers from 8.5 x 10 to the 6th to 225 x 10 to the 6th. Increasing the Mach number reduces the wall shear and promotes incipient separation. Reynolds number variations have little effect except at the lowest value where the flow abruptly approaches separation. Comparisons are made with solutions to the time-dependent, Reynolds-averaged, Navier-Stokes equations incorporating a two-equation, Wilcox-Rubesin turbulence model. The computations are in agreement with the experimental results. Additional numerical results indicated that the wind-tunnel walls constrained the flow and suppressed the formation of a separation bubble at the shock wave.

Mateer, G. G.↗

NASA NTRS1986

Coherent large-scale structures in high Reynolds number supersonic jet of Mach number 1.4

Large-scale coherent structures are found to exist in fully expanded shock free supersonic jets with Reynolds number 1.6 million and Mach number 1.4. The jet was excited by a weak upstream tone to extract the large-scale structures from the random turbulence. The most preferential excitation Strouhal number is found to be in the vicinity of 0.4, and the dependence of the large-scale structure phase velocity on the excitation Strouhal number shows a similar trend to that found for subsonic acoustically excited jets.

Lepicovsky, J.↗

NASA NTRS1987

A study of the effects of Reynolds number and Mach number on constant pressure coefficient jump for shock-induced trailing-edge separation

The influence of Mach and Reynolds numbers as well as airfoil and planform geometry on the phenomenon of constant shock jump pressure coefficient for conditions of shock induced trailing edge separation (SITES) was studied. It was demonstrated that the phenomenon does exist for a wide variety of two and three dimensional flow cases and that the influence of free stream Mach number was not significant. The influence of Reynolds number was found to be important but was not strong. Airfoil and planform geometric characteristics were found to be very important where the pressure coefficient jump was shown to vary with the sum of: (1) airfoil curvature at the upper surface crest, and (2) camber surface slope at the trailing edge. It was also determined that the onset of SITES could be defined as a function of airfoil geometric parameters and Mach number normal to the leading edge. This onset prediction was shown to predict the angle of onset to within + or - 1 deg accuracy or better for about 90% of the cases studied.

Cunningham, Atlee M., Jr.↗

NASA NTRS1953

The Aerodynamic Design and Calibration of an Asymmetric Variable Mach Number Nozzle with a Sliding Block for the Mach Number Range 1.27 to 2.75

A method of designing as asymmetric, fixed geometry, variable Mach number nozzle has been developed by using the method of characteristics. A small nozzle conforming to the analytically determined ordinates was constructed and calibrated over a range of Mach numbers extending from 1.27 to 2.75. The results show the variation in Mach number to be plus or minus 0.01 or less and in the flow direction to be plus or minus 0.2 degrees within the test section. The range of Mach numbers from 1.27 to 2.75 was obtained by translating the lower block in a straight line parallel to the test-section center line for a distance of 2.17 test-section heights.

NOZZLES↗

NASA NTRS1956

Investigation at Supersonic Speeds of the Variation with Reynolds Number and Mach Number of the Total, Base, and Skin-friction Drag of Seven Boattail Bodies of Revolution Designed for Minimum Wave Drag

Results are presented from an investigation of the variation with Reynolds number and Mach number of the total, base, and skin-friction drag of seven boattail bodies of revolution designed for minimum wave drag according to the theory of NACA-TN-2550. The tests covered a Reynolds number range from approximately 1.0 x 10(6) to 10.0 x 10(6) at Mach numbers of 1.62, 1.93, and 2.41 respectively.

Bromm, August, F , jr↗

NASA NTRS1994

Lewis number and Damkohler number effects in vort ex-flame interactions

A combined experimental-numerical study of the interaction of a two-dimensional vortex pair with a plane premixed laminar flame has been carried out. The selected geometry is two-dimensional in order to allow for the use of quantitative line-of-sight measurement techniques and for comparisons with two-dimensional direct numerical simulations. The experiment is used to identify possible effects of the Lewis number and of radiative heat losses, and direct numerical simulations reproducing the experimental conditions are used to investigate the role of the Lewis number, of heat losses, and of multi-step kinetics. The heat losses considered in this study are essentially radiative losses from the burnt gases and do not include conductive losses to the walls. This report presents the final results of this study of vortex-flame interactions. The main result is the creation of experimental data sets of vortex-flame interactions, which demonstrate the importance of the Lewis and Damkohler numbers to the flame and which allow for comparison with direct numerical simulations. In turn, the comparison between experimental and numerical results shows that the Lewis number effects and complex chemistry play significant roles, which can be reproduced using a two-step reaction mechanism.

Samaniego, Jean-Michael↗

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.↗