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At least 505 records · Page 28

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NASA NTRS1959

Aerodynamic Loading Characteristics Including Effects of Aeroelasticity of a Thin-Trapezoidal-Wing-Body Combination at Mach Number of 1.43

Results have been obtained in the Langley 8-foot transonic pressure tunnel at a Mach number of 1.43 and at angles of attack from 0 deg to about 24 deg which indicate the static-aerodynamic-loads characteristics for a 2-percent-thick trapezoidal wing in combination with a body. Included are the effects of changing Reynolds number and of fixing boundary-layer transition. The results show that aerodynamic loading characteristics at a Mach number of 1.43 are similar to those reported in NACA RM L56Jl2a for the same configuration at a Mach number of 1.115. Reducing the Reynolds number resulted in reductions in the deflection of the wing and caused a slight increase in the relative loading over the outboard wing sections since the deflections were in a direction to unload the tip sections. Little or no effects were seen to result from fixing boundary-layer transition at a tunnel stagnation pressure of 1,950 pounds per square foot.

Kelly, Thomas C.↗

NASA NTRS1959

Flow Characteristics About Two Thin Wings of Low Aspect Ratio Determined from Surface Pressure Measurements Obtained in Flight at Mach Numbers from 0.73 to 1.90

Surface pressure measurements were obtained at three chordwise stations on the wings of the X-3 and X-lE airplanes at Mach numbers from 0.73 to 1.13 for the X-3, and from 0.82 to 1.90 for the X-IE. Leading-edge separation is present on the X-3 wing at a Mach number of about 0.73 and an angle of attack of about 6 deg. However., when the Mach number is increased to 0.88, the trailing-edge separation dominates the pressure distribution and no leading-edge separation is visible although it is anticipated at the higher angles of attack shown. Conversely, the X-lE wing shows no indication of leading-edge separation within the scope of this investigation, but an overexpansion immediately behind the leading edge is present at a Mach number of approximately 0.82. Two separate normal shocks are present on the X-3 wing at a Mach number of about 0.88 and at a low angle of attack as an effect of wing geometry. These shocks merge to form a single shock when the angle of attack is increased to about 6 deg. At supersonic speeds the upper-surface expansion on the X-lE wing is limited by the approach of the pressure coefficients to the pressure coefficient for a vacuum.

Taillon, Norman V.↗

NASA NTRS1959

Estimation of Static Longitudinal Stability of Aircraft Configurations at High Mach Numbers and at Angles of Attack Between 0 deg and +/-180 deg

The possibility of obtaining useful estimates of the static longitudinal stability of aircraft flying at high supersonic Mach numbers at angles of attack between 0 and +/-180 deg is explored. Existing theories, empirical formulas, and graphical procedures are employed to estimate the normal-force and pitching-moment characteristics of an example airplane configuration consisting of an ogive-cylinder body, trapezoidal wing, and cruciform trapezoidal tail. Existing wind-tunnel data for this configuration at a Mach number of 6.86 provide an evaluation of the estimates up to an angle of attack of 35 deg. Evaluation at higher angles of attack is afforded by data obtained from wind-tunnel tests made with the same configuration at angles of attack between 30 and 150 deg at five Mach numbers between 2.5 and 3.55. Over the ranges of Mach numbers and angles of attack investigated, predictions of normal force and center-of-pressure locations for the configuration considered agree well with those obtained experimentally, particularly at the higher Mach numbers.

Dugan, Duane W.↗

NASA NTRS1958

Heat-Transfer Measurements on a 5.5- Inch-Diameter Hemispherical Concave Nose in Free Flight at Mach Numbers up to 6.6

The aerodynamic heat transfer to a hemispherical concave nose has been measured in free flight at Mach numbers from 3.5 to 6.6 with corresponding Reynolds numbers based on nose diameter from 7.4 x 10(exp 6) to 14 x 10(exp 6). Over the test Mach number range the heating on the cup nose, expressed as a ratio to the theoretical stagnation-point heating on a hemisphere nose of the same diameter, varied from 0.05 to 0.13 at the stagnation point of the cup, was approximately 0.1 at other locations within 40 deg of the stagnation point, and varied from 0.6 to 0.8 just inside the lip where the highest heating rates occurred. At a Mach number of 5 the total heat input integrated over the surface of the cup nose including the lip was 0.55 times the theoretical value for a hemisphere nose with laminar boundary layer and 0.76 times that for a flat face. The heating at the stagnation point was approximately 1/5 as great as steady-flow tunnel results. Extremely high heating rates at the stagnation point (on the order of 30 times the stagnation-point values of the present test), which have occurred in conjunction with unsteady oscillatory flow around cup noses in wind-tunnel tests at Mach and Reynolds numbers within the present test range, were not observed.

Levine, Jack↗

NASA NTRS1961

Aerodynamic Effects of Some Configuration Variables on the Aeroelastic Characteristics of Lifting Surfaces at Mach Numbers from 0.7 to 6.86

Results of flutter tests on some simple all-movable-control-type models are given. One set of models, which had a square planform with double-wedge airfoils with four different values of leading- and trailing-edge radii from 0 to 6 percent chord and airfoil thicknesses of 9, 11, 14, and 20 percent chord, was tested at Mach numbers from 0.7 to 6.86. The bending-to-torsion frequency ratio was about 0.33. The other set of models, which had a tapered planform with single-wedge and double-wedge airfoils with thicknesses of 3, 6, 9, and 12 percent chord, was tested at Mach numbers from 0.7 to 3.98 and a frequency ratio of about 0.42. The tests indicate that, in general, increasing thickness has a destabilizing effect at the higher Mach numbers but is stabilizing at subsonic and transonic Mach numbers. Double-wedge airfoils are more prone to flutter than single-wedge airfoils at comparable stiffness levels. Increasing airfoil bluntness has a stabilizing effect on the flutter boundary at supersonic speeds but has a negligible effect at subsonic speeds. However, increasing bluntness may also lead to divergence at supersonic speeds. Results of calculations using second-order piston-theory aerodynamics in conjunction with a coupled-mode analysis and an uncoupled-mode analysis are compared with the experimental results for the sharp-edge airfoils at supersonic speeds. The uncoupled-mode analysis more accurately predicted the flutter characteristics of the tapered-planform models, whereas the coupled-mode analysis was somewhat better for the square-planform models. For both the uncoupled- and coupled-mode analyses, agreement with the experimental results improved with increasing Mach number. In general, both methods of analysis gave unconservative results with respect to the experimental flutter boundaries.

Hanson, Perry W.↗

NASA NTRS1961

Aerodynamic Loading Characteristics at Mach Numbers from 0.80 to 1.20 of a 1/10-Scale Three-Stage Scout Model

Aerodynamic loads results have been obtained in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.80 to 1.20 for a 1/10-scale model of the upper three stages of the Scout vehicle. Tests were conducted through an angle-of-attack range from -8 deg to 8 deg at an average test Reynolds number per foot of about 4.0 x 10(exp 6). Results indicated that the peak negative pressures associated with expansion corners at the nose and transition flare exhibit sizeable variations which occur over a relatively small Mach number range. The magnitude of the variations may cause the critical local loading condition for the full-scale vehicle to occur at a Mach number considerably lower than that at which the maximum dynamic pressure occurs in flight. The addition of protuberances simulating antennas and wiring conduits had slight, localized effects. The lift carryover from the nose and transition flare on the cylindrical portions of the model generally increased with an increase in Mach number.

Kelly, Thomas C.↗

NASA NTRS1958

Effects of Body Shape on the Drag of a 45 degree Sweptback-Wing-Body Configuration at Mach Numbers from 0.90 to 1.43

An investigation was made of the effects of body shape on the drag of a 45 deg sweptback-wing-body combination at Mach numbers from 0.90 to 1.43. Both the expansion and compression fields induced by body indentation were swept back as the stream Mach number increased from 0.94. The line of zero pressure change was generally tangent to the Mach lines associated with the local velocities over the wing and body. The strength of the induced pressure fields over the wing were attenuated with spanwise distance and the major effects were limited to the inboard 60 percent of the wing semispan. Asymmetrical body indentation tended to increase the lift on the forward portion of the wing and reduce the lift on the rearward portion. This redistribution of lift had a favorable effect on the wave drag due to lift. Symmetrical body indentation reduced the drag loading near the wing-body juncture at all Mach numbers. The reduction in drag loading increased in spanwise extent as the Mach number increased and the line of zero induced pressure became more nearly aligned with the line of maximum wing thickness. Calculations of the wave drag due to thickness, the wave drag due to lift, and the vortex drag of the basic and symmetrical M = 1.2 body and wing combinations at an angle of attack of 0 deg predicted the effects of indentation within 11 percent of the wing-basic-body drag throughout the Mach number range from 1.0 to 1.43. Calculations of the wave drag due to thickness, the wave drag due to lift, and the vortex drag for the basic, symmetrical M = 1.2, and asymmetrical M = 1.4 body and wing combinations predicted the total pressure drag to within 8 percent of the experimental value at M = 1.43.

Olstad, Walter B.↗

NASA NTRS1998

Preston Probe Calibrations at High Reynolds Number

The overall goal of the research effort is to study the performance of two Preston probes designed by NASA Langley Research Center across an unprecedented range of Reynolds number (based on friction velocity and probe diameter), and perform an accurate calibration over the same Reynolds number range. Using the Superpipe facility in Princeton, two rounds of experiments were performed. In each round of experiments for each Reynolds number, the pressure gradient, static pressure from the Preston probes and the total pressure from the Preston probes were measured. In the first round, 3 Preston probes having outer diameters of 0.058 inches, 0.083 inches and 0.203 inches were tested over a large range of pipe Reynolds numbers. Two data reduction methods were employed: first, the static pressure measured on the Preston probe was used to calculate P (modified Preston probe configuration), and secondly, the static pressure measured at the reference pressure tap was used to calculate P (un-modified Preston probe configuration). For both methods, the static pressure was adjusted to correspond with the static pressure at the Preston probe tip using the pressure gradient. The measurements for Preston probes with diameters of 0.058 inches, and 0.083 inches respectively were performed in the test pipe before it was polished a second time. Therefore, the measurements at high pipe Reynolds numbers may have been affected by roughness. In the second round of experiments the 0.058 inches and 0.083 inches diameter, un-modified probes were tested after the pipe was polished and prepared to ensure that the surface was smooth. The average velocity was estimated by assuming that the connection between the centerline velocity and the average velocity was known, and by using a Pitot tube to measure the centerline velocity. A preliminary error estimate suggests that it is possible to introduce a 1% to 2% error in estimating the average velocity using this approach. The evidence on the errors attending the second data set is somewhat circumstantial, and the measurements have not been repeated using a better approach, it seems probable that the correlation given applies to un-modified Preston probes over the range 6.4 less than x* less than 11.3.

Smits, Alexander J.↗

NASA NTRS2000

Sweep and Compressibility Effects on Active Separation Control at High Reynolds Numbers

This paper explores the effects of compressibility, sweep and excitation location on active separation control at high Reynolds numbers. The model, which was tested in a cryogenic pressurized wind tunnel, simulates the upper surface of a 20% thick Glauert Goldschmied type airfoil at zero angle of attack. The flow is fully turbulent since the tunnel sidewall boundary layer flows over the model. Without control, the flow separates at the highly convex area and a large turbulent separation bubble is formed. Periodic excitation is applied to gradually eliminate the separation bubble. Two alternative blowing slot locations as well as the effect of compressibility, sweep and steady suction or blowing were studied. During the test the Reynolds numbers ranged from 2 to 40 million and Mach numbers ranged from 0.2 to 0.7. Sweep angles were 0 and 30 deg. It was found that excitation must be introduced slightly upstream of the separation region regardless of the sweep angle at low Mach number. Introduction of excitation upstream of the shock wave is more effective than at its foot. Compressibility reduces the ability of steady mass transfer and periodic excitation to control the separation bubble but excitation has an effect on the integral parameters, which is similar to that observed in low Mach numbers. The conventional swept flow scaling is valid for fully and even partially attached flow, but different scaling is required for the separated 3D flow. The effectiveness of the active control is not reduced by sweep. Detailed flow field dynamics are described in the accompanying paper.

Seifert, Avi↗

NASA NTRS2000

Sweep and Compressibility Effects on Active Separation Control at High Reynolds Numbers

This paper explores the effects of compressibility, sweep and excitation location on active separation control at high Reynolds numbers. The model, which was tested in a cryogenic pressurized wind tunnel, simulates the upper surface of a 20% thick GlauertGoldschmied type airfoil at zero angle of attack. The flow is fully turbulent since the tunnel sidewall boundary layer flows over the model. Without control, the flow separates at the highly convex area and a large turbulent separation bubble is formed. Periodic excitation is applied to gradually eliminate the separation bubble. Two alternative blowing slot locations as well as the effect of compressibility, sweep and steady suction or blowing were studied. During the test the Reynolds numbers ranged from 2 to 40 million and Mach numbers ranged from 0.2 to 0.7. Sweep angles were 0 and 30 deg. It was found that excitation must be introduced slightly upstream of the separation region regardless of the sweep angle at low Mach number. Introduction of excitation upstream of the shock wave is more effective than at its foot. Compressibility reduces the ability of steady mass transfer and periodic excitation to control the separation bubble but excitation has an effect on the integral parameters, which is similar to that observed in low Mach numbers. The conventional swept flow scaling is valid for fully and even partially attached flow, but different scaling is required for the separated 3D flow. The effectiveness of the active control is not reduced by sweep. Detailed flow field dynamics are described in the accompanying paper.

Seifert, Avi↗

NASA NTRS2002

Investigation of Reynolds Number Effects on a Generic Fighter Configuration in the National Transonic Facility

A wind tunnel test of a generic fighter configuration was tested in the National Transonic Facility through a cooperative agreement between NASA Langley Research Center and McDonnell Douglas. The primary purpose of the test was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The test included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: (1) Fuselage/Wing; (2) Fuselage/Wing/Centerline Vertical Tail/Horizontal Tail; and (3) Fuselage/Wing/Trailing-Edge Extension/Twin Vertical Tails. Reynolds number effects on the longitudinal aerodynamic characteristics are presented herein.

Tomek, W. G.↗

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 NTRS2002

Determination of the Absolute Number of Cytokine mRNA Molecules within Individual Activated Human T Cells

A primary function of activated T cells is the expression and subsequent secretion of cytokines, which orchestrate the differentiation of other lymphocytes, modulate antigen presenting cell activity, and alter vascular endothelium to mediate an immune response. Since many features of immune regulation probably result from modest alterations of endogenous rates of multiple interacting processes, quantitative analysis of the frequency and specific activity of individual T cells is critically important. Using a coordinated set of quantitative methods, the absolute number of molecules of several key cytokine mRNA species in individual T cells has been determined. The frequency of human blood T cells activated in vitro by mitogens and recall protein antigens was determined by intracellular cytokine protein staining, in situ hybridization for cytokine mRNA, and by limiting dilution analysis for cytokine mRNA+ cells. The absolute number of mRNA molecules was simultaneously determined in both homogenates of the entire population of cells and in individual cells obtained by limiting dilution, using a quantitative, competitive RT-PCR assay. The absolute numbers of mRNA molecules in a population of cells divided by the frequency of individual positive cells, yielded essentially the same number of mRNA molecules per cell as direct analysis of individual cells by limiting dilution analysis. Mean numbers of mRNA per positive cell from both mitogen and antigen activated T cells, using these stimulation conditions, were 6000 for IL-2, 6300 for IFN-gamma, and 1600 for IL-4.

Karr, Laurel J.↗

NASA NTRS1961

Pressure Distributions and Wave Drag Due to Two-Dimensional Fabrication-Type Surface Roughness on an Ogive Cylinder at Mach Numbers of 1.61 and 2.01

An investigation has been made at Mach numbers of 1.61 and 2.01 and over a range of free-stream Reynolds number per foot from about 1.2 x 10(exp 6) to 8.3 x 10(exp 6) to determine the pressure distributions and wave drags due to two-dimensional fabrication-type surface roughness. Ten types of surface roughness, including step, wave, crease, and swept configurations were investigated. The tests were made on an ogive cylinder of fineness ratio 12.2, the roughness elements covering the cylindrical portion of the model. The results indicate that wave drag is the major component of the drag due to roughness at supersonic speeds. The pressure distributions over the roughness elements were generally found to be in good agreement with linearized two-dimensional theory except for regions of the elements affected by boundary-layer separation and shock detachment. There was little or no effect of Reynolds number except on the pressures within the regions influenced by separation or shock detachment. Inasmuch as most of the roughness configurations were affected by flow separation and shock detachment, there was generally an effect of Reynolds number on the roughness wave drag. This wave drag decreased as the free-stream Reynolds number was decreased.

Czarnecki, K. R.↗

NASA NTRS1961

A Wind-Tunnel Investigation of a Transonic-Transport Configuration Utilizing Drag-Reducing Devices at Mach Numbers from 0.20 to 1.03

The static longitudinal stability and control and lateral characteristics of a transonic-transport model, incorporating recent drag-reducing devices, has been investigated in the Langley 8-foot transonic pressure tunnel. The wing was cambered, had a thickened root and a taper ratio of 0.3. Wing sweepback angles of 45 degrees and 40 degrees were investigated with corresponding aspect ratios of 7 and 8, respectively. Modifications to the model for reducing the drag were: a forward fuselage addition and special bodies (four big enough to house jet engines) added to the upper surface of the wing. Other components and changes investigated included an empennage, a wing-tip body, wing fences, wing trailing-edge flaps, horizontal-tail settings, and wing dihedral angle. The investigation covered the Mach number range from 0.20 to 1.03 for the angle-of-attack range from -5 degrees to 15.4 degrees, and a sideslip angle of -5 degrees, in the Reynolds number range from 0.52 times 10(exp 6) to 1.94 times 10(exp 6) based on the wing mean aerodynamic chord. The various fuselage and wing additions delayed the drag-rise Mach number and greatly reduced the drag beyond the drag rise. The wing bodies markedly alleviated unstable pitch tendencies throughout the test Mach number range. At low landing speeds, the wing bodies exhibited little interference with the ability of trailing-edge flaps to increase the lift near maximum lift coefficient; and the use of fences greatly reduced the severe longitudinal instability trend at landing attitudes. The model with a 6 degree dihedral angle exhibited positive lateral and directional stability characteristics in the presence of the fuselage and wing additions. An increase in drag-rise Mach number associated with the fuselage and wing additions on the 40 degree sweptback wing combination was similar to that for the comparable 45 degree combination. These additions did, however, reduce the drag of the 40 degree sweptback configurations more than the 45 degree configurations in the transonic speed range.

Loving, Donald L.↗

NASA NTRS2003

MHD Turbulence at Moderate Magnetic Reynolds Number

In the present article we will consider the decay of MHD turbulence under the influence of a strong external magnetic field at moderate magnetic Reynolds numbers. Typical values of R(sub m) that are considered here range from R(sub m) approx. 0.1 to R(sub m) approx. 20. As a comparison, the initial kinetic Reynolds number common to all our simulations is Re(sub L) = 199. This means that the range of Prandtl numbers explored is 5 x 10(exp -4) to 10(exp -1). Our motivation is mainly to exhibit how the transition from the QS approximation to FMHD occurs. At the lowest values of R(sub m) studied here, the QS approximation is shown to model the flow faithfully. However, for the higher values of R(sub m) considered, it is clearly inadequate but can be replaced by another approximation which will be referred to as the Quasi-Linear (QL) approximation. Another objective of the present study is to describe how variations in the magnetic Reynolds number (while maintaining all other parameters constant) affect the dynamics of the flow. This complements past studies where variations in either the strength of the external magnetic field or the kinetic Reynolds number were considered. This article is organized as follows. In section 2 we recall the definition of the quasi-static approximation. Section 3 is devoted to the description of the numerical experiments performed using the quasi-static approximation and full MHD. In section 4 we describe the quasi-linear approximation and test it numerically against full MHD. A concluding summary is given in section 5.

Knaepen, B.↗

NASA NTRS1961

Free-Flight Skin-Temperature and Surface-Pressure Measurements on a Highly Polished Nose Having a 100 deg Total-Angle Cone and a 10 deg Half-Angle Conical Flare Section up to a Mach Number of 4.08

The skin temperature and surface pressure were measured on a large-scale, highly polished nose having a relatively sharp-tipped 100 deg total-angle cone followed by a conical flare section of 10 deg half-angle. The measurements were obtained in flight from a rocket-propelled model up to a peak Mach number of 4.08 and a peak Reynolds number of 22 x 10(exp 6) per foot. Temperature distributions indicated that the heating on the forward 3.5 inches of the 100 deg cone was lower than the heating on the rearward portion. Likewise, measured temperatures on the flare portion of the test nose were generally lower than the temperatures on the 100 deg cone portion. The data indicated that the local Reynolds numbers of transition, based on calculated boundary-layer momentum thicknesses, ranged from 530 to 940 for a Mach number range from 2.72 to 3.75. Comparison of measured cone pressures with theory for a sharp cone showed that theory overestimates the cone pressures. Pressure measurements on the flare portion of the nose showed that in the lower speed range the flow expands below atmospheric pressure in going from the cone to the flare; however, as the speed increased, the expansion diminished and for speeds greater than a Mach number of approximately 3.0 the flare pressure coefficients were at or near a value of zero.

Rashis, Bernard↗

NASA NTRS1999

Oscillatory Excitation of Unsteady Compressible Flows over Airfoils at Flight Reynolds Numbers

An experimental investigation, aimed at delaying flow separation due to the occurrence of a shock-wave-boundary-layer interaction, is reported. The experiment was performed using a NACA 0012 airfoil and a NACA 0015 airfoil at high Reynolds number incompressible and compressible flow conditions. The effects of Mach and Reynolds numbers were identified, using the capabilities of the cryogenic-pressurized facility to maintain one parameter fixed and change the other. Significant Reynolds number effects were identified in the baseline compressible flow conditions even at Reynolds number of 10 and 20 million. The main objectives of the experiment were to study the effects of periodic excitation on airfoil drag-divergence and to alleviate the severe unsteadiness associated with shock-induced separation (known as "buffeting"). Zero-mass-flux oscillatory blowing was introduced through a downstream directed slot located at 10% chord on the upper surface of the NACA 0015 airfoil. The effective frequencies generated 2-4 vortices over the separated region, regardless of the Mach number. Even though the excitation was introduced upstream of the shock-wave, due to experimental limitations, it had pronounced effects downstream of it. Wake deficit (associated with drag) and unsteadiness (associated with buffeting) were significantly reduced. The spectral content of the wake pressure fluctuations indicates of steadier flow throughout the frequency range when excitation was applied. This is especially important at low frequencies which are more likely to interact with the airframe.

Seifert, Avi↗