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NASA NTRS1961

Low Subsonic Pressure Distributions on Three Rigid Wings Simulating Paragliders with Varied Canopy Curvature and Leading-Edge Sweep

An investigation has been made in the Langley 7- by 10-foot transonic tunnel to determine the subsonic pressure distribution of three paraglider models through an angle-of-attack range from 0 deg to 74 deg. Three rigid metal models simulated a 45 deg basic flat planform paraglider with leading-edge sweep angles of 61.6 deg, 52.5 deg, and 48.6 deg. These configurations resulted in one-half-circle, one-third-circle, and one-quarter-circle semispan trailing-edge curvature when viewed from downstream. The results of the investigation are presented as curves of chordwise pressure distributions at four spanwise locations.

Fournier, Paul G.↗

NASA NTRS1980

The prediction of pressure distributions on an arrow-wing configuration including the effect of camber, twist, and a wing fin

Wind tunnel tests of an arrow wing body configuration consisting of flat, twisted, and cambered twisted wings were conducted at Mach numbers from 0.40 to 2.50 to provide an experimental data base for comparison with theoretical methods. A variety of leading and trailing edge control surface deflections were included in these tests, and in addition, the cambered twisted wing was tested with an outboard vertical fin to determine its effect on wing and control surface loads. Theory experiment comparisons show that current state of the art linear and nonlinear attached flow methods were adequate at small angles of attack typical of cruise conditions. The incremental effects of outboard fin, wing twist, and wing camber are most accurately predicted by the advanced panel method PANAIR. Results of the advanced panel separated flow method, obtained with an early version of the program, show promise that accurate detailed pressure predictions may soon be possible for an aeroelasticity deformed wing at high angles of attack.

Bobbitt, P. J.↗

NASA NTRS2004

Electron Density and Two-Channel Neutron Emission Measurements in Steady-State Spherical Inertial-Electrostatically Confined Plasmas, with Review of the One-Dimensional Kinetic Model

Electron density measurements have been made in steady-state plasmas in a spherical inertial electrostatic confinement (IEC) discharge using microwave interferometry. Plasma cores interior to two cathodes, having diameters of 15 and 23 cm, respectively, were probed over a transverse range of 10 cm with a spatial resolution of about 1.4 cm for buffer gas pressures from 0.2 to 6 Pa in argon and deuterium. The transverse profiles are generally flat, in some cases with eccentric symmetric minima, and give mean densities of from approx. = 0.4 to 7 x 10(exp 10)/cu cm, the density generally increasing with the neutral gas pressure. Numerical solutions of the one-dimensional Poisson equation for IEC plasmas are reviewed and energy distribution functions are identified which give flat transverse profiles. These functions are used with the plasma approximation to obtain solutions which also give densities consistent with the measurements, and a double potential well solution is obtained which has minima qualitatively similar to those observed. Explicit consideration is given to the compatibility of the solutions interior and exterior to the cathode, and to grid transparency. Deuterium fusion neutron emission rates were also measured and found to be isotropic, to within the measurement error, over two simultaneous directions. Anisotropy was observed in residual emissions during operation with nonfusing hydrogen- 1. The deuterium rates are consistent with predictions from the model.

Dobson, Chris C.↗

NASA NTRS1999

Heat Transfer and Flow on the First Stage Blade Tip of a Power Generation Gas Turbine: Experimental Results - Part 1

A combined computational and experimental study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first stage blade tip surface for a geometry typical of large power generation turbines(>100MW). This paper is concerned with the design and execution of the experimental portion of the study. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57 x 10(exp 6), and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5% or 9%. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp-edge and rounded-edge tip geometries at each of the inlet turbulence intensity levels.

Bunker, Ronald S.↗

NASA NTRS1970

Flow characteristics of an air jet impinging on a flat surface

Survey develops adequate heat transfer correlations for design use. Flow characteristics studies include - potential core length, velocity and pressure distribution through the jet, and spread of jet and velocity decay along jet axis.

Gauntner, J. W.↗

NASA NTRS1997

Profiled Roller Stress/Fatigue Life Analysis Methodology and Establishment of an Appropriate Stress/Life Exponent

The objective of this work was to determine the three dimensional volumetric stress field, surface pressure distribution and actual contact area between a 0.50" square roller with different crown profiles and a flat raceway surface using Finite Element Analysis. The 3-dimensional stress field data was used in conjunction with several bearing fatigue life theories to extract appropriate values for stress-life exponents. Also, results of the FEA runs were used to evaluate the laminated roller model presently used for stress and life prediction.

Source record↗

NASA NTRS1976

Supersonic pressure measurements and comparison of theory to experiment for an arrow-wing configuration

A wind tunnel test of an arrow-wing-body configuration consisting of flat and twisted wings, as well as leading- and trailing-edge control surface deflections, was conducted at Mach numbers from 1.54 to 2.50 to provide an experimental pressure data base for comparison with theoretical methods. Theory-to-experiment comparisons of detailed pressure distributions were made using a state-of-the-art inviscid flow, constant-pressure-panel method. Emphasis was on conditions under which this theory is valid for both flat and twisted wings.

Manro, M. E.↗

NASA NTRS1980

Controlled supercritical crossflow on supersonic wings - An experimental validation

Leading edges of wings with high lift at supersonic speeds are discussed. Under these conditions, the leading edge must generate a controlled supercritical crossflow such that high lift is obtained without boundary layer separation. The leading edge concept that accomplishes this is the supersonic counterpart of the leading edge of the 2-D supercritical airfoil at transonic speeds. Wind tunnel test results on a cambered wing with a controlled supercritical crossflow leading edge successfully validated the concept. The main results are presented at the M = 1.62 design point, with off-design results given for the range of Mach numbers from 1.60 to 2.00. Surface pressure, oil flow photographs, and force data are presented and analyzed. A flat wing with the same thickness distribution was also tested and the results were used both for comparison with the designed wing and to illustrate the development of the supercritical crossflow region which, for the flat wing, is terminated by a distinct crossflow shock wave.

Mason, W. H.↗

NASA NTRS1985

Unsteady analysis of rotor blade tip flow

The development of the VSAERO-TS and VSAERO-H computer programs for calculating the unsteady aerodynamic characteristics of arbitrarily shaped wings oscillating in pitch is presented. The effect of several wake parameters on chordwise pressure distribution in VSAERO-TS is given and the convergence characteristics of both programs are discussed. In the program, the influence coefficient for each panel is formulated for a planar surface and so a skewed panel is represented by a projected flat quadrilateral lying in the mean plane. Since panels in the extreme roll-up region of the tip vortex are highly skewed the program was modified to treat each highly skewed panel as a pair of triangles. The programs are validated by comparing the chordwise pressure distribution of several blade tip planforms with experimental data. The comparison, for the most part, is good. The triangular panel representation improved the chordwise pressure distribution near the tip region for higher mean angle of attack.

Maskey, B.↗

NASA NTRS1974

A parametric study of effect of forebody shape on flow angularity at Mach 8

Flow angularity and static pressure measurements have been made on the lower surface of nine forebody models that simulate the bottom forward surface of a hypersonic aircraft. Measurements were made in an area of the forebody that represents the location of an inlet of a scramjet engine. A parametric variation of the forebody surface investigated the effect of: (1) spanwise curvature; (2) longitudinal curvature; and (3) planform shape on both flow angularity and static pressure distribution. Results of each of the three parametric variations of geometry were compared to those for the same flat delta forebody. Spanwise curvature results showed that a concave shape and the flat delta had the lowest flow angularity and lowest rate of increase in flow angularity with angle of attack. Longitudinal curvature results showed a convex surface to give the better flow at the higher angles of attack. The better of the two planform shapes tested was a convex elliptical shape. Limited flow field calculations were made at angles of attack using a three dimensional, method-of-characteristics program. In general, at all angles of attack there was agreement between data and theory.

Johnson, C. B.↗

NASA NTRS1978

Estimation of leading-edge thrust for supersonic wings of arbitrary planform

A numerical method for the estimation of leading edge thrust for supersonic wings of arbitrary planform was developed and was programmed as an extension to an existing high speed digital computer method for prediction of wing pressure distributions. The accuracy of the method was assessed by comparison with linearized theory results for a series of flat delta wings. Application of the method to wings of arbitrary planform, both flat and cambered, is illustrated in several examples.

Carlson, H. W.↗

NASA NTRS1975

Transonic pressure measurements and comparison of theory to experiment for an arrow-wing configuration. Volume 1: Experimental data report, base configuration and effects of wing twist and leading-edge configuration

A wind tunnel test of an arrow-wing-body configuration consisting of flat and twisted wings, as well as a variety of leading- and trailing-edge control surface deflections, was conducted at Mach numbers from 0.4 to 1.1 to provide an experimental pressure data base for comparison with theoretical methods. Theory-to-experiment comparisons of detailed pressure distributions were made using current state-of-the-art attached and separated flow methods. The purpose of these comparisons was to delineate conditions under which these theories are valid for both flat and twisted wings and to explore the use of empirical methods to correct the theoretical methods where theory is deficient.

Manro, M. E.↗

NASA NTRS1993

Flowfield measurements about a multi-element airfoil at high Reynolds numbers

This paper describes experimental data obtained with a multi-element airfoil at flight Reynolds numbers and lift coefficients including Clmax. The wind tunnel test was conducted in the NASA Langley Low Turbulence Pressure Tunnel as part of a cooperative effort between McDonnell Douglas Aerospace and NASA Langley. The airfoil model is a supercritical design configured with a leading-edge slat and a single-segment trailing-edge flap. Data include surface static-pressure distributions (integrated to obtain lift), drag data obtained with wake-rake surveys, and fbwfield surveys obtained with a flat-tube and five-hole probe at nine stations on the configuration's upper surface. Effects of variations in Reynolds number and flap gap on airfoil performance and flowfield survey data are presented.

Chin, Vincent D.↗

NASA NTRS1974

Pressure measurements on the leading edge of a swept wing at Mach 2.2

Detailed pressure measurements were made on a flat semispan swept wing with a rounded leading edge at Mach number 2.2 through a range of Reynolds numbers. Pressure orifices were distributed in the streamwise direction at five spanwise stations on the leading edge and on the upper and lower surfaces. No significant amount of leading-edge suction was found, but the pressures and integrated normal forces on the upper and lower surfaces indicate the presence of a vortex lift.

Sorrells, R. B., III↗

NASA NTRS1981

Pressure and force data for a flat wing and a warped conical wing having a shockless recompression at Mach 1.62

A conical nonlinear flow computer code was used to design a warped (cambered) wing which would produce a supercritical expansion and shockless recompression of the crossflow at a lift coefficient of 0.457, an angle of attack of 10 deg, and a Mach number of 1.62. This cambered wing and a flat wing the same thickness distribution were tested over a range of Mach numbers from 1.6 to 2.0. For both models the forward 60 percent is purely conical geometry. Results obtained with the cambered wing demonstrated the design features of a supercritical expansion and a shockless recompression, whereas results obtained with the flat wing indicated the presence of crossflow shocks. Tables of experimental pressure, force, and moment data are included, as well as selected oil flow photographs.

Miller, D. S.↗

NASA NTRS1992

The development of a turbulent junction vortex system

The growth and development of a horseshoe vortex system in an incompressible, three-dimensional turbulent junction flow were investigated experimentally. A streamlined cylinder mounted with its axis normal to a flat surface was used to generate the junction vortex flow. The flow environment was characterized by a body Reynolds number of 183,000, based on the leading edge diameter of the streamlined cylinder. The study included surface flow visualizations, surface pressure measurements, and mean flow measurements of total pressure, static pressure, and velocity distributions in three planes around the base of the streamlined cylinder, and in two planes in the wake flow. Some characterizations of vortex properties based on the measured mean cross-flow velocity components are presented. The results show the presence of a single large, dominant vortex, with strong evidence of a very small corner vortex in the junction between the cylinder and the flat surface. The center of the dominant vortex drifts away from both the body and the flat surface as the flow develops along and downstream of the body. The growth and development of the core of the large, dominant vortex are documented.

Pierce, F. J.↗

NASA NTRS2000

Flap Edge Aeroacoustic Measurements and Predictions

An aeroacoustic model test has been conducted to investigate the mechanisms of sound generation on high-lift wing configurations. This paper presents an analysis of flap side-edge noise, which is often the most dominant source. A model of a main element wing section with a half-span flap was tested at low speeds of up to a Mach number of 0.17, corresponding to a wing chord Reynolds number of approximately 1.7 million. Results are presented for flat (or blunt), flanged, and round flap-edge geometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database is obtained from a Small Aperture Directional Array (SADA) of microphones, which was constructed to electronically steer to different regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edge aerodynamics is established by static surface pressure data, as well as by Computational Fluid Dynamics (CFD) calculations and simplified edge flow analyses. Distributions of unsteady pressure sensors over the flap allow the noise source regions to be defined and quantified via cross-spectral diagnostics using the SADA output. It is found that shear layer instability and related pressure scatter is the primary noise mechanism. For the flat edge flap, two noise prediction methods based on unsteady-surface-pressure measurements are evaluated and compared to measured noise. One is a new causality spectral approach developed here. The other is a new application of an edge-noise scatter prediction method. The good comparisons for both approaches suggest that much of the physics is captured by the prediction models. Areas of disagreement appear to reveal when the assumed edge noise mechanism does not fully define, the noise production. For the different edge conditions, extensive spectra and directivity are presented. Significantly, for each edge configuration, the spectra for different flow speeds, flap angles, and surface roughness were successfully scaled by utilizing aerodynamic performance and boundary layer scaling method developed herein.

Brooks, Thomas F.↗

NASA NTRS2000

Flap Edge Aeroacoustic Measurements and Predictions

An aeroacoustic model test has been conducted to investigate the mechanisms of sound generation on high-lift wing configurations. This paper presents an analysis of flap side-edge noise, which is often the most dominant source. A model of a main element wing section with a half-span flap was tested at low speeds of up to a Mach number of 0.17, corresponding to a wing chord Reynolds number of approximately 1.7 million. Results are presented for flat (or blunt), flanged, and round flap-edge geometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database is obtained from a Small Aperture Directional Array (SADA) of microphones, which was constructed to electronically steer to different regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edge aerodynamics is established by static surface pressure data, as well as by Computational Fluid Dynamics (CFD) calculations and simplified edge flow analyses. Distributions of unsteady pressure sensors over the flap allow the noise source regions to be defined and quantified via cross-spectral diagnostics using the SADA output. It is found that shear layer instability and related pressure scatter is the primary noise mechanism. For the flat edge flap, two noise prediction methods based on unsteady surface pressure measurements are evaluated and compared to measured noise. One is a new causality spectral approach developed here. The other is a new application of an edge-noise scatter prediction method. The good comparisons for both approaches suggest that much of the physics is captured by the prediction models. Areas of disagreement appear to reveal when the assumed edge noise mechanism does not fully define the noise production. For the different edge conditions, extensive spectra and directivity are presented. Significantly, for each edge configuration, the spectra for different flow speeds, flap angles, and surface roughness were successfully scaled by utilizing aerodynamic performance and boundary layer scaling methods developed herein.

Brooks, Thomas F.↗