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At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Three sensor hot wire/film technique for three dimensional mean and turbulence flow field measurement

Methods of measuring the three dimensional flow field using a three-sensor, hot-wire probe is described in this paper, with emphasis on the techniques developed by the author's group at The Pennsylvania State University. The hot wire equations, data processing procedure, calibration techniques, and a discussion of various errors in the measurement are included. Some typical data acquired by this probe is also included.

Lakshminarayana, B.↗
NASA NTRS1987

An investigation of the flow characteristics in the blade endwall corner region

Studies were undertaken to determine the structure of the flow in the blade end wall corner region simulated by attaching two uncambered airfoils on either side of a flat plate with a semicircular leading edge. Detailed measurements of the corner flow were obtained with conventional pressure probes, hot wire anemometry, and flow visualization. The mean velocity profiles and six components of the Reynolds stress tensor were obtained with an inclined single sensor hot wire probe whereas power spectra were obtained with a single sensor oriented normal to the flow. Three streamwise vortices were identified based on the surface streamlines, distortion of total pressure profiles, and variation of mean velocity components in the corner. A horseshoe vortex formed near the leading edge of the airfoil. Within a short distance downstream, a corner vortex was detected between the horseshoe vortex and the surfaces forming the corner. A third vortex was formed at the rear portion of the corner between the corner vortex and the surface of the flat plate. Turbulent shear stress and production of turbulence are negligibly small. A region of negative turbulent shear stress was also observed near the region of low turbulence intensity from the vicinity of the flat plate.

Hazarika, Birinchi K.↗
NASA NTRS1934

Improved Apparatus for the Measurement of Fluctuations of Air Speed in Turbulent Flow

This report describes recent improvements in the design of the equipment associated with the hot-wire anemometer for the measurement of fluctuating air speeds in turbulent air flow, and presents the results of some experimental investigations dealing with the response of the hot wire to speed fluctuations of various frequencies. Attempts at measuring the frequency of the fluctuations encountered in the Bureau of Standards' 54-inch wind tunnel are also reported. In addition, the difficulties encountered in the use of such apparatus and the precautions found helpful in avoiding them are discussed.

Mock, W C , Jr↗
NASA NTRS2015

Validation of a Computational Model for the SLS Core Stage Oxygen Tank Diffuser Concept and the Low Profile Diffuser - An Advanced Development Design for the SLS

The Low Profile Diffuser (LPD) project originated as an award from the Marshall Space Flight Center (MSFC) Advanced Development (ADO) office to the Main Propulsion Systems Branch (ER22). The task was created to develop and test an LPD concept that could produce comparable performance to a larger, traditionally designed, ullage gas diffuser while occupying a smaller volume envelope. Historically, ullage gas diffusers have been large, bulky devices that occupy a significant portion of the propellant tank, decreasing the tank volume available for propellant. Ullage pressurization of spacecraft propellant tanks is required to prevent boil-off of cryogenic propellants and to provide a positive pressure for propellant extraction. To achieve this, ullage gas diffusers must slow hot, high-pressure gas entering a propellant tank from supersonic speeds to only a few meters per second. Decreasing the incoming gas velocity is typically accomplished through expansion to larger areas within the diffuser which has traditionally led to large diffuser lengths. The Fluid Dynamics Branch (ER42) developed and applied advanced Computational Fluid Dynamics (CFD) analysis methods in order to mature the LPD design from and initial concept to an optimized test prototype and to provide extremely accurate pre-test predictions of diffuser performance. Additionally, the diffuser concept for the Core Stage of the Space Launch System (SLS) was analyzed in a short amount of time to guide test data collection efforts of the qualification of the device. CFD analysis of the SLS diffuser design provided new insights into the functioning of the device and was qualitatively validated against hot wire anemometry of the exterior flow field. Rigorous data analysis of the measurements was performed on static and dynamic pressure data, data from two microphones, accelerometers and hot wire anemometry with automated traverse. Feasibility of the LPD concept and validation of the computational model were demonstrated by the test data.

Brodnick, Jacob↗
NASA NTRS1990

Transition and turbulence measurements in hypersonic flows

This paper reviews techniques for transitional- and turbulent-flow measurements and describes current research in support of turbulence modeling. Special attention is given to the potential of applying hot wire and laser velocimeter to measuring turbulent fluctuations in hypersonic flow fields. The results of recent experiments conducted in two hypersonic wind tunnels are presented and compared with previous hot-wire turbulence measurements.

Owen, F. K.↗
NASA NTRS1978

The free shear layer edgetone and instability measurements

The hot-wire probe is known to be capable of triggering and sustaining upstream instability modes like the slit jet-wedge edgetone termed the free shear layer tone. In this paper, the hot-wire probe is used to investigate the shear tone phenomenon in axisymmetric and plane free shear layers. For all the measurements reported, the initial boundary layers producing the free shear layers are documented (at 0.2 cm upstream from the lip) and classified as laminar. It is shown that the free shear layer, inherently different from the fully vortical slit jet, produces a different tone behavior. The data on the free shear layer tone eigenfunctions and eigenvalues suggest that the feedback is hydrodynamic rather than acoustic.

Zaman, K. B. M. Q.↗
NASA NTRS1993

Fluid dynamics and convective heat transfer in impinging jets through implementation of a high resolution liquid crystal technique

A combined convective heat transfer and fluid dynamics investigation in a turbulent round jet impinging on a flat surface is presented. The experimental study uses a high resolution liquid crystal technique for the determination of the convective heat transfer coefficients on the impingement plate. The heat transfer experiments are performed using a transient heat transfer method. The mean flow and the character of turbulent flow in the free jet is presented through five hole probe and hot wire measurements, respectively. The flow field character of the region near the impingement plate plays an important role in the amount of convective heat transfer. Detailed surveys obtained from five hole probe and hot wire measurements are provided. An extensive validation of the liquid crystal based heat transfer method against a conventional technique is also presented. After a complete documentation of the mean and turbulent flow field, the convective heat transfer coefficient distributions on the impingement plate are presented. The near wall of the impingement plate and the free jet region is treated separately. The current heat transfer distributions are compared to other studies available from the literature. The present paper contains complete sets of information on the three dimensional mean flow, turbulent velocity fluctuations, and convective heat transfer to the plate. The experiments also prove that the present nonintrusive heat transfer method is highly effective in obtaining high resolution heat transfer maps with a heat transfer coefficient uncertainty of 5.7 percent.

Kim, K.↗
NASA NTRS1998

Drag Force Anemometer Used in Supersonic Flow

To measure the drag on a flat cantilever beam exposed transversely to a flow field, the drag force anemometer (beam probe) uses strain gauges attached on opposite sides of the base of the beam. This is in contrast to the hot wire anemometer, which depends for its operation on the variation of the convective heat transfer coefficient with velocity. The beam probe retains the high-frequency response (up to 100 kHz) of the hot wire anemometer, but it is more rugged, uses simpler electronics, is relatively easy to calibrate, is inherently temperature compensated, and can be used in supersonic flow. The output of the probe is proportional to the velocity head of the flow, 1/2 rho u(exp 2) (where rho is the fluid density and u is the fluid velocity). By adding a static pressure tap and a thermocouple to measure total temperature, one can determine the Mach number, static temperature, density, and velocity of the flow.

Fralick, Gustave C.↗
NASA NTRS1995

Turbulence measurements in hypersonic boundary layers using constant-temperature anemometry and Reynolds stress measurements in hypersonic boundary layers

The primary objective in the two research investigations performed under NASA Langley sponsorship (Turbulence measurements in hypersonic boundary layers using constant temperature anemometry and Reynolds stress measurements in hypersonic boundary layers) has been to increase the understanding of the physics of hypersonic turbulent boundary layers. The study began with an extension of constant-temperature thermal anemometry techniques to a Mach 11 helium flow, including careful examinations of hot-wire construction techniques, system response, and system calibration. This was followed by the application of these techniques to the exploration of a Mach 11 helium turbulent boundary layer (To approximately 290 K). The data that was acquired over the course of more than two years consists of instantaneous streamwise mass flux measurements at a frequency response of about 500 kHz. The data are of exceptional quality in both the time and frequency domain and possess a high degree of repeatability. The data analysis that has been performed to date has added significantly to the body of knowledge on hypersonic turbulence, and the data reduction is continuing. An attempt was then made to extend these thermal anemometry techniques to higher enthalpy flows, starting with a Mach 6 air flow with a stagnation temperature just above that needed to prevent liquefaction (To approximately 475 F). Conventional hot-wire anemometry proved to be inadequate for the selected high-temperature, high dynamic pressure flow, with frequent wire breakage and poor system frequency response. The use of hot-film anemometry has since been investigated for these higher-enthalpy, severe environment flows. The difficulty with using hot-film probes for dynamic (turbulence) measurements is associated with construction limitations and conduction of heat into the film substrate. Work continues under a NASA GSRP grant on the development of a hot film probe that overcomes these shortcomings for hypersonic flows. Each of the research tasks performed during the NASA Langley research grants is discussed separately below.

Spina, Eric F.↗
NASA NTRS1989

An experimental investigation of a low Reynolds number turbulent boundary layer subject to an adverse pressure gradient

A very low Reynolds number turbulent boundary layer subject to an adverse pressure gradient is studied. The aim is to obtain highly accurate mean-flow and turbulence measurements under conditions that can be closely related to the numerical simulations of Philippe Spalart for the purposes of CFD validation. Much of the Boundary Layer Wind Tunnel was completely rebuilt with a new wider contraction and working section which will improve compatibility with the simulations. A unique sophisticated high-speed computer controlled 3-D probe traversing mechanism was integrated into the test section. Construction of the tunnel and traverse is discussed in some detail. The hardware is now complete, and measurements are in progress. The mean-flow data indicate that a suitably two-dimensional base flow was established. Automation of the probe positioning and data acquistion have led to a decreased running time for total pressure measurements. However, the most significant benefits are expected to occur when using hot-wire probes. Calibrations can be performed automatically and there is no need to handle fragile probes when moving between measuring stations. Techniques are being developed which require sampling of the signals from moving hot-wire probes on the basis of their position in the flow. Measurements can be made in high intensity turbulence by flying probes upstream at high speed so that the relative magnitude of the turbulent velocity fluctuations are reduced. In regions, where the turbulence intensity is not too large, the probe can also be repetitively scanned across very dense spatial grids in other directions. With this technique, a complete profile can be measured in about 1/3 the time and with a spatial density about 50 times that obtainable using a stationary probe.

Watmuff, Jonathan H.↗
NASA NTRS1972

An experimental investigation of turbulent boundary layers along curved surfaces

A curved wall tunnel was designed, and an equilibrium turbulent boundary layer was set up on the straight section preceding the curved test section. Turbulent boundary layer flows with uniform and adverse pressure distributions along convex and concave walls were investigated. Hot-wire measurements along the convex surface indicated that turbulent mixing between fluid layers was very much reduced. However, the law of the wall held and the skin friction, thus determined, correlated well with other measurements. Hot-wire measurements along the concave test wall revealed a system of longitudinal vortices inside the boundary layer and confirmed that concave curvature enhances mixing. A self-consistent set of turbulent boundary layer equations for flows along curved surfaces was derived together with a modified eddy viscosity. Solution of these equations together with the modified eddy viscosity gave results that correlated well with the present data on flows along the convex surface with arbitrary pressure distribution. However, it could only be used to predict the mean characteristics of the flow along concave walls because of the existence of the system of longitudinal vortices inside the boundary layer.

So, R. M. C.↗
NASA NTRS2019

Flow Quality Survey of the 9- by 15-Foot Low Speed Wind Tunnel (2015 and 2016 Tests) Prior to Acoustic Improvement Modifications

A flow quality survey (FQS) was conducted in December 2015 and January 2016 to characterize the flow quality of the 9- by 15-Foot Low-Speed Wind Tunnel (9×15 LSWT) test section prior to the start of significant structural and acoustic modifications. The data from this FQS will be compared with data collected during a replicated FQS after completion of the acoustic improvement modifications. This document contains only the FQS data collected in December 2015 and January 2016 using the 15-ft survey rake and the 18-in. boundary layer rakes in the 9×15 LSWT. The spatial variation at one axial test section station (TSTA), nominally TSTA 198, and five vertical heights of the 15-ft survey rake was computed for Mach number, velocity, total and static pressure, total temperature, and flow angularity (pitch and yaw). Three hot-wire probes mounted to the 15-ft survey rake were used to measure turbulence intensity; however, data from a previous test entry were used as the mean turbulence intensity baseline value, as there were issues with the hot-wire performance. The boundary layer pressure profiles around the perimeter of the test section at TSTA 199 were also measured and the boundary layer heights computed. All calculated results are tabulated for each facility condition set point.

Johnson, Aaron M.↗
NASA NTRS1974

Turbulent measurements in supersonic boundary layer flows using laser velocimetry

Laser velocimetry has been applied successfully in the measurement of mean velocities and Reynolds stresses in a zero pressure gradient, Mach 2.9 boundary layer, and upstream and downstream of a two-dimensional, shock-wave boundary layer interaction for the same Mach number. The Reynolds stresses were obtained using a one-component, dual-scatter laser velocimeter operated in a manner analogous to the 'slanted hot-wire' technique. Signal processing was of the single-particle counting type which permitted measurements to be made using only the naturally occurring particles in the tunnel air stream for light scattering. The results for the shock-wave boundary layer interaction are presented along with hot-wire anemometer and pitot-static pressure measurements obtained in the same flow. Also, a data acquisition scheme for use with a two-component laser velocimeter is presented which should provide even better accuracies in future studies.

Johnson, D. A.↗
NASA NTRS1991

Experimental investigation of turbulent flow through a circular-to-rectangular transition duct

Steady, incompressible, turbulent, swirl-free flow through a circular-to-rectangular transition duck was studied experimentally. The cross-sectional area remains the same at the exit as at the inlet, but varies through the transition section to a maximum value approximately 15 percent above the inlet value. The cross-sectional geometry everywhere along the duct is defined by the equation of a superellipse. Mean and turbulence data were accumulated utilizing pressure and hot-wire instrumentation at five stations along the test section. Data are presented for operating bulk Reynolds numbers of 88,000 and 390,000. Measured quantities include total and static pressure, the three components of the mean velocity vector, and the six components of the Reynolds stress tensor. In addition to the transition duct measurements, a hot-wire technique which relies on the sequential use of single rotatable normal and slant-wire probes was proposed. The technique is applicable for measurement of the total mean velocity vector and the complete Reynolds stress tensor when the primary flow is arbitrarily skewed relative to a plane which lies normal to the probe axis of rotation.

Davis, David O.↗
NASA NTRS1972

Experimental study of nozzle wall boundary layers at Mach numbers 20 to 47

The nozzle wall boundary layer of an M-50 helium tunnel was investigated with pitot pressure, total temperature, skin friction, and wall heat transfer measurements at five stations and hot wire measurements at two stations. The results indicated that the boundary layer was turbulent with a thick viscous sublayer. Pressure gradients were observed across the boundary layer; the effect of these gradients on the equations of motion are discussed. The direct skin friction measurements were higher than expected from empirical predictions; the Reynolds analogy factors however were lower than expected. Hot wire measurements indicated mass flow fluctuations as large as 80 percent of the local mean mass flow at the edge of the viscous sublayer with a maximum value relative to the edge mass flow of about 15 percent at gamma about equal to 0.8.

Kemp, J. H., Jr.↗
NASA NTRS1981

Effects of a modified leading edge on noise and boundary-layer transition in a rod-wall sound shield at Mach 5

A version of a rod wall sound shield was tested in the Mach 5 pilot quiet tunnel over a range of unit Reynolds numbers from 10 to 35 million per meter. The model was modified by inclining the leading edge plates to produce an initial 2 deg expansion to ascertain the sensitivity of boundary layer transition to leading edge disturbances. Rod surface pitot pressures, mean free stream pitot pressures, and static pressures on the rods and plenum walls were measured. Hot-wire measurements were also made in the model and nozzle free stream at a unit Reynolds number of 15 million per meter. The surface pitot pressures indicated that transition was much farther forward than for the previous tests due to the leading edge modification and minor fabrication flaws in the model. Early boundary layer transition on the rods was confirmed by hot-wire measurements which showed much higher noise levels in the free stream shield flow when compared with results from previous tests. Mean pitot pressure surveys within the shielded region inside the model indicated that there was an overexpansion and recompression that would limit the streamwise length of undisturbed flow to about 13 cm along the centerline.

Creel, T. R., Jr.↗
NASA NTRS1939

Air Flow in the Boundary Layer of an Elliptic Cylinder

The boundary layer of an elliptic cylinder of major and minor axis 11.78 and 3.98 inches, respectively, was investigated in air stream in which the turbulence could be varied. Conditions were arranged so that the flow was two-dimensional with the major axis of the ellipse parallel to the undisturbed stream. Speed distributions across the boundary layer were determined with a hot-wire anemometer at a number of positions about the surface for the lowest and highest intensities of turbulence, with the air speed in both cases sufficiently high to produce a turbulent boundary layer over the downstream part of the surface. The magnitude and the frequency of the speed fluctuations in the boundary layer were also measured by the use of the conventional type of hot-wire turbulence apparatus. Stream turbulence was found to affect both the nature of transition from laminar to turbulent flow in the layer and the position on the surface at which transition occurred. Transition was then investigated in detail with stream turbulence of several different scales and intensities.

Schubauer, G B↗
NASA NTRS1980

Experimental investigation of the asymmetric body vortex wake

An experimental investigation of the asymmetric body vortex wake of a circular cylinder in high subsonic flow is presented. Laser velocimeter, force and moment, and surface hot wire measurements were obtained for a freestream Mach number of 0.6 and Reynolds number (based on body diameter) of 0.62 x 10 to the 6th. Two component laser velocimeter measurements were made at three body cross-flow planes, x/d = 4, 8, and 12, and angles of attack of 25, 35, and 45 deg. Laser vapor screen photographs were also obtained at these body stations and angles of attack. Surface hot wire measurements were used to determine if any vortex switching occurred at various angles of attack of the body. The laser velocimeter measurements are related to the vapor screen photographs and side force measurements. These results show that more than one asymmetric body vortex wake configuration can exist for the same angle of attack and body roll angle.

Oberkampf, W. L.↗