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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Stability investigations of airfoil flow by global analysis

As the result of global, non-parallel flow stability analysis the single value of the disturbance growth-rate and respective frequency is obtained. This complex value characterizes the stability of the whole flow configuration and is not referred to any particular flow pattern. The global analysis assures that all the flow elements (wake, boundary and shear layer) are taken into account. The physical phenomena connected with the wake instability are properly reproduced by the global analysis. This enhances the investigations of instability of any 2-D flows, including ones in which the boundary layer instability effects are known to be of dominating importance. Assuming fully 2-D disturbance form, the global linear stability problem is formulated. The system of partial differential equations is solved for the eigenvalues and eigenvectors. The equations, written in the pure stream function formulation, are discretized via FDM using a curvilinear coordinate system. The complex eigenvalues and corresponding eigenvectors are evaluated by an iterative method. The investigations performed for various Reynolds numbers emphasize that the wake instability develops into the Karman vortex street. This phenomenon is shown to be connected with the first mode obtained from the non-parallel flow stability analysis. The higher modes are reflecting different physical phenomena as for example Tollmien-Schlichting waves, originating in the boundary layer and having the tendency to emerge as instabilities for the growing Reynolds number. The investigations are carried out for a circular cylinder, oblong ellipsis and airfoil. It is shown that the onset of the wake instability, the waves in the boundary layer, the shear layer instability are different solutions of the same eigenvalue problem, formulated using the non-parallel theory. The analysis offers large potential possibilities as the generalization of methods used till now for the stability analysis.

Morzynski, Marek↗
NASA NTRS2000

Measurements in a Transitional Boundary Layer Under Low-Pressure Turbine Airfoil Conditions

This report presents the results of an experimental study of transition from laminar to turbulent flow in boundary layers or in shear layers over separation zones on a convex-curved surface which simulates the suction surface of a low-pressure turbine airfoil. Flows with various free-stream turbulence intensity (FSTI) values (0.5%, 2.5% and 10%), and various Reynolds numbers (50,000, 100,000 200,000 and 300,000) are investigated. Reynold numbers in the present study are based on suction surface length and passage exit mean velocity. Flow separation followed by transition within the separated flow region is observed for the lower-Re cases at each of the FSTI levels. At the highest Reynolds numbers and at elevated FSn, transition of the attached boundary layer begins before separation, and the separation zone is small. Transition proceeds in the shear layer over the separation bubble. For both the transitional boundary layer and the transitional shear layer, mean velocity, turbulence intensity and intermittency (the fraction of the time the flow is turbulent) distributions are presented. The present data are compared to published distribution models for bypass transition, intermittency distribution through transition, transition start position, and transition length. A model developed for transition of separated flows is shown to adequately predict the location of the beginning of transition, for these cases, and a model developed for transitional boundary layer flows seems to adequately predict the path of intermittency through transition when the transition start and end are known. These results are useful for the design of low-pressure turbine stages which are known to operate under conditions replicated by these tests.

Simon, Terrence W.↗
NASA NTRS1997

The Flowfield Characteristics of a Mach 2 Diamond Jet

The potential for using a novel diamond-shaped nozzle which may allow for superior mixing characteristics of supersonic jets without significant thrust losses is explored. The results of flow visualization and pressure measurements indicate the presence of distinct structures in the shear layers, not normally observed in shear layers of axisymmetric and rectangular jets. As characteristics of these features suggests that they are a manifestation of significant streamwise vorticity in the shear layers. Despite the distinct nature of the flowfield structure of the present shear layer, the global growth rates of this shear layer were found to be very similar to its two-dimensional and axisymmetric counterparts. These and other observations suggest that the presence of streamwise vorticity may not play a significant role in the global development of a compressible shear layer.

Washington, Donnell↗
NASA NTRS1991

Supersonic mixing enhancement by vorticity for high-speed propulsion

The results of a three year study on vortex enhancement of supersonic mixing are discussed. Recent interests in compressible mixing has spurred research in the field of high speed shear layers. It was established that shear layer growth diminishes with increasing convective Mach number; this Mach number is the relative Mach number of the large scale structures in the shear layer with respect to the Mach numbers on either side of the shear layer. The purpose was to evaluate the effect of swirl on compressible mixing rates. Previously analytical and experimental results seem to indicate that swirling flow may significantly modify the shear layer, in some cases resulting in enhanced mixing. Previous studies of the effect of swirl on compressible mixing were incomplete since the amount of swirl in the flowfield was not quantified. This study was undertaken to conclusively determine the effect of swirl on supersonic mixing, including the quantification of the swirl. Preliminary results indicate that the swirl modestly enhances the mixing rates.

Settles, Gary S.↗
NASA NTRS1981

An experimental study of transmission, reflection and scattering of sound in a free jet flight simulation facility and comparison with theory

When a free jet (or open jet) is used as a wind tunnel to simulate the effects of flight on model noise sources, it is necessary to calibrate out the effects of the free jet shear layer on the transmitted sound, since the shear layer is absent in the real flight case. In this paper, a theoretical calibration procedure for this purpose is first summarized; following this, the results of an experimental program, designed to test the validity of the various components of the calibration procedure, are described. The experiments are conducted by using a point sound source located at various axial positions within the free jet potential core. By using broadband excitation and cross-correlation methods, the angle changes associated with ray paths across the shear layer are first established. Measurements are then made simultaneously inside and outside the free jet along the proper ray paths to determine the amplitude changes across the shear layer. It is shown that both the angle and amplitude changes can be predicted accurately by theory. It is also found that internal reflection at the shear layer is significant only for large ray angles in the forward quadrant where total internal reflection occurs. Finally, the effects of sound absorption and scattering by the shear layer turbulence are also examined experimentally.

Ahuja, K. K.↗
NASA NTRS1980

Lear jet boundary layer/shear layer laser propagation experiments

Optical degradations of aircraft turbulent boundary layers with shear layers generated by aerodynamic fences are analyzed. A collimated 2.5 cm diameter helium-neon laser (0.63 microns) traversed the approximate 5 cm thick natural aircraft boundary layer in double pass via a reflective airfoil. In addition, several flights examined shear layer-induced optical degradation. Flight altitudes ranged from 1.5 to 12 km, while Mach numbers were varied from 0.3 to 0.8. Average line spread function (LSF) and Modulation Transfer Function (MTF) data were obtained by averaging a large number of tilt-removed curves. Fourier transforming the resulting average MTF yields an LSF, thus affording a direct comparison of the two optical measurements. Agreement was good for the aerodynamic fence arrangement, but only fair in the case of a turbulent boundary layer. Values of phase variance inferred from the LSF instrument for a single pass through the random flow and corrected for a large aperture ranged from 0.08 to 0.11 waves (lambda = .63 microns) for the boundary layer. Corresponding values for the fence vary from 0.08 to 0.16 waves. Extrapolation of these values to 10.6 microns suggests negligible degradation for a CO2 laser transmitted through a 5 cm thick, subsonic turbulent boundary layer.

Gilbert, K.↗
NASA NTRS1987

Inlets, ducts and nozzles

The internal fluid mechanics research program in inlets, ducts, and nozzles is described. The program consists of a balanced effort between the development of computational tools and the conduct of experimental research. The experiments are designed to better understand the fluid flow physics, to develop new or improved flow models, and to provide benchmark quality data sets for validation of the computational methods. The inlet, duct, and nozzle research program is described according to three major classifications of flow phenomena: highly three-dimensional flow fields; shock-boundary layer interactions; and shear layer control. Specific examples of current and future elements of the research program are described for each of these phenomena. In particular, the highly three-dimensional flow field phenomena is highlighted by describing the computational and experiemental research program in transition ducts having a round-to-rectangular area variation. In the case of shock-boundary layer interactions, the specific details of research for normal shock-boundary layer interactions are described. For shear layer control research in vortex generators and the use of aerodynamic excitation for enhancement of the jet mixing process are described. Future research in inlets, ducts, and nozzles will include more emphasis on three-dimensional full Navier-Stokes methods and corresponding experiments designed to concentrate on the appropriate three-dimensional fluid flow physics.

Abbott, John M.↗
NASA NTRS2014

Visualization of Capsule Reentry Vehicle Heat Shield Ablation Using Naphthalene PLIF

The Orion Multi-Purpose Crew Vehicle (MPCV) will use an ablative heat shield and improved understanding of the ablation process would be beneficial for design purposes. Given that ablation is a multi-physics process involving heat and mass transfer, codes aiming to predict heat shield ablation are in need of experimental data pertaining to the turbulent transport of ablation products for validation. At The University of Texas at Austin, a technique is being developed that uses planar laser-induced fluorescence (PLIF) of a low-temperature sublimating ablator (naphthalene) to visualize the transport of ablation products in a supersonic flow. Since ablation at reentry temperatures can be difficult to recreate in a laboratory setting it is desirable to create a limited physics problem and simulate the ablation process at relatively low temperature conditions using naphthalene. A scaled Orion MPCV model with a solid naphthalene heat shield has been tested in a Mach 5 wind tunnel at various angles of attack in the current work. PLIF imaging reveals the distribution of the ablation products as they are transported into the heat-shield boundary layer and over the capsule shoulders into the separated shear layer and backshell recirculation region. Visualizations of the capsule shear layer using both naphthalene PLIF and Schlieren imaging compared favorably. High concentrations of naphthalene in the capsule separated flow region, intermittent turbulent structures on the heat shield surface, and interesting details of the capsule shear layer structure were observed using the naphthalene PLIF technique. The capsule shear layer was also shown to generally appear to be more turbulent at lower angles of attack. Furthermore, the PLIF signal increased steadily over the course of a run indicating that during a wind tunnel run the model heated up and the rate of naphthalene ablation increased. The shear layer showed increasing signs of turbulence over the course of a wind tunnel run as well, likely because of the combination of increased surface roughness and surface blowing rate. PLIF imaging also detected regions with a relatively low concentration of naphthalene in the capsule backshell recirculation region that are most likely the result of cross-flow-induced vortices on the capsule afterbody.

Combs, Christopher S.↗
NASA NTRS1996

Blunt body near wake flow field at Mach 6

Tests were conducted in a Mach 6 flow to examine the reattachment process of an axisymmetric free shear layer associated with the near wake of a 70 deg. half angle, spherically blunted cone with a cylindrical after body. Model angle of incidence was fixed at 0 deg. and free-stream Reynolds numbers based on body diameter ranged from 0.5 x 10(exp 6) to 4 x 10(exp 6). The sensitivity of wake shear layer transition on reattachment heating was investigated. The present perfect gas study was designed to compliment results obtained previously in facilities capable of producing real gas effects. The instrumented blunted cone model was designed primarily for testing in high enthalpy hypervelocity shock tunnels in both this country and abroad but was amenable for testing in conventional hypersonic blowdown wind tunnels as well. Surface heating rates were inferred from temperature - time histories from coaxial surface thermocouples on the model forebody and thin film resistance gages along the model base and cylindrical after body. General flow feature (bow shock, wake shear layer, and recompression shock) locations were visually identified by schlieren photography. Mean shear layer position and growth were determined from intrusive pitot pressure surveys. In addition, wake surveys with a constant temperature hot-wire anemometer were utilized to qualitatively characterize the state of the shear layer prior to reattachment. Experimental results were compared to laminar perfect gas predictions provided by a 3-D Navier Stokes code (NSHYP). Shear layer impingement on the instrumented cylindrical after body resulted in a localized heating maximum that was 21 to 29 percent of the forebody stagnation point heating. Peak heating resulting from the reattaching shear layer was found to be a factor of 2 higher than laminar predictions, which suggested a transitional shear layer. Schlieren flow visualization and fluctuating voltage time histories and spectra from the hot wire surveys across the shear layer substantiate this observation. The sensitivity of surface heating to forebody roughness was characterized for a reattaching shear layer. For example, at R(sub infinity), d = 4 x 10(exp 6), when the shear layer was transitional, the magnitude of peak heating from shear layer impingement was reduced by approximately 24 percent when transition grit was applied to the forebody. The spatial location of the local peak, however, remained unchanged.

Horvath, Thomas J.↗
NASA NTRS1996

NPARC Code Upgraded with Two-Equation Turbulence Models

The National PARC (NPARC) Alliance was established by the NASA Lewis Research Center and the Air Force Arnold Engineering Development Center to provide the U.S. aeropropulsion community with a reliable Navier-Stokes code for simulating the nonrotating components of propulsion systems. Recent improvements to the turbulence model capabilities of the NPARC code have significantly improved its capability to simulate turbulent flows. Specifically, the Chien k-epsilon and Wilcox k-omega turbulence models were implemented at Lewis. Lewis researchers installed the Chien k-epsilon model into NPARC to improve the code's ability to calculate turbulent flows with attached wall boundary layers and free shear layers. Calculations with NPARC have demonstrated that the Chien k-epsilon model provides more accurate calculations than those obtained with algebraic models previously available in the code. Grid sensitivity investigations have shown that computational grids must be packed against the solid walls such that the first point off of the wall is placed in the laminar sublayer. In addition, matching the boundary layer and momentum thicknesses entering mixing regions is necessary for an accurate prediction of the free shear-layer growth.

Source record↗
NASA NTRS2012

The Scaling of Broadband Shock-Associated Noise with Increasing Temperature

A physical explanation for the saturation of broadband shock-associated noise (BBSAN) intensity with increasing jet stagnation temperature has eluded investigators. An explanation is proposed for this phenomenon with the use of an acoustic analogy. For this purpose the acoustic analogy of Morris and Miller is examined. To isolate the relevant physics, the scaling of BBSAN at the peak intensity level at the sideline ( = 90 degrees) observer location is examined. Scaling terms are isolated from the acoustic analogy and the result is compared using a convergent nozzle with the experiments of Bridges and Brown and using a convergent-divergent nozzle with the experiments of Kuo, McLaughlin, and Morris at four nozzle pressure ratios in increments of total temperature ratios from one to four. The equivalent source within the framework of the acoustic analogy for BBSAN is based on local field quantities at shock wave shear layer interactions. The equivalent source combined with accurate calculations of the propagation of sound through the jet shear layer, using an adjoint vector Green s function solver of the linearized Euler equations, allows for predictions that retain the scaling with respect to stagnation pressure and allows for the accurate saturation of BBSAN with increasing stagnation temperature. This is a minor change to the source model relative to the previously developed models. The full development of the scaling term is shown. The sources and vector Green s function solver are informed by steady Reynolds-Averaged Navier-Stokes solutions. These solutions are examined as a function of stagnation temperature at the first shock wave shear layer interaction. It is discovered that saturation of BBSAN with increasing jet stagnation temperature occurs due to a balance between the amplification of the sound propagation through the shear layer and the source term scaling.A physical explanation for the saturation of broadband shock-associated noise (BBSAN) intensity with increasing jet stagnation temperature has eluded investigators. An explanation is proposed for this phenomenon with the use of an acoustic analogy. For this purpose the acoustic analogy of Morris and Miller is examined. To isolate the relevant physics, the scaling of BBSAN at the peak intensity level at the sideline psi = 90 degrees) observer location is examined. Scaling terms are isolated from the acoustic analogy and the result is compared using a convergent nozzle with the experiments of Bridges and Brown and using a convergent-divergent nozzle with the experiments of Kuo, McLaughlin, and Morris at four nozzle pressure ratios in increments of total temperature ratios from one to four. The equivalent source within the framework of the acoustic analogy for BBSAN is based on local field quantities at shock wave shear layer interactions. The equivalent source combined with accurate calculations of the propagation of sound through the jet shear layer, using an adjoint vector Green s function solver of the linearized Euler equations, allows for predictions that retain the scaling with respect to stagnation pressure and allows for the accurate saturation of BBSAN with increasing stagnation temperature. This is a minor change to the source model relative to the previously developed models. The full development of the scaling term is shown. The sources and vector Green s function solver are informed by steady Reynolds-Averaged Navier-Stokes solutions. These solutions are examined as a function of stagnation temperature at the first shock wave shear layer interaction. It is discovered that saturation of BBSAN with increasing jet stagnation temperature occurs due to a balance between the amplification of the sound propagation through the shear layer and the source term scaling.

Miller, Steven A.↗
NASA NTRS1987

The effect of flow oscillations on cavity drag

An experimental investigation of flow over an axisymmetric cavity shows that self-sustained, periodic oscillations of the cavity shear layer are associated with low cavity drag. In this low-drag mode the flow regulates itself to fix the mean-shear-layer stagnation point at the downstream corner. Above a critical value of the cavity width-to-depth ratio there is an abrupt and large increase of drag due to the onset of the 'wake mode' of instability. It is also shown by measurement of the momentum balance how the drag of the cavity is related to the state of the shear layer, as defined by the mean momentum transport and the Reynolds stress, and how these are related to the amplifying oscillations in the shear layer. The cavity shear layer is found to be different, in several respects, from a free shear layer.

Gharib, M.↗
NASA NTRSDate not supplied

High-Frequency Secondary Instabilities Downstream of a Forward-Facing Step

Localized high-frequency velocity fluctuations were measured downstream of a forward-facing step on a 30◦swept flat plate with an imposed pressure gradient. Phase-locked time-resolved particle image velocimetry measurements were performed to determine the source of the fluctuations, which ultimately lead to transition. The steps result in localized regions of reversed flow for a short region downstream of the step. These regions are highly localized due to the impact of the stationary crossflow vortices. The velocity fluctuations typically increase significantly in amplitude shortly downstream of reattachment in these localized regions, indicating that the fluctuations are related to reattachment of the shear layer. The phase-averaged measurements confirm that the unsteadiness in the shear layer results in the shedding of a plethora of vortices, some oriented in the streamwise direction, and others in the spanwise direction. These vortices are impacted by the local properties of the mean flow, such as streamwise vorticity and streamwise velocity gradient. The origin of the unsteadiness in the shear layer is not definitively shown, but evidence suggests it is due to a shear-layer instability. Ultimately, the unsteadiness of the shear layer, and the vortices that are shed as a result of this unsteadiness, are responsible for transition in this scenario.

Jenna L Eppink↗
NASA NTRS1997

Noise from Supersonic Coaxial Jets: Inverted Velocity Profile - Part 3

The instability wave noise generation model is used to study the instability waves in the two shear layers of an inverted velocity profile, supersonic, coaxial jet and the noise radiated from the dominant wave. The inverted velocity profile jet has a high speed outer stream surrounding a low speed inner stream and the outer shear layer is always larger than the inner shear layer. The jet mean flows are calculated numerically. The operating conditions are chosen to exemplify the effect of the coaxial jet outer shear layer initial spreading rates. Calculations are made for the stability characteristics in the coaxial jet shear layers and the noise radiated from the instability waves for different operating conditions with the same total thrust, mass flow and exit area as a single reference jet. Results for inverted velocity profile jets indicate that relative maximum instability wave amplitudes and far field peak noise levels can be reduced from that of the reference jet by having higher spreading rates for the outer shear layer, low velocity ratios, and outer streams hotter than the inner stream.

Dahl, Milo D.↗
NASA NTRSDate not supplied

Investigation of the 30P30N Slat Flow Field with Passive Control Devices Using Particle Image Velocimetry

The leading-edge slat of a high-lift wing is one of the main noise contributors during approach and landing. This paper describes an experimental investigation of the velocity field associated with multiple passive noise treatments, including slat-cusp extensions, a cove filler, and a gap filler, on a two-dimensional multi-element high-lift 30P30N airfoil. Previous work documented comparisons of both surface and far field pressure fluctuations in the presence of these devices with those for a baseline case at different flow conditions. However, important information related to the velocity fields was missing from the previous measurements, hindering our ability to elucidate the changes in the flow physics associated with the noise treatments. Therefore, two-component Particle Image Velocity has been used to investigate the influence of passive noise treatments on the flow fields. All measurements are taken at an effective, free-air angle of attack of 5.5 degrees and a stowed-chord-based Reynolds number of 1.71 × 10e6. The measurements show that the slat extensions shorten the slat-cove shear layer trajectory, resulting in reduced growth of disturbances within the slat-cove shear layer. This leads to a shift of tonal peaks to higher frequencies and a reduction in the tonal amplitudes. The gap filler blocks the flow path through the gap, causing the reattachment location to shift to the main wing leading edge lower surface. Consequently, the feedback loop associated with the flow-acoustic interaction in the baseline case is eliminated, and the turbulent kinetic energy in the slat-cove shear layer is significantly reduced. However, extensive flow separation is observed on the suction side of the gap filler, which does not eliminate the noise reduction benefit due to the gap filler, but does degrade the aerodynamic performance of the high-lift configuration. Finally, the overall flow field in the presence of the cove filler is similar to that in the baseline case at the design angle of attack, but the slat-cove shear layer is eliminated leading to a suppression of the cavity tones associated with the shear layer. This change accounts for the reduction in slat noise as measured in previous work.

Aeroacoustics↗
NASA NTRS1990

Inlets, ducts, and nozzles

The internal fluid mechanics research program in inlets, ducts, and nozzles consists of a balanced effort between the development of computational tools (both parabolized Navier-Stokes and full Navier-Stokes) and the conduct of experimental research. The experiments are designed to better understand the fluid flow physics, to develop new or improved flow models, and to provide benchmark quality data sets for validation of the computational methods. The inlet, duct, and nozzle research program is described according to three major classifications of flow phenomena: (1) highly 3-D flow fields; (2) shock-boundary-layer interactions; and (3) shear layer control. Specific examples of current and future elements of the research program are described for each of these phenomenon. In particular, the highly 3-D flow field phenomenon is highlighted by describing the computational and experimental research program in transition ducts having a round-to-rectangular area variation. In the case of shock-boundary-layer interactions, the specific details of research for normal shock-boundary-layer interactions are described. For shear layer control, research in vortex generators and the use of aerodynamic excitation for enhancement of the jet mixing process are described.

Abbott, John M.↗