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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Skin-Friction Measurements in a 3-D, Supersonic Shock-Wave/Boundary-Layer Interaction

The experimental documentation of a three-dimensional shock-wave/boundary-layer interaction in a nominal Mach 3 cylinder, aligned with the free-stream flow, and 20 deg. half-angle conical flare offset 1.27 cm from the cylinder centerline. Surface oil flow, laser light sheet illumination, and schlieren were used to document the flow topology. The data includes surface-pressure and skin-friction measurements. A laser interferometric skin friction data. Included in the skin-friction data are measurements within separated regions and three-dimensional measurements in highly-swept regions. The skin-friction data will be particularly valuable in turbulence modeling and computational fluid dynamics validation.

Wideman, J. K.↗
NASA NTRS1972

Sideband growth in nonlinear Landau wave-particle interaction.

The distortion of the electron velocity distribution caused by a large amplitude Landau wave is determined analytically for the initial-value problem. The resulting stability of electrostatic perturbations impressed on the evolving plasma is studied. Narrow sidebands of the applied frequency experience consecutive growths of large magnitude during the early stages of the nonlinear wave-particle interaction. The significance of the derived results to both wave propagation experiments and triggered VLF emissions in the magnetosphere is discussed.

Brinca, A. L.↗
NASA NTRS1986

The PROUST radar: First results

Two campaigns took place in 1984 with the PROUST Radar operating in a bistatic mode, the transmitting antenna pointing at the vertical and the receiving one, 1 deg. off the vertical axis. The antenna beam intersection covers an altitude range between 3 and 9 km. The first of these campaigns are analyzed. The results analyzed show the capability of the PROUST Radar to measure the turbulent parameters and study the turbulence-wave interaction. In its present configuration (bistatic mode and 600 m vertical resolution), it has been necessary to make some assumptions that are known not to be truly fulfilled: homogeneous turbulence and constant vertical wind intensity over a 600-m thickness. It is clear that a more detailed study of the interaction between wave and turbulence will be possible with the next version of PROUST Radar (30-m altitude resolution and monostatic mode) that will soon be achieved.

Bertin, F.↗
NASA NTRS1986

Interferometric data for a shock-wave/boundary-layer interaction

An experimental study of the axisymmetric shock-wave / boundary-layer strong interaction flow generated in the vicinity of a cylinder-cone intersection was conducted. The study data are useful in the documentation and understanding of compressible turbulent strong interaction flows, and are part of a more general effort to improve turbulence modeling for compressible two- and three-dimensional strong viscous/inviscid interactions. The nominal free stream Mach number was 2.85. Tunnel total pressures of 1.7 and 3.4 atm provided Reynolds number values of 18 x 10(6) and 36 x 10(6) based on model length. Three cone angles were studied giving negligible, incipient, and large scale flow separation. The initial cylinder boundary layer upstream of the interaction had a thickness of 1.0 cm. The subsonic layer of the cylinder boundary layer was quite thin, and in all cases, the shock wave penetrated a significant portion of the boundary layer. Owing to the thickness of the cylinder boundary layer, considerable structural detail was resolved for the three shock-wave / boundary-layer interaction cases considered. The primary emphasis was on the application of the holographic interferometry technique. The density field was deduced from an interferometric analysis based on the Able transform. Supporting data were obtained using a 2-D laser velocimeter, as well as mean wall pressure and oil flow measurements. The attached flow case was observed to be steady, while the separated cases exhibited shock unsteadiness. Comparisons with Navier-Stokes computations using a two-equation turbulence model are presented.

Dunagan, Stephen E.↗
NASA NTRS2015

CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave/Boundary-Layer Interaction

Experimental investigations of specific flow phenomena, e.g., Shock Wave Boundary-Layer Interactions (SWBLI), provide great insight to the flow behavior but often lack the necessary details to be useful as CFD validation experiments. Reasons include: 1.Undefined boundary conditions Inconsistent results 2.Undocumented 3D effects (CL only measurements) 3.Lack of uncertainty analysis While there are a number of good subsonic experimental investigations that are sufficiently documented to be considered test cases for CFD and turbulence model validation, the number of supersonic and hypersonic cases is much less. This was highlighted by Settles and Dodsons [1] comprehensive review of available supersonic and hypersonic experimental studies. In all, several hundred studies were considered for their database.Of these, over a hundred were subjected to rigorous acceptance criteria. Based on their criteria, only 19 (12 supersonic, 7 hypersonic) were considered of sufficient quality to be used for validation purposes. Aeschliman and Oberkampf [2] recognized the need to develop a specific methodology for experimental studies intended specifically for validation purposes.

Turbulence Modeling↗
NASA NTRS1994

Experimental studies of shock-wave/wall-jet interaction in hypersonic flow

Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not significantly reduced by even the largest levels of film cooling. For the case studies in the absence of film cooling, the interaction regions were unseparated. However, adding film cooling resulted in regions of boundary layer separation induced in the film cooling layer -- the size of which regions first increased and then decreased with increased film cooling. Surprisingly, the size of the separated regions and the magnitude of the recompression heating were not strongly influenced by the thickness of the cooling film, nor by the point of shock impingement relative to the exit plane of the nozzles. The lip thickness was found to have little effect on cooling effectiveness. Measurements with and in the absence of shock interaction were compared with the results of earlier experimental studies and correlated in terms of the major parameters controlling these flows.

Holden, Michael S.↗
NASA NTRS1994

Experimental studies of shock-wave/wall-jet interaction in hypersonic flow, part A

Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident-shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not significantly reduced by even the largest levels of film cooling. For the case studies in the absence of film cooling, the interaction regions were unseparated. However, adding film cooling resulted in regions of boundary layer separation induced in the film cooling layer, the size of which regions first increased and then decreased with increased film cooling. Surprisingly, the size of the separated regions and the magnitude of the recompression heating were not strongly influenced by the thickness of the cooling film, nor by the point of shock impingement relative to the exit plane of the nozzles. The lip thickness was found to have little effect on cooling effectiveness. Measurements with and in the absence of shock interaction were compared with the results of earlier experimental studies and correlated in terms of the major parameters controlling these flows.

Holden, Michael S.↗
NASA NTRS1991

Evidence of gravity wave-tidal interaction observed near the summer mesopause at Poker Flat, Alaska

An analysis of gravity wave-tidal interaction observed near the mesopause by the MST radar at Poker Flat in July of 1986 is presented. The observations revealed daily mean wind maxima of about 60 m/sec westward and 20 m/sec southward with daily mean momentum fluxes, contributed by gravity waves with periods less than 1 hour of 4-5 sq m/sec sq eastward and 1-2 sq m/sec sq northward. Considerable hourly height variability was found to exist for both winds and momentum fluxes. A significant modulation of the fluxes by tidal winds was observed, characterized by out-of-phase correlations over a number of heights.

Wang, Ding-Yi↗
NASA NTRSDate not supplied

Numerical Simulations of an Axisymmetric Shock-Wave/Boundary-Layer Interaction

A series of Reynolds-averaged Navier-Stokes (RANS) simulations were performed of two Mach 2.5 axisymmetric shock-wave/boundary-layer interactions (SWBLIs) based off of the NASA Turbulent Computational Fluid Dynamics Validation Experiment (TCFDVE). The goals of the simulations were to (1) determine how well RANS can predict the flow physics of an axisymmetric SWBLI and (2) generate a new validation case for the NASA Turbulence Modeling Resource (TMR). The simulations were run with two different turbulence models and three different computational fluid dynamics (CFD) codes while looking at two different conical shock generator angles, one of which was shown to cause flow separation in the experiment. The results in this paper show that, as expected, the CFD simulations agreed well with each other and the experiment when SWBLI-induced separation was not present. When separation was present, differences in the predicted separation characteristics were noted. These included a stronger interaction and subsequently longer separation extent for the Menter Shear-Stress Transport k-ω (SST) turbulence model simulations compared to both the Spalart-Allmaras (SA) turbulence model simulations and the experiment.

Computational Fluid Dynamics↗
NASA NTRSDate not supplied

Shock-Wave/Boundary-Layer Interaction (SWBLI) Experiments in the Presence of Transition-to-Turbulence on a Flat Plate Model in the NASA LaRC 31-Inch Mach 10 Air Tunnel

Historically, there have been a limited number of studies involving Transitional Shock-Wave/Boundary-Layer Interactions (XSWBLI) in hypersonic flows, the majority of which have been performed at Mach Numbers of 7.5 and below. Additionally, there have been even fewer such experiments which incorporate non-intrusive and optical diagnostics. NASA Langley Research Center (NASA LaRC), in collaboration with The University of Texas at San Antonio (UTSA) and The University of Tennessee Space Institute (UTSI), is currently planning and preparing for a series of experimental aerodynamic tests using the NASA LaRC 31-inch Mach 10 Air Tunnel. The experiments will primarily focus on studying the dynamics of Shock-Wave/Boundary-Layer Interactions (SWBLI) in the presence of transition (XSWBLI) and turbulence (SWTBLI) on a large, flat plate model at a freestream Mach number of 10. Besides continued interest in understanding the flow at high Mach numbers, executing such tests at Mach 10 avoids some known concerns and will aid in solving new problems. XSWBLI are a highly unsteady phenomena and the generation of XSWBLI at lower Mach numbers can prove to be a significant challenge. This concern will be somewhat mitigated in the Mach 10 flow, as boundary layers become increasingly stabilized, but some difficulty in achieving transition is still expected. Modelling SWBLI in the presence of transitional and turbulent boundary layers has also proven to be difficult, so a high-Mach number experiment which applies non-intrusive and optical diagnostics will aid in solving a unique problem as well as advancing the understanding and characterization of an aerodynamic surface at Mach 10.The NASA LaRC 31-inch Wind Tunnel incorporates a test section having a cross-section of 31-inches × 31-inches and provides optical access to the test section via three (3) ultra-violet (UV) transmitting windows on the top, side, and bottom. The Mach 10 operating conditions will consist of pressures (P0) ranging from 2.4 to 10MPa (348 to 1,450 psi) at a temperature (T0) of 1,000 K (1,800 °R). The NASA LaRC 31-inch wind tunnel will deliver 1-minute blow-down runs, of which 30-45 seconds will be consumed to reach the aforementioned test conditions. The pump down time between runs will be approximately 45-60 minutes, which will provide for 6-10 test runs per day over the course of a weeklong entry. The proposed flat plate model will be constructed of stainless steel with geometry consisting of a10-inch × 30-inch upper surface and a 2-inch thickness. The model will be designed to support multiple leading-edge inserts in order to examine blunt, round, and sharp leading edges at Mach 10. Adjustments will be made to manage boundary layer thickness as well as the strength and size of the shock interaction region for each test. Measurements and diagnostics will be performed though high-speed Schlieren, on-body high-speed pressure transducers, IR thermography, and oil flow visualization. Hypersonic vehicle applications have and will continue to emerge at the forefront of aerospace. The NASA-UTSA-UTSI team realizes the value of characterizing SWBLI, XSWBLI, and SWTBLI behaviors well beyond single-digit hypersonic Mach numbers. As hypersonic vehicle speed capabilities continue to increase, measurement and diagnostic methods for Mach numbers of 10 and beyond will provide the foundation for a firm understanding of flow field behavior which will directly influence the advancement of technologies towards the design and manufacturing of high-speed aerodynamic surfaces, controls, thermal protection systems, acoustic treatments, and structural components. The use of non-intrusive and optical diagnostic methods in such experimentation is pivotal in developing the visualization and empirical data necessary to advance the aforementioned technology areas.

Shockwave/Boundary-Layer Interactions↗
NASA NTRS1988

Ocean wave directional spectra and wave-current interaction in the Agulhas from the Shuttle Imaging Radar-B synthetic aperture radar

SIR-B synthetic aperture radar ocean-wave spectra for a 200-km pass crossing the Agulhas current off the coast of Africa are analyzed. A significant enhancement of one spectral peak along the northern edge of the current is attributed both to amplification of the waves by refraction-dominated wave-current interaction and to transient specular backscatter contributions spatially correlated to the waves.

Irvine, D. E.↗
NASA NTRS1990

Wave-particle interactions for relativistic electrons in a recirculation acceleration model

Evidence is presented for the presence of a multistep process, leading to relativistic electron populations, in the magnetospheres of earth and Jupiter. After an initial injection of a spectrally soft electron component in the outer magnetosphere, inward radial diffusion acts to accelerate the electrons perpendicular to the local magnetic field; strong wave-particle interactions deep in the magnetosphere is of primary significance in the pitch-angle scattering of electrons, thereby lowering their mirror points. An additional model element of critical significance, again involving wave-particle interactions, is energy-preserving outward transport near the mirror points where the field strength is large and trans-L distances are small.

Baker, D. N.↗
NASA NTRS1994

Laser interferometer skin-friction measurements of crossing-shock-wave/turbulent-boundary-layer interactions

Wall shear stress measurements beneath crossing-shock-wave/turbulent boundary-layer interactions have been made for three interactions of different strengths. The interactions are generated by two sharp fins at symetric angles of attack mounted on a flat plate. The shear stress measurements were made for fin angles of 7 and 11 deg at Mach 3 and 15 deg at Mach 3.85. The measurements were made using a laser interferometer skin-friction meter, a device that determines the wall shear by optically measuring the time rate of thinning of an oil film placed on the test model surface. Results of the measurements reveal high skin-friction coefficients in the vicinity of the fin/plate junction and the presence of quasi-two-dimensional flow separation on the interaction center line. Additionally, two Navier-Stokes computations, one using a Baldwin-Lomax turbulence model and one using a k-epsilon model, are compared with the experimental results for the Mach 3.85, 15-deg interaction case. Although the k-epsilon model did a reasonable job of predicting the overall trend in portions of the skin-friction distribution, neither computation fully captured the physics of the near-surface flow in this complex interaction.

Garrison, T. J.↗