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At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Reynolds Number Effects on the Richtmyer-Meshkov Instability

This presentation compares the results of two very different experimental studies of Richtmyer-Meshkov instability: shock tube experiments in which an air/SF6 interface is accelerated by a weak shock wave; and incompressible experiments in which a box containing two different density miscible liquids is impulsively accelerated by bouncing it off of a fixed coil spring. Both experiments are initiated with sinusoidal initial perturbations. The interface perturbation initially remains sinusoidal as it grows in amplitude, but eventually the interfacial vorticity concentrates into points, forming a row of line vortices of alternating sign. The Reynolds number based on vortex circulation ranges from 1,000 to 45,000 in these experiments. It is found that viscous effects have a large, quantifiable effect on the evolution of the individual vortices. The effects of viscosity on the overall perturbation amplitude, however, are small and will be compared to theory.

Niederhaus, Charles↗
NASA NTRS2000

High Resolution Experiments of Compressible Turbulence Interacting With Shock Waves

Most of the second year of our research program focused on exploring the potentially favorable the effects of expansion waves on homogeneous and isotropic turbulence, which is formed downstream of a grid. Expansion waves are associated with compressible flows and may reduce the drag over airfoils by suppressing turbulence. In the very few previous investigations of interactions of turbulence with expansion waves the effects due to stabilizing streamline curvature substantially masked the effects of turbulence suppression due to flow expansion though the waves. In the present flow configuration planar expansion waves interact with grid generated turbulence in our high-resolution shock tube research facility. This approach will assess directly the effects of the interaction on turbulence. The first objective of our study was to identify the nature of expansion waves present in our shock tube facility. Our time-dependent numerical simulations of the flow in our facility indicated the existence of two regions of traveling expansion waves. The system of expansion waves utilized in this investigation is generated by the exiting shock wave and the induced flow behind it at the end of the driver. Several new measuring techniques are being developed which are capable of providing velocity-gradient-related quantities in compressible flows for the first time.

Agui, Juan↗
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 NTRS1991

Estimates of thermochemical relaxation lengths behind normal shock waves relevant to manned lunar and Mars return missions, the aeroassist flight experiment, and Mars entry

Thermochemical relaxation distances behind the strong normal shock waves associated with vehicles that enter the Earth atmosphere upon returning from a manned lunar or Mars mission are estimated. The relaxation distances for a Mars entry are estimated as well, in order to highlight the extent of the relaxation phenomena early in currently envisioned space exploration studies. The thermochemical relaxation length for the Aeroassist Flight Experiment is also considered. These estimates provide an indication as to whether finite relaxation needs to be considered in subsequent detailed analyses. For the Mars entry, relaxation phenomena that are fully coupled to the flow field equations are used. The relaxation-distance estimates can be scaled to flight conditions other than those discussed.

Howe, John T.↗
NASA NTRS1989

Transient hot-film sensor response in a shock tube

Shock tube experiments were performed to determine the response of a hot-film sensor, mounted flush on the side wall of a shock tube, to unsteady flow behind a normal shock wave. The present experiments attempt to isolate the response of the anemometer due only to the change in convective heat transfer at the hot-film surface. The experiments, performed at low supersonic shock speeds in air, are described along with the data acquisition procedure. The change in convective heat transfer is deduced from the data and the results are compared with those from transient boundary layer theory and another set of experimental results. Finally, a transient local heat transfer coefficient is formulated for use as the forcing function in a hot-film sensor instrument model simulation.

Roberts, A. S., Jr.↗
NASA NTRS2008

Experimental Plans for Subsystems of a Shock Wave Driven Gas Core Reactor

This Contractor Report proposes a number of plans for experiments on subsystems of a shock wave driven pulsed magnetic induction gas core reactor (PMI-GCR, or PMD-GCR pulsed magnet driven gas core reactor). Computer models of shock generation and collision in a large-scale PMI-GCR shock tube have been performed. Based upon the simulation results a number of issues arose that can only be addressed adequately by capturing experimental data on high pressure (approx.1 atmosphere or greater) partial plasma shock wave effects in large bore shock tubes ( 10 cm radius). There are three main subsystems that are of immediate interest (for appraisal of the concept viability). These are (1) the shock generation in a high pressure gas using either a plasma thruster or pulsed high magnetic field, (2) collision of MHD or gas dynamic shocks, their interaction time, and collision pile-up region thickness, and (3) magnetic flux compression power generation (not included here).

Kazeminezhad, F.↗
NASA NTRS1987

Numerical simulation of shock wave diffraction by TVD schemes

An upwind total variation diminishing (TVD) scheme and a predictor-corrector symmetric TVD scheme were used to numerically simulate the blast wave diffraction on a stationary object. The objective is to help design an optimum configuration so that lateral motion is minimized and at the same time vortex shedding and flow separation are reduced during a blast wave encounter. Results are presented for a generic configuration for both a coarse grid and a fine grid to illustrate the global and local diffraction flow fields. Numerical experiments for the shock wave reflection on a wedge are also included to validate the current approach. Numerical study indicated that these TVD schemes are more stable and produced higher shock resolution than classical shock capturing methods such as the explicit MacCormack scheme.

Young, Victor Y. C.↗
NASA NTRS2018

Conceptual Design of an Experiment to Study Dust Destruction by Astrophysical Shock Waves

A novel laboratory experimental design is described that will investigate the processing of dust grains in astrophysical shocks. Dust is a ubiquitous ingredient in the interstellar medium (ISM) of galaxies; however, its evolutionary cycle is still poorly understood. Especially shrouded in mystery is the efficiency of grain destruction by astrophysical shocks generated by expanding supernova remnants. While the evolution of these remnants is fairly well understood, the grain destruction efficiency in these shocks is largely unknown. The experiments described herein will fill this knowledge gap by studying the dust destruction efficiencies for shock velocities in the range of approximately 10-30 kilometers per second (microns per nanosecond), at which most of the grain destruction and processing in the ISM takes place. The experiments focus on the study of grain-grain collisions by accelerating small (1 millimeter) dust particles into a large (approximately 5-10 millimeter diameter) population; this simulates the astrophysical system well in that the more numerous, small grains impact and collide with the large population. Facilities that combine the versatility of high-power optical lasers with the diagnostic capabilities of X-ray free-electron lasers, e.g., the Matter in Extreme Conditions instrument at the SLAC (originally named Stanford Linear Accelerator Center) National Accelerator Laboratory, provide an ideal laboratory environment to create and diagnose dust destruction by astrophysically relevant shocks at the micron scale.

Manuel, M. J.-E.↗
NASA NTRS1981

Experimental studies of scale effects on oscillating airfoils at transonic speeds

Scale effects are discussed with reference to a conventional airfoil (NACA 64A010) and a supercritical airfoil (NLR 7301) at mean flow conditions that support both weak and strong shock waves. During the experiment the Reynolds number was varied from 3 x 10 to the sixth power at time history data are presented over the range of reduced frequencies that are important in aeroelastic applications. The experimental data show that viscous effects are important in the case of the supercritical airfoil at all flow conditions and in the case of the conventional airfoil under strong shock wave conditions. Some frequency dependent viscous effects were also observed.

Davis, S. S.↗
NASA NTRS1980

Experimental studies of scale effects on oscillating airfoils at transonic speeds

Experimental data are presented on the effect of Reynolds number on unsteady pressures induced by the pitching motion of an oscillating airfoil. Scale effects are discussed with reference to a conventional airfoil (NACA 64A010) and a supercritical airfoil (NLR 7301) at mean-flow conditions that support both weak and strong shock waves. During the experiment the Reynolds number was varied from 3,000,000 to 12,000,000 at a Mach number and incidence necessary to induce the required flow. Both fundamental frequency and complete time history data are presented over the range of reduced frequencies that is important in aeroelastic applications. The experimental data show that viscous effects are important in the case of the supercritical airfoil at all flow conditions and in the case of the conventional airfoil under strong shock-wave conditions. Some frequency-dependent viscous effects were also observed.

Davis, S. S.↗
NASA NTRS2015

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

Preliminary results of an experimental investigation of a Mach 2.5 two-dimensional axisymmetric shock-wave/boundary-layer interaction (SWBLI) are presented. The purpose of the investigation is to create a SWBLI dataset specifically for CFD validation purposes. Presented herein are the details of the facility and preliminary measurements characterizing the facility and interaction region. The results will serve to define the region of interest where more detailed mean and turbulence measurements will be made.

Superonic Flow↗
NASA NTRS2015

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

Preliminary results of an experimental investigation of a Mach 2.5 two-dimensional axisymmetric shock-wave/ boundary-layer interaction (SWBLI) are presented. The purpose of the investigation is to create a SWBLI dataset specifically for CFD validation purposes. Presented herein are the details of the facility and preliminary measurements characterizing the facility and interaction region. These results will serve to define the region of interest where more detailed mean and turbulence measurements will be made.

Shock-Wave/Boundary Layer Interaction↗
NASA NTRS1985

Computation of three-dimensional shock-wave/turbulent boundary-layer interaction flows

Solutions of the Reynolds-averaged Navier-Stokes equations, employing a two-equation turbulence model, are presented and compared with measurements from a series of supersonic shock-wave/boundary-layer interaction experiments. The test flows include swept-compression corners and axisymmetric bodies with either skewed or segmented conical flares. The solutions correctly predict all major qualitative features of the flow field. However, for flow fields with large separated zones, significant quantitative differences are observed between the computed and experimental results. Possible reasons for these differences are discussed.

Horstman, C. C.↗
NASA NTRS2013

Investigation of Materials for Boundary Layer Control in a Supersonic Wind Tunnel

During operation of the NASA Glenn Research Center 15- by 15-Centimeter Supersonic Wind Tunnel (SWT), a significant, undesirable corner flow separation is created by the three-dimensional interaction of the wall and floor boundary layers in the tunnel corners following an oblique-shock/ boundary-layer interaction. A method to minimize this effect was conceived by connecting the wall and floor boundary layers with a radius of curvature in the corners. The results and observations of a trade study to determine the effectiveness of candidate materials for creating the radius of curvature in the SWT are presented. The experiments in the study focus on the formation of corner fillets of four different radii of curvature, 6.35 mm (0.25 in.), 9.525 mm (0.375 in.), 12.7 mm (0.5 in.), and 15.875 mm (0.625 in.), based on the observed boundary layer thickness of 11.43 mm (0.45 in.). Tests were performed on ten candidate materials to determine shrinkage, surface roughness, cure time, ease of application and removal, adhesion, eccentricity, formability, and repeatability. Of the ten materials, the four materials which exhibited characteristics most promising for effective use were the heavy body and regular type dental impression materials, the basic sculpting epoxy, and the polyurethane sealant. Of these, the particular material which was most effective, the heavy body dental impression material, was tested in the SWT in Mach 2 flow, and was observed to satisfy all requirements for use in creating the corner fillets in the upcoming experiments on shock-wave/boundary-layer interaction.

Braafladt, Alexander↗
NASA NTRS1998

Shock Wave Dynamics in Weakly Ionized Gases

We have begun a comprehensive series of analyses and experiments to study the basic problem of shock wave dynamics in ionized media. Our objective is to isolate the mechanisms that are responsible for the decrease in the shock amplitude and also to determine the relevant plasma parameters that will be required for a drag reduction scheme in an actual high altitude hypersonic flight. Specifically, we have initiated a program of analyses and measurements with the objective of (i) fully characterizing the propagation dynamics in plasmas formed in gases of aerodynamic interest, (ii) isolating the mechanisms responsible for the decreased shock strength and increased shock velocity, (iii) extrapolating the laboratory observations to the technology of supersonic flight.

Johnson, Joseph A., III↗
NASA NTRSDate not supplied

Implicit Large-Eddy Simulation of Axisymmetric and Asymmetric Turbulent Shock Wave Boundary Layer Interaction at Mach 2.5

High-speed inlet flows feature turbulent shock wave boundary layer interactions at supersonic Mach numbers. A canonical Mach 2.5 shock wave boundary layer interaction experiment at the NASA Glenn Research Center was designed for axisymmetric and asymmetric interactions depending on the placement of the shock generator. Four different wall-resolved implicit large-eddy simulations of the experiment were carried out at one-tenth of the Reynolds number of the experiment: (1) Two axisymmetric interactions with shock generator radii of 9.2% and 14.7% of the test section diameter and (2) two asymmetric interactions with 14.7% shock generator radius and shock generator centerline offsets of one-third and one-fourth of the test section radius. For the two axisymmetric interactions, various sub-domains upstream of the interaction, within the interaction, and downstream of the interaction were analyzed with the proper orthogonal decomposition. For all three regions, the dominant modes are streamwise coherent structures with Strouhal numbers based on separation length of 0.15 and below. Fourier spectra of the wall-pressure coefficient for the interaction region have peaks at Strouhal numbers based on separation length of 0.15 and 0.04. For the two asymmetric interactions, the streamwise extent of the separation varies considerably in the azimuthal direction. For both shock generator offsets, the interaction is strongest for the azimuthal location that is closest to the shock generator. Compared to the axisymmetric interaction, the streamwise extent of the separated flow region is noticeably reduced and a substantial crossflow is observed. Fourier analyses of the unsteady flow indicate low-frequency content f or the upstream part of the interaction which is more two-dimensional. Away from the upstream region, with increasing sweep angle and crossflow, the low-frequency content is diminished.

Computational Fluid Dynamics↗
NASA NTRS1975

An experimental documentation of a hypersonic shock-wave turbulent boundary layer interaction flow: With and without separation

An experiment, thoroughly documenting the flow field resulting from the interaction of a shock wave with a nonadiabatic hypersonic turbulent boundary layer, is described. Detailed mean flow and surface data are presented for two shock strengths resulting in attached and separated flows, respectively. The surface measurements include continuous pressure, shear and heat-flux distributions upstream, in, and downstream of the interaction regions. At closely spaced intervals along the surface, boundary-layer profiles of static and pitot pressure and total temperature were obtained from which velocity, density and static temperature profiles were derived. The data are presented in both graphical and tabular form. These data are of sufficient detail to validate advanced computer codes and their associated turbulence models.

Kussoy, M. I.↗