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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
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↗
DOE OSTI2026

A Tutorial on Bayesian analysis of linear shock compression data

Gas gun and other shock compression experiments often produce shock wave velocity measurements that are linearly associated with particle velocity. Traditionally, this empirical relationship is quantified with a single Hugoniot curve that is estimated using least squares regression. However, for downstream modeling and simulation tasks, it is often more useful to have multiple Hugoniot curves in the pressure–volume plane that are consistent with the data. We employ Bayesian uncertainty quantification methods as a framework for propagating measurement uncertainty through to model parameters and predictions. Specifically, this Tutorial shows how to sample multiple Hugoniot curves in the pressure–volume plane that are consistent with the shock wave-particle velocity measurements in a two-step Bayesian approach. First, we obtain an analytical expression for the posterior distribution of the linear model parameters using Bayesian linear regression. Second, we propagate samples from the posterior distribution through the Rankine–Hugoniot equations to yield Hugoniot curves in the pressure–volume plane. The procedure is demonstrated with publicly available data on argon, copper, and nickel, and compared against bootstrapping and linear regression. The Bayesian procedure is shown to be interpretable, computationally inexpensive, and less sensitive than an alternative bootstrapping approach to the removal of the point in the copper dataset that has the largest particle velocity. As a Tutorial on Bayesian methodology for the shock compression community, we provide several derivations and explanations that make this paper self-contained, and make all code and data available at github.com/llnl/BALSCD.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
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.↗
NASA NTRS1981

Calculation of Unsteady Transonic Flow over an Airfoil

An implicit finite-difference solver for either the Euler equations or the "thin-layer" Navier-Stokes equations was used to calculate a transonic flow over the NACA 64A010 airfoil pitching about its one-quarter chord. An unsteady automatic grid-generation procedure that will improve significantly the computational efficiency of various unsteady flow problems is described. The calculated results for both inviscid and viscous flows at Much number 0.8 over the airfoil oscillating with reduced frequency referenced to one-half chord, 0.2, are compared with experimental data measured in the Ames 11 x 11-ft Transonic Wind Tunnel. Nonlinear, unsteady effects of the flow on the surface pressure variations, shock-wave excursions, and overall airloads are examined. Good agreements between the results of computations and experiments were obtained. In the shock-wave region, however, the results of the viscous-flow computations showed closer agreement with the experimental data.

Chyu, W. J.↗
NASA NTRS1981

Calculation of Unsteady Transonic Flow Over an Airfoil

An implicit finite-difference solver for either the Euler equations or the "thin-layer" Navier-Stokes equations was used to calculate a transonic flow over the NACA 64A010 airfoil pitching about its one-quarter chord. An unsteady automatic grid-generation procedure that will improve significantly the computational efficiency of various unsteady flow problems is described. The calculated results for both inviscid and viscous flows at Mach number 0.8 over the airfoil oscillating with reduced frequency referenced to one-half chord, 0.2, are compared with experimental data measured in the Ames 11 x 11 ft Transonic Wind Tunnel. Nonlinear, unsteady effects of the flow on the surface pressure variations, shock-wave excursions, and overall airloads are examined. Good agreements between the results of computations and experiments were obtained. In the shock-wave region, however, the results of the viscous-flow computations showed closer agreement with the experimental data.

Chyu, W. J.↗
NASA NTRS1975

Shock-wave-induced turbulent boundary-layer separation at hypersonic speeds

An experiment is described that tests and guides computations of the interaction of a shock wave with a turbulent boundary layer. Numerical solutions of the time-averaged Navier-Stokes equations for the entire flow field employing algebraic eddy viscosity and turbulent Prandtl number models for shear stress and heat flux are presented and used to illustrate the dependence of the computations on the particulars of the turbulence models. To guide modifications in the models, the mean flow profiles and surface measurements of pressure, shear, and heat flux are analyzed critically. The results show that the models of eddy viscosity require substantial modifications in the interaction region. Improved solutions employing the experimentally modified models are presented.

Horstman, C. C.↗
NASA NTRS1975

An experimental and numerical investigation of shock-wave induced turbulent boundary-layer separation at hypersonic speeds

An experiment designed to test and guide computations of the interaction of an impinging shock wave with a turbulent boundary layer is described. Detailed mean flow-field and surface data are presented for two shock strengths which resulted in attached and separated flows, respectively. Numerical computations, employing the complete time-averaged Navier-Stokes equations along with algebraic eddy-viscosity and turbulent Prandtl number models to describe shear stress and heat flux, are used to illustrate the dependence of the computations on the particulars of the turbulence models. Models appropriate for zero-pressure-gradient flows predicted the overall features of the flow fields, but were deficient in predicting many of the details of the interaction regions. Improvements to the turbulence model parameters were sought through a combination of detailed data analysis and computer simulations which tested the sensitivity of the solutions to model parameter changes. Computer simulations using these improvements are presented and discussed.

Marvin, J. G.↗
NASA NTRS2013

Experimental Database with Baseline CFD Solutions: 2-D and Axisymmetric Hypersonic Shock-Wave/Turbulent-Boundary-Layer Interactions

A database compilation of hypersonic shock-wave/turbulent boundary layer experiments is provided. The experiments selected for the database are either 2D or axisymmetric, and include both compression corner and impinging type SWTBL interactions. The strength of the interactions range from attached to incipient separation to fully separated flows. The experiments were chosen based on criterion to ensure quality of the datasets, to be relevant to NASA's missions and to be useful for validation and uncertainty assessment of CFD Navier-Stokes predictive methods, both now and in the future. An emphasis on datasets selected was on surface pressures and surface heating throughout the interaction, but include some wall shear stress distributions and flowfield profiles. Included, for selected cases, are example CFD grids and setup information, along with surface pressure and wall heating results from simulations using current NASA real-gas Navier-Stokes codes by which future CFD investigators can compare and evaluate physics modeling improvements and validation and uncertainty assessments of future CFD code developments. The experimental database is presented tabulated in the Appendices describing each experiment. The database is also provided in computer-readable ASCII files located on a companion DVD.

Database↗
NASA NTRS1989

Nonequilibrium molecular motion in a hypersonic shock wave

Molecular velocities have been measured inside a hypersonic, normal shock wave, where the gas experiences rapid changes in its macroscopic properties. As first hypothesized by Mott-Smith, but never directly observed, the molecular velocity distribution exhibits a qualitatively bimodal character that is derived from the distribution functions on either side of the shock. Quantitatively correct forms of the molecular velocity distribution function in highly nonequilibrium flows can be calculated, by means of the Direct Simulation Monte Carlo technique.

Pham-Van-diep, G.↗
NASA NTRS1994

Airfoil pressure measurements during oblique shcok-wave/vortex interaction in a Mach 3 stream

An experimental investigation of the interaction between streamwise discrete vortices and oblique shocks formed over a two-dimensional wedge surface was conducted in a Mach 3 blowdown wind tunnel. An instrumented two-dimensional wedge was placed downstream of a semiispan wing so that the trailing tip vortex would arbitrarily interact with the shock wave formed over the wedge surface. The experiments were designed to simulate interaction of streamwise vortices from upstream bodies with shock waves formed over aft surfaces as might be encountered in supersonic flight of aircraft and missiles. The influence of vortex strength and vortex-airfoil vertical separation distance on the interaction was examined.

Kalkhoran, I. M.↗
NASA NTRS1989

The computation of thermo-chemical nonequilibrium hypersonic flows

Several conceptual designs for vehicles that would fly in the atmosphere at hypersonic speeds have been developed recently. For the proposed flight conditions the air in the shock layer that envelops the body is at a sufficiently high temperature to cause chemical reaction, vibrational excitation, and ionization. However, these processes occur at finite rates which, when coupled with large convection speeds, cause the gas to be removed from thermo-chemical equilibrium. This non-ideal behavior affects the aerothermal loading on the vehicle and has ramifications in its design. A numerical method to solve the equations that describe these types of flows in 2-D was developed. The state of the gas is represented with seven chemical species, a separate vibrational temperature for each diatomic species, an electron translational temperature, and a mass-average translational-rotational temperature for the heavy particles. The equations for this gas model are solved numerically in a fully coupled fashion using an implicit finite volume time-marching technique. Gauss-Seidel line-relaxation is used to reduce the cost of the solution and flux-dependent differencing is employed to maintain stability. The numerical method was tested against several experiments. The calculated bow shock wave detachment on a sphere and two cones was compared to those measured in ground testing facilities. The computed peak electron number density on a sphere-cone was compared to that measured in a flight test. In each case the results from the numerical method were in excellent agreement with experiment. The technique was used to predict the aerothermal loads on an Aeroassisted Orbital Transfer Vehicle including radiative heating. These results indicate that the current physical model of high temperature air is appropriate and that the numerical algorithm is capable of treating this class of flows.

Candler, Graham↗
NASA NTRS1975

Electrostatic and electromagnetic turbulence associated with the earth's bow shock

Simultaneous measurements were made of the electric and magnetic field spectral densities in the earth's bow shock by a plasma wave experiment on the Imp 6 spacecraft. The frequency range of the plasma wave detector was 20 Hz to 200 kHz. Electric fields were measured with high-sensitivity 100-m long dipole antennas and magnetic fields were measured with single-turn loop antennas. Two components are distinguished in the electric field spectrum in the bow shock: one component has a broad peak centered in the region 200-800 Hz, while the other component increases monotonically with decreasing frequency. The magnetic field spectrum has only one component that increases monotonically with decreasing frequency and has an upper cutoff frequency near the local electron gyrofrequency. This magnetic field turbulence is judged to be caused by whistler mode waves. The monotonic component of the electric field spectrum is thought to be the electric field spectrum of these whistler mode waves.

Rodriguez, P.↗