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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Hybrid LES/RANS Simulation of Transverse Sonic Injection into a Mach 2 Flow

A computational study of transverse sonic injection of air and helium into a Mach 1.98 cross-flow is presented. A hybrid large-eddy simulation / Reynolds-averaged Navier-Stokes (LES/RANS) turbulence model is used, with the two-equation Menter baseline (Menter-BSL) closure for the RANS part of the flow and a Smagorinsky-type model for the LES part of the flow. A time-dependent blending function, dependent on modeled turbulence variables, is used to shift the closure from RANS to LES. Turbulent structures are initiated and sustained through the use of a recycling / rescaling technique. Two higher-order discretizations, the Piecewise Parabolic Method (PPM) of Colella and Woodward, and the SONIC-A ENO scheme of Suresh and Huyhn are used in the study. The results using the hybrid model show reasonably good agreement with time-averaged Mie scattering data and with experimental surface pressure distributions, even though the penetration of the jet into the cross-flow is slightly over-predicted. The LES/RANS results are used to examine the validity of commonly-used assumptions of constant Schmidt and Prandtl numbers in the intense mixing zone downstream of the injection location.

Boles, John A.↗
NASA NTRS1978

Inviscid, nonadiabatic flow fields over blunt, sonic corner bodies for outer planet entry conditions by a method of integral relations

An investigation has been made into the ability of a method of integral relations to calculate inviscid zero degree angle of attack, radiative heating distributions over blunt, sonic corner bodies for some representative outer planet entry conditions is investigated. Comparisons have been made with a more detailed numerical method, a time asymptotic technique, using the same equilibrium chemistry and radiation transport subroutines. An effort to produce a second order approximation (two-strip) method of integral relations code to aid in this investigation is also described and a modified two-strip routine is presented. Results indicate that the one-strip method of integral relations cannot be used to obtain accurate estimates of the radiative heating distribution because of its inability to resolve thermal gradients near the wall. The two-strip method can sometimes be used to improve these estimates; however, the two-strip method has only a small range of conditions over which it will yield significant improvement over the one-strip method.

Gnoffo, P. A.↗
NASA NTRS1977

Detailed design of a quiet high flow fan

A single stage fan was designed to demonstrate the noise abatement properties of near-sonic inlet flow and long-chord stator vanes for the reduction of both upstream and downstream propagated fan source noise. It is designed to produce a pressure ratio of 1.653:1 with an adiabatic efficiency of 83.9%. The fan has a 508 mm inlet diameter with a hub/tip ratio of 0.426 and a design tip speed of 533.4 m/sec. The design inlet specific flow rate is 219.71 kg/sec sq m and there are 10 tandem stator vanes with a combined aspect ratio of 0.54.

Soltau, J. D.↗
NASA NTRS1988

Laser transit anemometer experiences in supersonic flow

The purpose of this paper is to present examples of velocity measurements obtained in supersonic flow fields with the laser transit anemometer system. Velocity measurements of a supersonic jet exhausting in a transonic flow field, a cone boundary survey in a Mach 4 flow field, and a determination of the periodic disturbance frequencies of a sonic nozzle flow field are presented. Each of the above three cases also serves to illustrate different modes of laser transit anemometer operation. A brief description of the laser transit anemometer system is also presented.

Hunter, William W., Jr.↗
NASA NTRS1975

Effect of rod gap spacing on a suction panel for laminar flow and noise control in supersonic wind tunnels

Results are presented of a coordinated experimental and theoretical study of a sound shield concept which aims to provide a means of noise reduction in the test section of supersonic wind tunnels at high Reynolds numbers. The model used consists of a planar array of circular rods aligned with the flow, with adjustable gaps between them for boundary layer removal by suction, i.e., laminar flow control. One of the basic requirements of the present sound shield concept is to achieve sonic cross flow through the gaps in order to prevent lee-side flow disturbances from penetrating back into the shielded region. Tests were conducted at Mach 6 over a local unit Reynolds number range from about 1.2 x 10 to the 6th power to 13.5 x 10 to the 6th power per foot. Measurements of heat transfer, static pressure, and sound levels were made to establish the transition characteristics of the boundary layer on the rod array and the sound shielding effectiveness.

Harvey, W. D.↗
NASA NTRS1997

Numerical Study of a Continuum Sonic Jet Interacting with a Rarefield Flow

The results of a numerical study with flow and boundary conditions based on an experiment of a continuum sonic jet interacting with rarefied flow about a sharp leading edge flat plate at zero incidence are presented. Comparisons are made between computational fluid dynamics (CFD) and direct simulation Monte Carlo (DSMC) solutions which provide an assessment of applying each technique to the flow conditions of the experiment. An analysis of the CFD results revealed a correlation between the interaction interface of the jet continuum breakdown surface and a non-dimensional parameter derived from jet and free stream flow conditions. Using the breakdown surface from the correlation, the continuum jet was uncoupled from the interaction, thus allowing an uncoupled CFD-DSMC solution to be obtained. Also, a nearest neighbor collision algorithm, similar to the subcell technique, was implemented in the DSMC solution technique. The comparison between CFD and DSMC results shows good qualitative agreement in the interaction region and good quantitative agreement elsewhere.

Glass, Christoper E.↗
NASA NTRS1997

A Numerical Simulation of a Normal Sonic Jet into a Hypersonic Cross-Flow

This study involves numerical modeling of a normal sonic jet injection into a hypersonic cross-flow. The numerical code used for simulation is GASP (General Aerodynamic Simulation Program.) First the numerical predictions are compared with well established solutions for compressible laminar flow. Then comparisons are made with non-injection test case measurements of surface pressure distributions. Good agreement with the measurements is observed. Currently comparisons are underway with the injection case. All the experimental data were generated at the Southampton University Light Piston Isentropic Compression Tube.

Jeffries, Damon K.↗
NASA NTRS1960

The Blunt Plate In Hypersonic Flow

The sonic-wedge characteristics method has been used to obtain the shock shapes and surface pressure distributions on several blunt two-dimensional shapes in a hypersonic stream for several values of the ratio of specific heats. These shapes include the blunt slab at angle of attack and power profiles of the form yb = a)P, where 0 les than m less than 1, Yb and x are coordinates of the body surface, and a is a constant. These numerical results have been compared with the results of blast-wave theory, and methods of predicting the pressure distributions and shock shapes are proposed in each case. The effects of a free-stream conical-flow gradient on the pressure distribution on a blunt slab in hypersonic flow were investigated by the sonic-wedge characteristics method and were found to be sizable in many cases. Procedures which are satisfactory for reducing pressure data obtained in conical flows with small gradients are presented.

Baradell, Donald L.↗
NASA NTRS1991

A powered lift experiment for CFD validation

The transitional flight characteristics of a geometrically simplified STOVL aircraft configuration were measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment is the first in a sequence of tests designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important flow parameters for powered lift. The model consists of a 60-deg delta wing planform with two circular high-pressure air jets located in a blended fuselage. The measured flows have a maximum freestream Mach number of 0.2. The flow is sonic and at ambient temperature in both jets. The data presented include forces and moments measured using an internal balance, pressures measured at the 281 surface pressure ports, and jet pressures and temperatures. Measurements of the flow are also made in the tunnel test section upstream and downstream of the model and at the jet exists to provide boundary conditions for the planned computations.

Roth, Karlin R.↗
NASA NTRS1976

Sound propagation in choked ducts

The linearized equations describing the propagation of sound in variable area ducts containing flow are shown to be singular when the duct mean flow is sonic. The singularity is removed when previously ignored nonlinear terms are retained. The results of a numerical study, for the case of plane waves propagating in a one-dimensional converging-diverging duct, show that the sound field is adequately described by the linearized equations only when the axial mean flow Mach number at the duct throat M sub th 0.6. For M sub th 0.6, the numerical results showed that acoustic energy flux was not conserved. An attempt was made to extend the study to include the nonlinear behavior of the sound field. Meaningful results were not obtained due, primarily, to numerical difficulties.

Hersh, A. S.↗
NASA NTRS1976

Sonic environment of aircraft structure immersed in a supersonic jet flow stream

Test methods for determining the sonic environment of aircraft structure that is immersed in the flow stream of a high velocity jet or that is subjected to the noise field surrounding the jet, were investigated. Sonic environment test data measured on a SCAT 15-F model in the flow field of Mach 1.5 and 2.5 jets were processed. Narrow band, lateral cross correlation and noise contour plots are presented. Data acquisition and reduction methods are depicted. A computer program for scaling the model data is given that accounts for model size, jet velocity, transducer size, and jet density. Comparisons of scaled model data and full size aircraft data are made for the L-1011, S-3A, and a V/STOL lower surface blowing concept. Sonic environment predictions are made for an engine-over-the-wing SST configuration.

Guinn, W. A.↗