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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
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 NTRS1990

Qualitative investigation of cryogenic fluid injection into a supersonic flow field

The behavior of liquid nitrogen injected into a supersonic nitrogen flow field was investigated using an experimental apparatus in which a Mach 2.7 2D gas nitrogen tunnel is coupled with a high-pressure cryogenic source. Observations were monitored and recorded via a video camera and a motion picture camera. It was found that the penetration of a supersonic flow field by injection of liquid nitrogen is strongly dependent on the flow Mach number, the cryogen injection pressure (Pi/Pc), the injector configuration, and the cryogen temperature. For a 2D gaseous N2, Mach 2.7 tunnel, with cryogen injection Pi/Pc approaching 2, the injected fluid penetration for the 1/8-in. injection port approached one half of the tunnel width at 90-deg injection, and one fourth of the tunnel width at 20-deg injection.

Hendricks, R. C.↗
NASA NTRS1976

An experimental investigation of supersonic flow past a wedge-cylinder configuration

An experimental investigation of supersonic flow past double-wedge configurations was conducted. Over the range of geometries tested, it was found that, while theoretical solutions both for a Type V pattern and for a Type VI pattern could be generated for a particular flow condition (as defined by the geometry and the free-stream conditions), the weaker, Type VI pattern was observed experimentally. More rigorous flow-field solutions were developed for the flow along the wing leading-edge. Solutions were developed for the three-dimensional flow in the plane of symmetry of a swept cylinder (which represented the wing leading-edge) which was mounted on a wedge (which generated the "bow" shock wave). A numerical code was developed using integral techniques to calculate the flow in the shock layer upstream of the interaction region (i.e., near the wing root). Heat transfer rates were calculated for various free stream conditions. The present investigation was undertaken to examine the effects of crossflow on the resultant flow-field and to verify the flow model used in theoretical calculations.

Barnette, D. W.↗
NASA NTRS1976

Supersonic Flow of Chemically Reacting Gas-Particle Mixtures. Volume 2: RAMP - A Computer Code for Analysis of Chemically Reacting Gas-Particle Flows

A computer program written in conjunction with the numerical solution of the flow of chemically reacting gas-particle mixtures was documented. The solution to the set of governing equations was obtained by utilizing the method of characteristics. The equations cast in characteristic form were shown to be formally the same for ideal, frozen, chemical equilibrium and chemical non-equilibrium reacting gas mixtures. The characteristic directions for the gas-particle system are found to be the conventional gas Mach lines, the gas streamlines and the particle streamlines. The basic mesh construction for the flow solution is along streamlines and normals to the streamlines for axisymmetric or two-dimensional flow. The analysis gives detailed information of the supersonic flow and provides for a continuous solution of the nozzle and exhaust plume flow fields. Boundary conditions for the flow solution are either the nozzle wall or the exhaust plume boundary.

Penny, M. M.↗
NASA NTRS1978

A Mach line panel method for computing the linearized supersonic flow over planar wings

A method is described for solving the linearized supersonic flow over planar wings using panels bounded by two families of Mach lines. Polynomial distributions of source and doublet strength lead to simple, closed form solutions for the aerodynamic influence coefficients, and a nearly triangular matrix yields rapid solutions for the singularity parameters. The source method was found to be accurate and stable both for analysis and design boundary conditions. Similar results were obtained with the doublet method for analysis boundary conditions on the portion of the wing downstream of the supersonic leading edge, but instabilities in the solution occurred for the region containing a portion of the subsonic leading edge. Research on the method was discontinued before this difficulty was resolved.

Ehlers, F. E.↗
NASA NTRS1994

Development of quiet-flow supersonic wind tunnels for laminar-turbulent transition research

This grant supported research into quiet-flow supersonic wind-tunnels, between May 1990 and December 1994. Quiet-flow nozzles operate with laminar nozzle-wall boundary layers, in order to provide low-disturbance flow for studies of laminar-turbulent transition under conditions comparable to flight. Major accomplishments include: (1) the design, fabrication, and performance-evaluation of a new kind of quiet tunnel, a quiet-flow Ludweig tube; (2) the integration of preexisting codes for nozzle design, 2D boundary-layer computation, and transition-estimation into a single user-friendly package for quiet-nozzle design; and (3) the design and preliminary evaluation of supersonic nozzles with square cross-section, as an alternative to conventional quiet-flow nozzles. After a brief summary of (1), a description of (2) is presented. Published work describing (3) is then summarized. The report concludes with a description of recent results for the Tollmien-Schlichting and Gortler instability in one of the square nozzles previously analyzed.

Schneider, Steven P.↗
NASA NTRS1978

Supersonic flow over ablated nosetips using an unsteady, implicit numerical procedure

The axisymmetric supersonic flow over passive, that is, nonablating, indented nosetips of reentry vehicles is determined using an unsteady implicit numerical algorithm which solves either the inviscid Euler equations or the 'thin-layer' Navier-Stokes equations. A nonorthogonal independent variable transformation is used to map the distorted physical region, containing multiple zones of embedded subsonic flow into a rectangular computational domain at whose boundaries the required permeable or impermeable boundary conditions are simulated. Use of the implicit algorithm results in faster convergence to the steady state because of a larger allowable time step over conventional explicit schemes. The numerical results obtained compare favorably with existing experimental data for very mildly and severely indented blunt nosetips.

Kutler, P.↗
NASA NTRS1993

Investigation of three-dimensional separation at wing/body junctions in supersonic flows using TVD MacCormack's scheme

A Total Variation Diminishing (TVD) MacCormack scheme is used to study the three-dimensional separation at wing/body junctions in supersonic flows. The new scheme is robust, resolves discontinuities well with-out any numerical oscillations present in the flow field. Moreover, this new scheme does not contain any problem dependent parameters to be adjusted and many production codes employing MacCormack algorithm can be easily updated to good effect. Numerical simulation carried out in laminar supersonic junction flows using the new scheme yields improved prediction for the separation location and the axial velocity profiles in the separated flow region.

Lakshmanan, B.↗
NASA NTRS1971

Optical probing of supersonic flows with statistical correlation

Remote sensing tool reliably measures statistical properties of supersonic turbulence. Tool neither affects nor is adversely affected by flow field. Device determines characteristics of supersonic flow with optical system and provides method and apparatus for separating translational and rotational motions of turbulent structures in supersonic flow.

Funk, B. H.↗
NASA NTRS1946

Flutter and oscillating air-force calculations for an airfoil in a two-dimensional supersonic flow

An account is given of the Possio theory of non stationary flow for small disturbances in a two-dimensional supersonic flow and of its application to determination of the aerodynamic forces on an oscillating airfoil. Further application is made to the problem of wing flutter in the degrees of freedom - torsion, bending, and aileron torsion. Numerical tables for flutter calculations are provided for numerous values of the Mach number greater than unity. Results for bending-torsion wing flutter are discussed. The static instabilities of divergence and aileron reversal are examined as is a one degree of freedom case of torsional oscillatory instability.

Garrick, I E↗
NASA NTRS1946

Flutter and oscillating air-force calculations for an airfoil in two-dimensional supersonic flow

A connected account is given of the Possio theory of non-stationary flow for small disturbances in a two-dimensional supersonic flow and of its application to the determination of the aerodynamic forces on an oscillating airfoil. Further application is made to the problem of wing flutter in the degrees of freedom - torsion, bending, and aileron rotations. Numerical tables for flutter calculations are provided for various values of the Mach number greater than unity. Results for bending-torsion wing flutter are shown in figures and are discussed. The static instabilities of divergence and aileron reversal are examined as is a one-degree-of-freedom case of torsional oscillatory instability.

Garrick, I E↗