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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Computation of three-dimensional, inviscid supersonic flows

The paper sets forth in detail a method for the finite-difference computation of three-dimensional supersonic fields in an Eulerian mesh. First-, second-, and third-order finite difference schemes are examined. Attention is given to proper treatment of the impermeable and permeable boundaries encompassing the computational plane. Numerical results are presented for certain specific configurations: a conical wing-body combination, internal corner flow, a two-dimensional blunt body, an interfering shock problem, and three-dimensional inviscid supersonic flow past a shuttle-orbiter type vehicle.

Kutler, P.↗
NASA NTRS1979

A higher order panel method for linearized supersonic flow

The basic integral equations of linearized supersonic theory for an advanced supersonic panel method are derived. Methods using only linear varying source strength over each panel or only quadratic doublet strength over each panel gave good agreement with analytic solutions over cones and zero thickness cambered wings. For three dimensional bodies and wings of general shape, combined source and doublet panels with interior boundary conditions to eliminate the internal perturbations lead to a stable method providing good agreement experiment. A panel system with all edges contiguous resulted from dividing the basic four point non-planar panel into eight triangular subpanels, and the doublet strength was made continuous at all edges by a quadratic distribution over each subpanel. Superinclined panels were developed and tested on s simple nacelle and on an airplane model having engine inlets, with excellent results.

Ehlers, F. E.↗
NASA NTRS2008

Boundary Layer Receptivity due to Roughness and Freestream Sound for Supersonic Flows over Axisymmetric Cones

The receptivity of supersonic boundary layers due to the interaction of a two-dimensional acoustic wave with a three-dimensional roughness is numerically investigated over a 5-degree sharp tipped straight cone at a free stream Mach number of 3.5 and a high Reynolds number of 10(exp 6)/inch. Both the steady and unsteady solutions are obtained by solving the full Navier-Stokes equations using the fifth-order accurate weighted essentially non-oscillatory (WENO) scheme for spatial discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for temporal integration. The simulations are performed with different roughness heights and roughness locations. The steady computations with the roughness reveal that the perturbations induced by the roughness remain very close to the wall and do not penetrate to the outer part of the boundary layer. The effects of the roughness persist for about 10 boundary layer thicknesses downstream of each element. The unsteady simulation shows the generation of unstable three-dimensional first mode disturbances within a short distance from the roughness. However, the amplitudes of the instability waves are very small near the neutral point, on the order of 10(exp -4) compared to the free stream disturbances. The receptivity coefficients due to the roughness elements located very close to the tip are one order of magnitude smaller than those due to the roughness elements located closer to the neutral point.

Balakumar, P.↗
NASA NTRS1992

The flip-flop nozzle extended to supersonic flows

An experiment studying a fluidically oscillated rectangular jet flow was conducted. The Mach number was varied over a range from low subsonic to supersonic. Unsteady velocity and pressure measurements were made using hot wires and piezoresistive pressure transducers. In addition smoke flow visualization using high speed photography was used to document the oscillation of the jet. For the subsonic flip-flop jet it was found that the apparent time-mean widening of the jet was not accompanied by an increase in mass flux. It was found that it is possible to extend the operation of these devices to supersonic flows. Most of the measurements were made for a fixed nozzle geometry for which the oscillations ceased at a fully expanded Mach number of 1.58. By varying the nozzle geometry this limitation was overcome and operation was extended to Mach 1.8. The streamwise velocity perturbation levels produced by this device were much higher than the perturbation levels that could be produced using conventional excitation sources such as acoustic drivers. In view of this ability to produce high amplitudes, the potential for using small scale fluidically oscillated jet as an unsteady excitation source for the control of shear flows in full scale practical applications seems promising.

Raman, Ganesh↗
NASA NTRS1992

The flip flop nozzle extended to supersonic flows

An experiment studying a fluidically oscillated rectangular jet flow was conducted. The Mach number was varied over a range from low subsonic to supersonic. Unsteady velocity and pressure measurements were made using hot wires and piezoresistive pressure transducers. In addition smoke flow visualization using high speed photography was used to document the oscillation of the jet. For the subsonic flip-flop jet it was found that the apparent time-mean widening of the jet was not accompanied by an increase in mass flux. It was found that it is possible to extend the operation of these devices to supersonic flows. Most of the measurements were made for a fixed nozzle geometry for which the oscillations ceased at a fully expanded Mach number of 1.58. By varying the nozzle geometry this limitation was overcome and operation was extended to Mach 1.8. The streamwise velocity perturbation levels produced by this device were much higher than the perturbation levels that could be produced using conventional excitation sources such as acoustic drivers. In view of this ability to produce high amplitudes, the potential for using small scale fluidically oscillated jet as an unsteady excitation source for the control of shear flows in full scale practical applications seems promising.

Raman, Ganesh↗
NASA NTRS1972

A stagnation pressure probe for use in supersonic flow

Probe was developed which is capable of recording true stagnation pressure in supersonic and subsonic flows. Static pressure orifices in the compression surface allow probe to be used as combined flow direction and stagnation pressure sensor.

Goodyer, M. J.↗
NASA NTRS1976

Quasi-three-dimensional laminar boundary-layer separations in supersonic flow

A numerical algorithm is presented for solving laminar, steady, supersonic interacting boundary-layer flows for quasi-three-dimensional configurations. The interaction problem is treated as a boundary-value problem and a salient feature of the scheme is the direct implementation of the downstream boundary condition. Solutions are presented for axisymmetric and swept (yawed) compression ramps for both adiabatic and heat transfer conditions over a Mach number range of 2-6. The results are in good agreement with experimental data and existing theories for axisymmetric cases. For the swept (yawed) configurations, lack of experimental data makes a direct comparison impossible, but the present solutions are found to be in qualitative agreement with earlier studies. In addition it is shown that the trends obtained here for the sweep effects are well predicted by a simple extension of the two-dimensional asymptotic theory.

Vatsa, V. N.↗
NASA NTRS1973

Boundary condition calculation procedures for inviscid supersonic flow fields.

Results are given of a comparative study of numerical procedures for computing solid wall and sharp shock boundary points in supersonic inviscid flow calculations. Thirty different calculation procedures were tested on four sample problems in two dimensional steady flow. A new simple but accurate calculation procedure was developed for solid wall boundary point calculations. The merits and shortcomings of the various procedures are thoroughly discussed.

Abbett, M. J.↗
NASA NTRS2002

A Quantitative Comparison of Leading-edge Vortices in Incompressible and Supersonic Flows

When requiring quantitative data on delta-wing vortices for design purposes, low-speed results have often been extrapolated to configurations intended for supersonic operation. This practice stems from a lack of database owing to difficulties that plague measurement techniques in high-speed flows. In the present paper an attempt is made to examine this practice by comparing quantitative data on the nearwake properties of such vortices in incompressible and supersonic flows. The incompressible flow data are obtained in experiments conducted in a low-speed wind tunnel. Detailed flow-field properties, including vorticity and turbulence characteristics, obtained by hot-wire and pressure probe surveys are documented. These data are compared, wherever possible, with available data from a past work for a Mach 2.49 flow for the same wing geometry and angles-of-attack. The results indicate that quantitative similarities exist in the distributions of total pressure and swirl velocity. However, the streamwise velocity of the core exhibits different trends. The axial flow characteristics of the vortices in the two regimes are examined, and a candidate theory is discussed.

Wang, F. Y.↗
NASA NTRS1973

An improved method for the aerodynamic analysis of wing-body-tail configurations in subsonic and supersonic flow. Part 1: Theory and application

A new method has been developed for calculating the pressure distribution and aerodynamic characteristics of wing-body-tail combinations in subsonic and supersonic potential flow. A computer program has been developed to perform the numerical calculations. The configuration surface is subdivided into a large number of panels, each of which contains an aerodynamic singularity distribution. A constant source distribution is used on the body panels, and a vortex distribution having a linear variation in the streamwise direction is used on the wing and tail panels. The normal components of velocity induced at specified control points by each singularity distribution are calculated and make up the coefficients of a system of linear equations relating the strengths of the singularities to the magnitude of the normal velocities. The singularity strengths which satisfy the boundary condition of tangential flow at the control points for a given Mach number and angle of attack are determined by solving this system of equations using an interactive procedure. Once the singularity strengths are known, the pressure coefficients are calculated, and the forces and moments acting on the configuration determined by numerical integration.

Woodward, F. A.↗
NASA NTRS1989

The response of a laminar boundary layer in supersonic flow to small amplitude progressive waves

The effect of a small amplitude progressive wave on the laminar boundary layer on a semi-infinite flat plate, due to a uniform supersonic freestream flow, is considered. The perturbation to the flow divides into two streamwise zones. In the first, relatively close to the leading edge of the plate, on a transverse scale comparable to the boundary layer thickness, the perturbation flow is described by a form of the unsteady linearized compressible boundary layer equations. In the freestream, this component of flow is governed by the wave equation, the solution of which provides the outer velocity conditions for the boundary layer. This system is solved numerically, and also the asymptotic structure in the far downstream limit is studied. This reveals a breakdown and a subsequent second streamwise zone, where the flow disturbance is predominantly inviscid. The two zones are shown to match in a proper asymptotic sense.

Duck, Peter W.↗
NASA NTRS1958

The similarity rules for second-order subsonic and supersonic flow

The similarity rules for linearized compressible flow theory (Gothert's rule and its supersonic counterpart) are extended to second order. It is shown that any second-order subsonic flow can be related to "nearly incompressible" flow past the same body, which can be calculated by the Janzen-Rayleigh method.

Van Dyke, Milton D↗