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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1998

Flowfield Measurements in a Slot-Bled Oblique Shock Wave and Turbulent Boundary-Layer Interaction

An experimental investigation was conducted to determine the flowfield inside a bleed slot used to control an oblique shock-wave and turbulent boundary-layer interaction. The slot was oriented normal to the primary flow direction and had a width of 1.0 cm (primary flow direction), a length of 2.54 cm, and spanned 16.5 cm. The approach boundary layer upstream of the interaction was nominally 3.0 cm thick. Two operating conditions were studied: M = 1.98 with a shock generator deflection angle of 6 deg and M= 2.46 with a shock generator deflection angle of 8 deg. Measurements include surface and flowfield static pressure, Pitot pressure, and total mass-flow through the slot. The results show that despite an initially two-dimensional interaction for the zero bleed-flow case, the slot does not remove mass uniformly in the spanwise direction. Inside the slot, the flow is characterized by two separation regions which significantly reduce the effective flow area. The upper separation region acts as an aerodynamic throat resulting in supersonic flow through much of the slot.

Davis, D. O.↗
NASA NTRS1993

Development of separation due to interaction between a shock wave and a turbulent boundary layer perturbed by rarefaction waves

The interaction of a turbulent boundary layer with rarefaction and shock waves in flows past inclined steps was investigated experimentally and theoretically. The experiments were carried out in supersonic wind tunnels with test sections of 0.6 x 0.6 m and 0.2 x 0.2 m, with adiabatic conditions on the surface model. Based on detailed measurements of pressure fields, velocity, and surface friction, combined with results of optical visualization and analysis of limiting flow lines, three characteristic flow regimes are identified. These are (1) nonseparated flow at small step angles, (2) formation of a local separation zone with a free separation point at moderate step angles, and (3) formation of a large-scale separated flow with a fixed separation point at sufficiently large step angles. The quantitative gasdynamic schemes and surface pressure distributions for these flow regimes are shown for Mach 2.85.

Zheltovodov, A. A.↗
NASA NTRS1991

Hypersonic shock-wave turbulent-boundary-layer interaction flows - Experiment and computation

Solutions of the Reynolds-averaged Navier-Stokes equations are presented and compared with experimental surface data for a series of hypersonic shock-wave/turbulent-boundary-layer interaction flows. The turbulence models used include the standard k-epsilon two-equation eddy viscosity model, a two-layer modification to this model, and a third model with several modifications to account for compressibility effects. Both modified models gave significant improvements for all the test flows.

Horstman, C. C.↗
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 NTRS1981

Simple approximations for the asymptotic description of the interaction between a normal shock wave and a turbulent boundary layer at transonic speeds

The asymptotic description of the interaction between a normal shock wave and a turbulent boundary layer is reviewed. The layers necessary in a rational analysis of the interaction are discussed with emphasis on the differences from an interaction with a laminar boundary layer, the uncoupling of solutions for the distribution of pressure and skin friction at the wall, and the role of the Reynolds shear stress in these solutions. The accuracy of asymptotic solutions in flows at Reynolds numbers of technical interest is discussed. Solutions for the distribution of pressure and skin friction at the wall and the shape of the shock are considered for the case where the flow is near separation. For the pressure and skin friction, it is possible to write two simplified partial solutions, one valid at the beginning of the interaction and one valid somewhat downstream of the shock wave. A solution composed of these two parts and a linear interpolation between them appears to give good comparison with experiment; one unknown constant, independent of the parameters of the interaction, must be found from experiment. The simplified relations are presented. Comparison of numerical computations with experimental data indicates a possible value for the constant and shows quite satisfactory results.

Adamson, T. C., Jr.↗
NASA NTRS1988

Glancing shock wave-turbulent boundary layer interaction with boundary layer suction

Tests conducted to ascertain the stagnation pressure and flow angularity profiles of a turbulent boundary layer subjected to boundary layer suction (BLS) as it crosses a glancing sidewall shock wave have determined that the boundary layer does not separate upon crossing the shock wave. Without BLS, the upstream influence of the shock wave-induced wall static pressure rise was extensive, of the order of four bloundary layer thicknesses; for the same case, with suction, the extent of upstream influence was 50 percent lower. In addition, flow angularities at the wall were found to be smaller with suction than without it.

Barnhart, P. J.↗
NASA NTRS1984

Numerical solution for the interaction of shock wave with laminar boundary layer in two-dimensional flow on a flat plate

The finite difference computation method was investigated for solving problems of interaction between a shock wave and a laminar boundary layer, through solution of the complete Navier-Stokes equations. This method provided excellent solutions, was simple to perform and needed a relatively short solution time. A large number of runs for various flow conditions could be carried out from which the interaction characteristics and principal factors that influence interaction could be studied.

Landau, U.↗
NASA NTRS1981

Simulation of ideal-gas flow by nitrogen and other selected gases at cryogenic temperatures

The real gas behavior of nitrogen, the gas normally used in transonic cryogenic tunnels, is reported for the following flow processes: isentropic expansion, normal shocks, boundary layers, and interactions between shock waves and boundary layers. The only difference in predicted pressure ratio between nitrogen and an ideal gas which may limit the minimum operating temperature of transonic cryogenic wind tunnels occur at total pressures approaching 9 atm and total temperatures 10 K below the corresponding saturation temperature. These pressure differences approach 1 percent for both isentropic expansions and normal shocks. Alternative cryogenic test gases were also analyzed. Differences between air and an ideal diatomic gas are similar in magnitude to those for nitrogen and should present no difficulty. However, differences for helium and hydrogen are over an order of magnitude greater than those for nitrogen or air. It is concluded that helium and cryogenic hydrogen would not approximate the compressible flow of an ideal diatomic gas.

Hall, R. M.↗
NASA NTRS1959

The Interaction of an Oblique Shock Wave with a Laminar Boundary Layer

The results of some experimental and theoretical studies of the interaction of oblique shock waves with laminar boundary layers are presented. Detailed measurements of pressure distribution, shear distribution, and velocity profiles were made during the interaction of oblique shock waves with laminar boundary layers on a flat plate. From these measurements a model was derived to predict the pressure levels characteristic of separation and the length of the separated region.

Hakkinen, R. J.↗
NASA NTRS1986

On the inception lengths of swept shock-wave/turbulent boundary-layer interactions

Experimental data are shown for the inception lengths of swept compression corner-generated and fin-generated shock/boundary layer interactions at Mach 2.95. These results are found to correlate on the basis of three different flow regimes. The inception lengths of these flows are dominated by a singularity at the cylindrical/conical boundary for swept corners and by an elongation due to shock wave sweepback for fin interactions. Similarity rules for both Re(delta)and shock generator geometry effects on inception lengths are demonstrated.

Settles, G. S.↗
NASA NTRS1979

Effect of Mach number and Reynolds number on a normal shock-wave/turbulent boundary-layer interaction

Experimental values of wall pressure and skin friction have been obtained for Mach numbers from 1.32 to 1.48 and for Reynolds numbers from 8.5 x 10 to the 6th to 225 x 10 to the 6th. Increasing the Mach number reduces the wall shear and promotes incipient separation. Reynolds number variations have little effect except at the lowest value where the flow abruptly approaches separation. Comparisons are made with solutions to the time-dependent, Reynolds-averaged, Navier-Stokes equations incorporating a two-equation, Wilcox-Rubesin turbulence model. The computations are in agreement with the experimental results. Additional numerical results indicated that the wind-tunnel walls constrained the flow and suppressed the formation of a separation bubble at the shock wave.

Mateer, G. G.↗
NASA NTRS1976

A normal shock-wave turbulent boundary-layer interaction at transonic speeds

Experimental results, including surveys of the mean and fluctuating flow, and measurements of surface pressure, skin friction, and separation length, are compared with solutions to the Navier-Stokes equations utilizing various algebraic eddy viscosity models to describe the Reynolds shear stresses. The experimental data, obtained at a free-stream Mach number of 1.5 and Reynolds numbers between 10 million and 80 million, show that a separated zone forms near the foot of the shock and that its length is proportional to the initial boundary-layer thickness; that a supersonic region forms downstream of the shock; and that the shear stress increases significantly through the interaction and subsequently decays downstream. The computations adequately represent the qualitative features of the flow field throughout the interaction but quantitatively underpredict the extent of separation and the downstream level of skin friction.

Mateer, G. G.↗
NASA NTRS1989

Numerical computation of shock wave-turbulent boundary layer interaction in transonic flow over an axisymmetric curved hill

A control-volume based finite difference computation of a turbulent transonic flow over an axisymmetric curved hill is presented. The numerical method is based on the SIMPLE algorithm, and hence the conservation of mass equation is replaced by a pressure correction equation for compressible flows. The turbulence is described by a k-epsilon turbulence model supplemented by a near-wall turbulence model. In the method, the dissipation rate in the region very close to the wall is obtained from an algebraic equation and that for the rest of the flow domain is obtained by solving a partial differential equation for the dissipation rate. The other flow equations are integrated up to the wall. It is shown that the present turbulence model yields the correct location of the compression shock. The other computational results are also in good agreement with experimental data.

Kim, S.-W.↗