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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 NTRS1974

Comparison of several methods for predicting separation in a compressible turbulent boundary layer

Several methods for predicting the separation point for a compressible turbulent boundary layer were applied to the flow over a bump on a wind-tunnel wall. Measured pressure distributions were used as input. Two integral boundary-layer methods, three finite-difference boundary-layer methods, and three simple methods were applied at five free-stream Mach numbers ranging from 0.354 to 0.7325. Each of the boundary-layer methods failed to explicitly predict separation. However, by relaxing the theoretical separation criteria, several boundary-layer methods were made to yield reasonable separation predictions, but none of the methods accurately predicted the important boundary-layer parameters at separation. Only one of the simple methods consistently predicted separation with reasonable accuracy in a manner consistent with the theory. The other methods either indicated several possible separation locations or only sometimes predicted separation.

Gerhart, P. M.↗
NASA NTRS1985

Mean-flow measurements in a supersonic three-dimensional turbulent boundary layer

Measurements are presented of the mean flow in a supersonic turbulent boundary layer subjected to a constant weak transverse pressure gradient. The temperature and longitudinal velocity and the axial shear stress were only slightly affected by the three-dimensionality but a clearly defined crossflow component appeared, in the manner suggested by theory and confirmed in earlier experiments. The flow deflection angle and transverse-velocity component achieved a self-preserving form so long as the transverse-pressure gradient remained constant, and both achieved a maximum at the sublayer edge; the deflection angle seemed to decrease again between the latter and the surface. An empirical relation was found between the pressure-gradient strength and the maximum in the crossflow, and the dependence of the latter on distance from the surface was used to test analytical crossflow predictions. The data are in general agreement with Van Den Berg's law of the wall. The data also support the so-called parabolic law following a relaxation distance, especially if inviscid gradients are accounted for, and if the normal coordinate is contracted by the compressibility transformation.

Demetriades, A.↗
NASA NTRS1991

Acoustic sources in the low Mach number turbulent boundary layer

The sources of sound production in a low Mach number turbulent boundary layer are examined. The sources are shown to be quadrupole in nature and to result from supersonically convecting wave-number components of the fluctuating Reynolds' normal stresses. The primary Tollmien-Schlichting instability of the boundary layer is found to radiate no sound. Analysis of various vortical phenomena suggests that the primary source is the process of formation of horseshoe vortices, with viscous sublayer bursts a possible secondary source.

Hardin, Jay C.↗
NASA NTRS1973

Flowfield analysis for successive oblique shock wave interactions with a turbulent boundary layer

A computation procedure is described for predicting flow fields which develop when successive interactions between oblique shock waves and a turbulent boundary layer occur. Computations were carried out for axisymmetric internal flows at free stream Mach numbers = 3.82 and 2.82. The effect of boundary layer bleed was considered for the free stream Mach number = 2.82 flow. A control volume analysis is used to predict changes in the flow field across the interactions. Two bleed flow models were considered. A turbulent boundary layer program was used to compute changes in the boundary layer between the interactions. The results given are for flows with two shock wave interactions and for bleed at the second interaction site. In principle the method described may be extended to account for additional interactions. The predicted results are compared with measured results and are shown to be in good agreement when the bleed flow rate is low (on the order of 3 percent of the boundary layer mass flow), or when there is no bleed. As the bleed flow rate is increased, differences between the predicted and measured results become larger. Shortcomings of the bleed flow models at higher bleed flow rates are discussed.

Sun, C. C.↗
NASA NTRS1991

Exploratory study of in-plane streamline curvature effects on a turbulent boundary layer at a Mach number of 3

This report on a program to study in-plane streamline curvature effects in a turbulent boundary layer at a Mach number of 3. The original proposal, for a 3-year program to explore in-plane streamline curvature effects on a supersonic turbulent boundary layer using three-dimensional pressure fields generated by fins and wall geometry, ended after one year. The purpose of these tests was to compare these streamline curvature effects to the more classical two-dimensional curvature generated by wall shape and imposed pressure gradients, previously considered primarily in a plane normal to the floor. The studies were carried out in the Mach number of 3, 8 x 8 inch High Reynolds Number Supersonic Tunnel. The usual surface visualization and mean wall static pressures were supplemented by the use of many small high frequency wall static pressure gauges (Kulites) to get some indication of the amplification of boundary layer disturbances by the in-plane streamline curvature caused by the three-dimensional pressure fields imposed on the boundary layer.

Bogdonoff, Seymour M.↗
NASA NTRS2006

Optimal Control of Shock Wave Turbulent Boundary Layer Interactions Using Micro-Array Actuation

The intent of this study on micro-array flow control is to demonstrate the viability and economy of Response Surface Methodology (RSM) to determine optimal designs of micro-array actuation for controlling the shock wave turbulent boundary layer interactions within supersonic inlets and compare these concepts to conventional bleed performance. The term micro-array refers to micro-actuator arrays which have heights of 25 to 40 percent of the undisturbed supersonic boundary layer thickness. This study covers optimal control of shock wave turbulent boundary layer interactions using standard micro-vane, tapered micro-vane, and standard micro-ramp arrays at a free stream Mach number of 2.0. The effectiveness of the three micro-array devices was tested using a shock pressure rise induced by the 10 shock generator, which was sufficiently strong as to separate the turbulent supersonic boundary layer. The overall design purpose of the micro-arrays was to alter the properties of the supersonic boundary layer by introducing a cascade of counter-rotating micro-vortices in the near wall region. In this manner, the impact of the shock wave boundary layer (SWBL) interaction on the main flow field was minimized without boundary bleed.

Anderson, Bernhard H.↗
NASA NTRS1978

Effect of wall temperature on a supersonic turbulent boundary layer

Measurement of mean flow profiles in a fully developed Mach 3 turbulent boundary layer with negligible pressure gradient is reported. Data were acquired at several streamwise locations for wall-to-total temperature ratios of 0.94, 0.71, and 0.54. The results demonstrate that the velocity defect formulation of the law-of-the-wake, which successfully correlates compressible, adiabatic boundary layers, is also valid for nonadiabatic flows. It is also shown that for adiabatic walls, the linear Crocco relation between total temperature and velocity does not provide a valid test of the nature of the boundary-layer flow for practical cases where the Prandtl number departs from unity. Finally, the turbulent shear stress, mixing length, and eddy viscosity were extracted from the 'time-averaged' conservation equations using the measured mean flow profiles and found to be insensitive to wall temperature. In particular, the latter properties are in good agreement with earlier compressible, adiabatic correlations of turbulent transport properties.

Laderman, A. J.↗
NASA NTRS1976

Flowfield analysis for successive oblique shock wave-turbulent boundary layer interactions

A computation procedure is described for predicting the flowfields which develop when successive interactions between oblique shock waves and a turbulent boundary layer occur. Such interactions may occur, for example, in engine inlets for supersonic aircraft. Computations are carried out for axisymmetric internal flows at M 3.82 and 2.82. The effect of boundary layer bleed is considered for the M 2.82 flow. A control volume analysis is used to predict changes in the flow field across the interactions. Two bleed flow models have been considered. A turbulent boundary layer program is used to compute changes in the boundary layer between the interactions. The results given are for flows with two shock wave interactions and for bleed at the second interaction site. In principle the method described may be extended to account for additional interactions. The predicted results are compared with measured results and are shown to be in good agreement when the bleed flow rate is low (on the order of 3% of the boundary layer mass flow), or when there is no bleed. As the bleed flow rate is increased, differences between the predicted and measured results become larger. Shortcomings of the bleed flow models at higher bleed flow rates are discussed.

Sun, C. C.↗
NASA NTRS1972

Pitot-probe displacement in a supersonic turbulent boundary layer

Eight circular pitot probes ranging in size from 2 to 70 percent of the boundary-layer thickness were tested to provide experimental probe displacement results in a two-dimensional turbulent boundary layer at a nominal free-stream Mach number of 2 and unit Reynolds number of 8 million per meter. The displacement obtained in the study was larger than that reported by previous investigators in either an incompressible turbulent boundary layer or a supersonic laminar boundary layer. The large probes indicated distorted Mach number profiles, probably due to separation. When the probes were small enough to cause no appreciable distortion, the displacement was constant over most of the boundary layer. The displacement in the near-wall region decreased to negative displacement in some cases. This near-wall region was found to extend to about one probe diameter from the test surface.

Allen, J. M.↗
NASA NTRS1975

Experimental results for the transpired turbulent boundary layer in an adverse pressure gradient

The fluid mechanics of transpired incompressible turbulent boundary layers under zero and adverse pressure gradient conditions is investigated using an open-ended wind tunnel with a porous floor in the test section and a secondary air system for supply and metering of the transpiration air. All velocity profiles and turbulence profiles are obtained by linearized constant-temperature hot-wire anemometry. The wall shear stress is determined by measuring the shear stress away from the wall and extrapolating to the wall by integrating the boundary layer equations for the shear-stress profile. Equilibrium boundary layers are obtained when the transpiration velocity is varied such that the blowing parameter and the Clauser pressure gradient parameter are held constant. The experimental results obtained are presented in tabular and graphical forms.

Andersen, P. S.↗
NASA NTRS2001

A High-Lift Building Block Flow: Turbulent Boundary Layer Relaminarization

A working wind tunnel test facility has been constructed at the University of Notre Dame's Hessert Center. The relaminarization test facility has been constructed in the 1.5m x 1.5m (5ft x 5 ft) atmospheric wind tunnel and generates a Re(theta)=4694 turbulent boundary layer in nominally zero-pressure gradient before it is exposed to the Case #1 pressure gradient (K approximately equal to 4.2 x 10(exp -6), which is believed to be sufficient to achieve relaminarization. Future work to be conducted will include measuring the response of the turbulent boundary layer to the favorable pressure gradients created in the test facility and documenting this response in order to understand the underlying flow physics responsible for relaminarization. It is the goal of this research to have a better understanding of accelerated turbulent boundary layers which will aid in the development of future flow diagnostic utilities to be implemented in applied aerodynamic research.

Bourassa, Corey↗
NASA NTRS1985

An experimental and computational study of turbulent boundary layers in oscillating flows

An experimental-computational study of the behavior of turbulent boundary layers for oscillating air flows over a plane surface with a small favorable mean pressure gradient is described. Experimental data were obtained for boundary layers generated on the test section wall of a new type of facility that produces a mean flow velocity with a superposed sinusoidal component over a wide frequency range. Experimental results for unsteady velocity profiles and turbulence intensity obtained from hot wire measurements at a fixed stream-wise position on the surface are presented for frequencies ranging from 5 to 21 Hz. Numerical predictions were obtained using an unsteady boundary layer computational code and the Cebeci-Smith and Glushko turbulence models. Predicted quantities related to unsteady velocity profiles exhibit fair agreement with experiment when the Cebeci-Smith model is used.

Cook, W. J.↗