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At least 1,693 records · Page 94
All sourcesNASA NTRSDOE OSTI
NASA NTRS1998

Computational Study of Surface Tension and Wall Adhesion Effects on an Oil Film Flow Underneath an Air Boundary Layer

The fringe-imaging skin friction (FISF) technique, which was originally developed by D. J. Monson and G. G. Mateer at Ames Research Center and recently extended to 3-D flows, is the most accurate skin friction measurement technique currently available. The principle of this technique is that the skin friction at a point on an aerodynamic surface can be determined by measuring the time-rate-of-change of the thickness of an oil drop placed on the surface under the influence of the external air boundary layer. Lubrication theory is used to relate the oil-patch thickness variation to shear stress. The uncertainty of FISF measurements is estimated to be as low as 4 percent, yet little is known about the effects of surface tension and wall adhesion forces on the measured results. A modified version of the free-surface Navier-Stokes solver RIPPLE, developed at Los Alamos National Laboratories, was used to compute the time development of an oil drop on a surface under a simulated air boundary layer. RIPPLE uses the volume of fluid method to track the surface and the continuum surface force approach to model surface tension and wall adhesion effects. The development of an oil drop, over a time period of approximately 4 seconds, was studied. Under the influence of shear imposed by an air boundary layer, the computed profile of the drop rapidly changes from its initial circular-arc shape to a wedge-like shape. Comparison of the time-varying oil-thickness distributions computed using RIPPLE and also computed using a greatly simplified numerical model of an oil drop equation which does not include surface tension and wall adhesion effects) was used to evaluate the effects of surface tension on FISF measurement results. The effects of surface tension were found to be small but not necessarily negligible in some cases.

Celic, Alan↗
NASA NTRS1987

A three-dimensional dual potential procedure with applications to wind tunnel inlets and interacting boundary layers

A dual potential decomposition of the velocity field into a scalar and a vector potential function is extended to three dimensions and used in the finite-difference simulation of steady three-dimensional inviscid rotational flows and viscous flow. The finite-difference procedure was used to simulate the flow through the 80 by 120 ft wind tunnel at NASA Ames Research Center. Rotational flow produced by the stagnation pressure drop across vanes and screens which are located at the entrance of the inlet is modeled using actuator disk theory. Results are presented for two different inlet vane and screen configurations. The numerical predictions are in good agreement with experimental data. The dual potential procedure was also applied to calculate the viscous flow along two and three dimensional troughs. Viscous effects are simulated by injecting vorticity which is computed from a boundary layer algorithm. For attached flow over a three dimensional trough, the present calculations are in good agreement with other numerical predictions. For separated flow, it is shown from a two dimensional analysis that the boundary layer approximation provides an accurate measure of the vorticity in regions close to the wall; whereas further away from the wall, caution has to be exercised in using the boundary-layer equations to supply vorticity to the dual potential formulation.

Rao, K. V.↗
NASA NTRSDate not supplied

Modelling Sonic Boom Propagation through Planetary Boundary Layer Turbulence near the Lateral Extent of the Carpet

Nonlinear acoustic propagation of sonic booms in the atmospheric boundary layer is considered in the context of a one-way solution to a third order wave equation. A split-step approach is utilized to compute different physical effects efficiently. Unlike previous approaches, the diffraction effects are computed exactly in the forward direction with no restriction on the angle of propagation. This results in more accurate modelling of sonic boom near the carpet edge. Heterogenous flow effects are incorporated with a wide-angle parabolic approximation. Nonlinear propagation is computed with a Burgers-Hayes method. Turbulence in the medium is constructed by the method of random Fourier modes. The turbulence spectra are constructed using an altitude dependent 3-dimensional von Karman spectrum. The turbulence field is considered frozen, as the eddy turnover time in the atmospheric boundary layer is generally much larger than the propagation time of the acoustic wave. Comparisons with benchmark predictions from the PCBoom software are conducted. The benchmark cases provide insight into the accuracy of the current prediction code for non-turbulent atmospheres. Future work is discussed regarding the prediction of shaped booms, boundary conditions, and predictions near the lateral extent of the boom carpet where previous approaches fail to capture the diffraction effects.

sonic boom↗
NASA NTRS1997

Boundary Layer Receptivity to Three-Dimensional Freestream Disturbances at Two-Dimensional Roughness

The receptivity of a laminar boundary layer to an isolated three-dimensional convected disturbance is investigated in a low-speed wind tunnel experiment. The disturbance is created by the short-duration pulsed displacement of a small low-aspect-ratio wing located upstream of a flat plate. The height of the wing is set so that the convected disturbance grazes the edge of the flat-plate boundary layer. A receptivity site is provided by a two-dimensional roughness strip on the surface of the plate. The different propagation speeds of acoustic, convected and instability waves cause the various wave packets from the pulsed displacement to arrive at a downstream measurement station at different times, separating the phenomena and allowing them to be studied independently. Ensemble- averaged measurements are made with and without roughness on the plate. Preliminary analysis of the measurements suggest the presence of a two-dimensional T-S wave packet arising from an interaction between an acoustic wave and the roughness, and a three-dimensional T-S wave packet arising from an interaction between the localized convected disturbance and the roughness strip. The growth rates and spatial characteristics of the disturbances and the instability wave packets are measured as they propagate downstream.

Dietz, Anthony↗
NASA NTRS1972

A study of turbulence generation in a hypersonic boundary layer.

Extensive hot wire correlation and disturbance convection velocity measurements have been obtained in the fully developed turbulent hypersonic boundary layer on a cone-ogive-cylinder at a free-stream Mach number of 7.2 in the Ames 3.5-foot hypersonic wind tunnel. These data suggest that turbulence production in hypersonic boundary layers may be created by highly three-dimensional disturbances originating close to the wall, similar to previous observations in incompressible flows.

Owen, F. K.↗
NASA NTRS1977

Prediction of laminar and turbulent boundary layer flow separation in V/STOL engine inlets

The paper provides a theoretical description of the development of the boundary layer on the lip and diffuser surface of a subsonic inlet at arbitrary operating conditions of mass flow rate, freestream velocity and incidence angle. Both laminar separation on the lip and turbulent separation in the diffuser are discussed. The agreement of the theoretical results with model experimental data illustrates the capability of the theory to predict separation. The effects of throat Mach number, inlet size, and surface roughness on boundary-layer development and separation are illustrated.

Chou, D. C.↗
NASA NTRS1974

Shock-induced separation of adiabatic turbulent boundary layers in supersonic axially symmetric internal flow

An experimental investigation at Mach 4 of shock-induced turbulent boundary layer separation at the walls of axially symmetric flow passages is discussed, with particular emphasis placed on determining the shock strengths required for incipient separation. The shock waves were produced by interchangeable sting-mounted cones placed on the axes of the flow passages and aligned with the freestream flow. The interactions under study simulate those encountered in axially symmetric engine inlets of supersonic aircraft. Knowledges of the shock strengths required for boundary layer separation in inlets is important since for shocks of somewhat greater strength rather drastic alterations in the inlet flow field may occur.

Page, R. J.↗
NASA NTRS1950

Extension of boundary-layer heat-transfer theory to cooled turbine blades

An equation for average heat transfer of a surface was derived when the boundary layer changed from laminar to turbulent. Influences on the heat transfer through a laminar boundary layer of Mach number, temperature ratio (gas temperature divided by wall temperature), and exponents of gas-property temperature relations were shown to be relatively small for air with Mach numbers less than 2 and temperature ratios between 1 and 4. Good agreement was obtained with experimental results from cylinders, an airfoil, and turbine blades.

HEAT TRANSFER, AERODYNAMIC↗
NASA NTRS1993

Boundary layer transition detection by luminescence imaging

In recent experiments we have demonstrated the feasibility of a new approach to boundary layer transition detection. This new approach employs the temperature dependence of certain photoluminescent materials in the form of a surface coating or 'paint' to detect the change in heat transfer characteristics that accompany boundary layer transition. The feasibility experiments were conducted for low subsonic to transonic Mach numbers on two-dimensional airfoil and flat plate configurations. Paint derived transition locations were determined and compared to those obtained from Preston pressure probe measurements. Artificial heating of the models was used to obtain transition temperature signatures suitable for the instrumentation available to us. Initial estimates show, however, that passive kinetic heating at high Mach numbers is a promising alternative.

Mclachlan, B. G.↗
NASA NTRS1981

A quasi-three-dimensional blade surface boundary layer analysis for rotating blade rows

A quasi-three-dimensional approximation has been developed for a blade boundary layer which involves the calculation of the effect of nonzero pressure gradients, turbulent flow, and blade twist, but includes only a simple coupling between streamlines. The resulting set of equations is solved using Keller's box scheme. The solution scheme is checked against available incompressible flow solutions and then applied to a NASA low aspect ratio transonic compressor stage for which extensive experimental and computational data are available. It is found that the three-dimensional boundary layer separates significantly sooner and has a much larger influence on rotor performance than would be expected from a two-dimensional analysis.

Thompkins, W. T., Jr.↗
NASA NTRS1987

Atmospheric surface and boundary layers of the Amazon Basin

Three phases of work were performed: design of and preparation for the Amazon Boundary Layer Experiment (ABLE 2-A); execution of the ABLE 2-A field program; and analysis of the ABLE 2-A data. Three areas of experiment design were dealt with: surface based meteorological measurements; aircraft missions; and project meteorological support. The primary goal was to obtain a good description of the structure of the atmosphere immediately above the rain forest canopy (top of canopy to a few thousand meters), to describe this region during the growing daytime phase of the boundary layer; and to examine the nighttime stratified state. A secondary objective was to examine the role that deep convective storms play in the vertical transport of heat, water vapor, and other trace gases. While significant progress was made, much of the analysis remains to be done.

Garstang, Michael↗
NASA NTRS2003

Optimal Disturbances in Boundary Layers Subject to Streamwise Pressure Gradient

An analysis of the non-modal growth of perturbations in a boundary layer in the presence of a streamwise pressure gradient is presented. The analysis is based on PSE equations for an incompressible fluid. Examples with Falkner- Skan profiles indicate that a favorable pressure gradient decreases the non-modal growth while an unfavorable pressure gradient leads to an increase of the amplification. It is suggested that the transient growth mechanism be utilized to choose optimal parameters of tripping elements on a low-pressure turbine (LPT) airfoil. As an example, a boundary-layer flow with a streamwise pressure gradient corresponding to the pressure distribution over a LPT airfoil is considered. It is shown that there is an optimal spacing of the tripping elements and that the transient growth effect depends on the starting point. The amplification is found to be small at the LPT s very low Reynolds numbers, but there is a possibility to enhance the transient energy growth by means of wall cooling.

Ashpis, David E.↗
NASA NTRS1985

A three-dimensional boundary-layer method for flow over delta wings with leading-edge separation

A three-dimensional, laminar boundary-layer method is applied to the incompressible flow over a slender delta wing at incidence. The predictor-corrector finite-difference scheme of Matsuno is used to difference the governing equations. The method has the advantages that no iterations are required to advance the solution and the cross-flow derivatives are formed independent of the cross-field direction. The difference scheme is demonstrated to yield accurate numerical results when compared to the exact solution of the three-dimensional boundary-layer equations for parabolic flow over a moving flat plate. The method is applied to delta wings of various sweep angles at angles of attack up to 20 deg., with the inviscid solution determined using a higher-order, three-dimensional panel method.

Woodson, S. H.↗
NASA NTRS1982

An experimental investigation and a numerical prediction of a transonic normal shock wave/turbulent boundary layer interaction

Detailed pitot, static and wall pressure measurements have been obtained for a transonic normal shock wave/turbulent boundary layer interaction at free-stream Mach numbers of 1.28, 1.37 and 1.48 and at a constant unit Reynolds number of 4.92 x 10 to the 6th per meter in an axisymmetric, internal flow. Measurements have also been obtained at a unit Reynolds number of 9.84 x 10 to the 6th per meter at a free-stream Mach number of 1.29. The interaction depends very strongly on the Mach number. The effect of Reynolds number under study is small. Flow confinement due to the wind-tunnel wall boundary layer produces a weaker interaction and a much larger supersonic tongue than is observed for planar flows. Comparisons are made with solutions to the time-dependent, mass-averaged, Navier-Stokes equations incorporating a two-equation, Wilcox-Rubesin turbulence model (1980). The computations are in agreement with the experimental results.

Om, D.↗
NASA NTRS1989

Three-dimensional compressible boundary layer calculations to fourth order accuracy on wings and fuselages

Laminar flow control and drag reduction research requires accurate boundary layer solutions as input to the three-dimensional stability analysis procedures currently under development. In support of these major programs, a fourth-order accurate finite difference scheme for solving the three-dimensional, compressible boundary layer equations has been developed and is presented in this paper. The method employs a two-point scheme in the wall normal direction and second order zigzag scheme in the cross flow direction. Accurate procedures to interface with the inviscid results are also presented. The results of applying the procedure to laminar flow on wings and fuselages are presented.

Iyer, Venkit↗
NASA NTRS1974

Comparison of techniques for determination of boundary layer transition in shock wave induced flows

Three methods of determining transition times in shock wave-induced boundary layers are compared. Experimental transition data for these comparisons were obtained in two different shock tube facilities for shock wave Mach numbers from 3.0 to 5.5. Bias in determining transition times by the conventional thin-film temperature history and difficulty in reading schlieren photographs suggest the use of a new heat-flux technique for determining transition. It is suggested that the heat-flux technique be used when possible, because of better sensitivity and a view of the entire transition process and because it reveals important information concerning departures from theoretical laminar boundary-layer development. Transition results presented extend the range of data available in the literature and are in good agreement with existing data through transition Reynolds numbers of 1,000,000.

Hall, J. L.↗