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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

Effect of Free Stream Turbulence and Other Vortical Disturbances on a Laminar Boundary Layer

This paper is concerned with the effect of free-stream turbulence on the pretransitional flat-plate boundary layer. It is assumed that either the turbulence Reynolds number or the downstream distance (or both) is small enough so that the flow can be linearized. The dominant disturbances in the boundary layer, which are of the Klebanoff type, are governed by the linearized unsteady boundary-region equations, i.e., the Navier Stokes equations with the streamwise derivatives neglected in the viscous and pressure-gradient terms. The turbulence is represented as a superposition of vortical free-stream Fourier modes, and the corresponding individual Fourier component solutions to the boundary-region equations are obtained numerically. The results are then superposed to compute the root mean square of the fluctuating streamwise velocity in the boundary layer produced by the actual free-stream turbulence. The calculated boundary-layer disturbances are in good quantitative agreement with the experimentally observed Klebanoff modes when strong low-frequency anisotropic effects are included in the free-stream turbulence spectrum. We discuss some additional effects that may need to be accounted for in order to obtain a complete description of the Klebanoff modes.

Goldstein, Marvin E.↗
NASA NTRS1959

A Simple Method for Determining Heat Transfer, Skin Friction, and Boundary-Layer Thickness for Hypersonic Laminar Boundary-Layer Flows in a Pressure Gradient

A procedure based on the method of similar solutions is presented by which the skin friction, heat transfer, and boundary-layer thickness in a laminar hypersonic flow with pressure gradient may be rapidly evaluated if the pressure distribution is known. This solution, which at present is. restricted to power-law variations of pressure with surface distance, is presented for a wide range of exponents in the power law corresponding to both favorable and adverse pressure gradients. This theory has been compared to results from heat-transfer experiments on blunt-nose flat plates and a hemisphere cylinder at free-stream Mach numbers of 4 and 6.8. The flat-plate experiments included tests made at a Mach number of 6.8 over a range of angle of attack of +/- 10 deg. Reasonable agreement of the experimental and theoretical heat-transfer coefficients has been obtained as well as good correlation of the experimental results over the entire range of angle of attack studied. A similar comparison of theory with experiment was not feasible for boundary-layer-thickness data; however, the hypersonic similarity theory was found to account satisfactorily for the variation in boundary-layer thickness due to local pressure distribution for several sets of measurements.

Bertram, Mitchel H.↗
NASA NTRS1979

A model for the noise radiated by turbulent boundary layers and its interaction with laminar layers in supersonic flow

Motivated by the noise-boundary layer transition problem in supersonic flow, the interaction between incoming aerodynamic noise and the laminar boundary layer was studied. The noise field was modeled as a Mach wave radiation field consisting of discrete waves emanating from coherent turbulent entities moving downstream within the supersonic turbulent boundary layer. The individual disturbances are likened to miniature sonic booms and the laminar boundary layer is strafed by the waves as the sources move downstream. The mean, autocorrelation, and power spectral density of the field are expressed in terms of the wave shapes and their average arrival rates. The emphasis in the interaction analysis is on the behavior of the shocklets in the noise field. The shocklets are shown to be focused by the laminar boundary layer in its outer region. Borrowing wave-propagation terminology, this region is termed the caustic region. Using scaling laws from sonic boom work, focus factors at the caustic are estimated to vary from 2 to 6 for incoming shocklet strengths of 1 to .01 percent of the free stream pressure level. The situation regarding experimental evidence of the caustic region is reviewed. It is not clear whether the focusing effectively promotes early transition. Additional observations concerning the interaction problem are presented.

Schopper, M. R.↗
NASA NTRS1979

A model for the noise radiated by turbulent boundary layers and its interaction with laminar layers in supersonic flow

Motivated by the noise-boundary layer transition problem in supersonic flow, the interaction between incoming aerodynamic noise and the laminar boundary layer has been studied. The noise field has been modeled as a Mach wave radiation field consisting of discrete waves emanating from coherent turbulent entities moving downstream within the supersonic turbulent boundary layer. The individual disturbances are likened to miniature sonic booms and the laminar boundary layer is strafed by the waves as the sources move downstream. The mean, autocorrelation, and power spectral density of the field are expressed in terms of the wave shapes and their average arrival rates. The emphasis in the interaction analysis is on the behavior of the shocklets in the noise field. The shocklets are shown to be focused by the laminar boundary layer in its outer region. Borrowing wave-propagation terminology, this region is termed the caustic region. Using scaling laws from sonic boom work, focus factors at the caustic are estimated to vary from 2 to 6 for incoming shocklet strengths of 1 to .01 percent of the free stream pressure level.

Schopper, M. R.↗
NASA NTRS1983

A study of non-unique solutions of the two-dimensional boundary layer equations at laminar separation and reattachment points

Nonunique laminar boundary layer equation solutions in direct problem calculations are identified for the case of accelerating flow. As a separation or reattachment point is approached, the multiple solutions approach each other and become identical. The computer code used to generate these results was developed for the solution of compressible, laminar or turbulent boundary layer, and free wake problems, in either direct or inverse mode. Similarity solutions in either a primitive variable or a stream function form are possible, and the resulting equations are solved by means of a modified Keller's Box scheme in which the energy equation and turbulence modeling equations are solved simultaneously with the continuity and momentum equations. Examples illustrating the nature of the solutions at the separation and reattachment points are presented.

Drela, M.↗
NASA NTRS1978

Compressible laminar boundary-layer flow

Program computes compressible laminar boundary-layer flow over yawed infinite wing including distributed suction. Computations are useful in efforts to implement boundary-layer suction on wing to maintain laminar flow in order to reduce net drag. Program contains number of approximations to simplify analysis; yet results are accurate enough to be useful, particularly in preliminary design phase.

Carter, J. E.↗
NASA NTRS1984

Interacting boundary-layer solutions for laminar separated flow past airfoils

Numerical solutions of the interacting laminar boundary layer equations are presented for two symmetric airfoils at zero incidence: the NACA 0012 and the NACA 66 sub 3-108 airfoils. The potential flow was computed using Carlson's code, and viscous interaction was treated following a Hilbert integral scheme due to Veldman. Effects of various grid parameters are studied, and pressure and skin friction distributions are compared at several Reynolds numbers. For the NACA 0012 airfoil, Reynolds number is varied from a value just below separation (R sub N = 3000) to a value for which extensive separation occurs (R sub N = 100,000). For the 66 sub 3-018 airfoil, results are given at intermediate values (R sub N - 10,000 and 40,000). The method fails to converge for greater values of Reynolds number, corresponding to the development of very thin well separated shear layers where transition to turbulence would occur naturally.

Burggraf, O. R.↗
NASA NTRS1935

An Application of the Von Karman-millikan Laminar Boundary-layer Theory and Comparison with Experiment

The von Karman-Millikan theory of laminar boundary layers presented in NACA Technical Report No. 504 is applied to the laminar boundary layer about an elliptic cylinder on which boundary-layer and pressure-distribution measurements were made. An outline of the procedure of the von Karman-Millikan method is given. Good agreement is obtained between the calculated and experimental results, indicating that the method may be applied generally to the laminar boundary layer about any body provided that an experimentally determined pressure distribution is available. It appears that for all Reynolds Number above 24,000 the separation point for the elliptic cylinder should occur at a constant distance behind the point of minimum pressure, provided that the boundary layer does not become turbulent.

Von Doenhoff, Albert E↗
NASA NTRS1991

Flight experiments measuring boundary-layer disturbances in laminar flow and correlation with stability analysis

Flight test results are reported from an experiment designed to study the detailed growth of disturbances in the laminar boundary layer. A gloved wing section incorporating closely-spaced flush-mounted streamwise-located instrumentation for measuring instability frequencies and amplitude growths as well as pressure distributions was used. The growth of Tollmien-Schlichting (T-S) and crossflow instabilities is predicted by the linear e exp n method and compared to the measured boundary-layer disturbance frequencies. The predictions showed good agreement with the measured data. The results exhibited fair agreement with previous n(T-S) and n(CF) flight correlations for several of the conditions analyzed. It is inferred from the high n(T-S) values for these data that moderately swept wings at compressible speeds can withstand higher combinations of n(T-S) and n(CF) values and still remain laminar than previously thought.

Lee, Cynthia C.↗
NASA NTRS1976

Real gas scale effects on hypersonic laminar boundary-layer parameters including effects of entropy-layer swallowing

Inviscid and viscous (laminar boundary-layer) flow-field calculations under perfect gas hypersonic wind tunnel and equilibrium real gas flight conditions are presented for the windward centerline of the Rockwell International 139 Space Shuttle Orbiter at a 30-deg angle of attack. Correlation parameters for laminar boundary-layer edge quantities and surface heat transfer are developed which properly account for entropy-layer-swallowing effects under both real and perfect gas conditions. Some implications of the proposed correlation parameters on boundary-layer transition are discussed.

Adams, J. C., Jr.↗