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Results for “BOUNDARY LAYER - EFFECT OF COMPRESSIBILITY”

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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

A fourth-order scheme for the unsteady compressible Navier-Stokes equations

A computational scheme is described which is second-order accurate in time and fourth-order accurate in space (2-4). This method is applied to study the stability of compressible boundary layers. The laminar compressible Navier-Stokes equations are solved with a time harmonic inflow superimposed on the steady state solution. This results in spatially unstable modes. It is shown that the second-order methods are inefficient for calculating the growth rates and phases of the unstable modes. In contrast the fourth-order method yields accurate results on relatively course meshes.

Bayliss, A.↗
NASA NTRS1986

Space-time correlation measurements in a compressible turbulent boundary layer

Single and dual hot-wire probes have been used to accomplish a first look at the large-eddy structure of a supersonic turbulent boundary layer. The investigation was performed in an axisymmetric turbulent boundary layer in the Ames High Reynolds Number wind tunnel. Freestream Mach number was 2.97, the boundary layer was approximately 1.2 cm thick, and Reynolds number = 15,000. Quantities measured were streamwise turbulence intensity and higher order statistics of mass-flow fluctuations, and space-time correlations of instantaneous streamwise velocity at two locations. Conditional sampling of the hot-wire signals was performed to detect the passage and describe the character of the large turbulent eddies. Results indicate that the large scale structure in compressible turbulence is similar to that at low subsonic flow speeds except at the boundary layer edge, where the intermittency region is narrowed in compressible flow.

Robinson, S. K.↗
NASA NTRS1981

Compressible boundary-layer stability calculations for sweptback wings with suction

The stability of the laminar boundary layers on two transonic wings of infinite span with distributed suction is investigated with the compressible, parallel-flow stability theory. Both wings have supercritical airfoil sections; one has a sweep angle of 23 deg, the other of 35 deg. Zero-frequency disturbances are used to represent cross-flow instability, and disturbances with the wavenumber vector aligned with the local flow direction represent traveling-wave instability. In both cases, the maximum spatial amplification rate is used as a measure of the instability. For the suction, distributions with constant mass flux downstream of the starting point are used. The main objective is to determine how the maximum amplification rate varies with the magnitude and starting point of the suction. It is found for both types of disturbances that the maximum amplification rate varies almost linearly with the suction magnitude up to at least the point where the amplification rate is halved. Different starting locations for the suction in the first 4% of the chord were found to affect cross-flow instability, but to have little influence on traveling-wave instability.

Mack, L. M.↗
NASA NTRS2007

On an Asymptotically Consistent Unsteady Interacting Boundary Layer

This paper develops the asymptotic matching of an unsteady compressible boundary layer to an inviscid flow. Of particular importance is the velocity injection or transpiration boundary condition derived by this theory. It is found that in general the transpiration will contain a slope of the displacement thickness and a time derivative of a density integral. The conditions under which the second term may be neglected, and its consistency with the established results of interacting boundary layer are discussed.

Bartels, Robert E.↗
NASA NTRS1984

Simulation of transonic separated airfoil flow by finite difference viscous-inviscid interaction

A finite difference viscous inviscid interaction program has been developed for simulating the separated transonic flow about lifting airfoils, including the wake. In contrast to most interaction programs, this code combines a finite difference boundary layer algorithm with the inviscid program. The recently developed finite difference boundary layer code efficiently simulates attached and reversed compressible boundary layer and wake flows. New viscous inviscid interaction algorithms were also developed to couple the boundary layer code with the inviscid transonic full potential program. Transonic cases with shock induced and trailing edge separation are computed and compared with experimental and Navier-Stokes results.

Vandalsem, W. R.↗
NASA NTRS1988

Slip velocity method for three-dimensional compressible turbulent boundary layers

A slip velocity method for 2-D incompressible turbulent boundary layers was presented in AIAA Paper 88-0137. The inner part of the boundary layer was characterized by a law of the wall and a law of the wake, and the outer part was characterized by an arbitrary eddy viscosity model. In the present study for compressible flows, only a law of the wall is considered. The problem of 2-D compressible flow is treated first; then the extension to 3-D flow is addressed. A formulation for primitive variables is presented.

Barnwell, Richard W.↗
NASA NTRS1990

A defect stream function formulation for compressible turbulent boundary layers

Progress to date on the development of a method for turbulent, wall-bounded flow which uses the defect stream function formulation in the outer layer and an analytic law of the wall and wake formulation in the inner region is reviewed. This two-formulation approach avoids the need to computationally resolve the high-gradient inner layer. One of the most appealing recent developments is the transformation of the compressible governing equation for the defect stream function into a linear, second-order differential equation which has analytic solutions for many problems of practical interest. Numerical and analytic results for incompressible and compressible flows are shown to be in excellent agreement with experimental results. In this paper the two-formulation approach is applied to primitive-variable computations. Excellent comparisons with experiment are presented for two compressible flat plate flows.

Barnwell, Richard W.↗
NASA NTRS1979

On the stability of the boundary layer on a transonic swept wing

Both incompressible and compressible linear stability theory are applied to the three-dimensional compressible boundary layer on a particular transonic sweptback wing of infinite span. A spatial stability theory is used which identifies the growth direction with the real part of the complex angle of the group velocity. It is found that in the forward, but not the rear, crossflow instability region, the maximum amplification rates of the steady disturbances may be calculated to within about 10% by the incompressible stability theory. There is little difference between the sixth and eighth-order compressible theories. The maximum amplification rate of the steady disturbances at any chordwise station is closely related to the maximum crossflow at that station independent of the Reynolds number. For other than crossflow instability, there can be large differences between the incompressible and compressible theories, both as to the amplification rate and the angle of the wavenumber vector for maximum instability. Amplitude ratios of individual wave components are obtained by integrating the spatial amplification rate along the growth direction subject to the constraint that the wavenumber vector is irrotational. This procedure yields steady disturbances aligned with the local potential flow direction whose wavelengths are nearly independent of downstream distance.

Mack, L. M.↗
NASA NTRS1971

Calculation of compressible nonadiabatic boundary layers in laminar, transitional and turbulent flow by the method of integral relations

A computer program was developed to do the calculations for two-dimensional or axisymmetric configurations from low speeds to hypersonic speeds with arbitrary streamwise pressure, temperature, and Mach number distributions. Options are provided for obtaining initial conditions either from experimental information or from a theoretical similarity solution. The transition region can be described either by an arbitrary distribution of intermittency or by a function based on Emmons' probability theory. Correlations were developed for use in estimating the parameters of the theoretical intermittency function. Correlations obtained from other sources are used for estimating the transition point. Comparisons were made between calculated and measured boundary layer quantities for laminar, transitional, and turbulent flows on flat plates, cones, cone flares, and a waisted body of revolution. Excellent agreement was obtained between the present theory and two other theories based on the method of finite differences. The intermittency required to reproduce some experimental heat transfer results in hypersonic flow was found to be quite different from the theoretical function. It is suggested that the simple probability theory of Emmons may not be valid for representing the intermittency of hypersonic transitional boundary layers and that the program could be useful as a tool for detailed study of the intermittency of the transition region.

Kuhn, G. D.↗
NASA NTRS1978

In-flight Compressible Turbulent Boundary Layer Measurements on a Hollow Cylinder at a Mach Number of 3.0

Skin temperatures, shearing forces, surface static pressures, and boundary layer pitot pressures and total temperatures were measured on a hollow cylinder 3.04 meters long and 0.437 meter in diameter mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 and at wall-to-recovery temperature ratios of 0.66 to 0.91. The free stream Reynolds number had a minimal value of 4.2 million per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. Boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor of 1.11 was obtained from the measured heat transfer and skin friction data. The skin friction coefficients predicted by the theory of van Driest were in excellent agreement with the measurements. Theoretical heat transfer coefficients, in the form of Stanton numbers calculated by using a modified Reynolds analogy between skin friction and heat transfer, were compared with measured values. The measured velocity profiles were compared to Coles' incompressible law-of-the-wall profile.

Quinn, R. D.↗
NASA NTRS1984

Remarks on disputed numerical results in compressible boundary-layer stability theory

Past work on the influence of Mach number on the viscous and inviscid instability of flat-plate boundary layers is reviewed, and new spatial calculations are presented. These calculations support the previous view that viscosity is only stabilizing for both two- and three-dimensional first-mode waves above M1 = 3.0, and for second-mode waves at all Mach numbers. It is concluded that the calculations of Wazzan, Taghavi, and Keltner that show viscous instability at M1 = 6.0 for first-mode 50 deg waves, and at M1 = 3.0 for two-dimensional second-mode waves, are not correct.

Mack, L. M.↗
NASA NTRS1985

Numerical study of the response of a compressible, turbulent boundary layer to a short region of surface curvature

Calculations were performed to investigate the supersonic flow of a turbulent boundary layer over short regions of concave surface curvature. Upstream of each curved surface the freestream Mach number was 2.9. Three different constant radii of curvature models were investigated to cover a range of curvatures and turning angles. The numerical technique solved the full, Reynolds-averaged Navier-Stokes equations using two different turbulence models: the algebraic eddy viscosity model due to Baldwin and Lomax (1978), and the one-equation model to Rubesin (1976). The calculations were compared with the experimental data of Taylor and Smits (1984), and the agreement was very satisfactory, especially for the computations using the one-equation model.

Degani, D.↗
NASA NTRS1987

The response of a compressible turbulent boundary layer to short regions of concave surface curvature

An experimental investigation of the supersonic flow of a boundary layer over short regions of concave surface curvature, using radii of curvature equivalent to 10 and 50 (models I and II, respectively) initial boundary layer thicknesses and 8-deg turning angle, notes that the boundary layer experienced a strong adverse pressure gradient as the boundary layer passed through the curved region. Attention is given to the mean and turbulent field for each flow, with comparisons of results to those for an 8-deg ramp. In the ramp and model I, turbulence levels increased significantly, together with structural parameters; in model II, despite an increase in turbulence levels, structural parameters remained essentially unchanged. The former results suggest that the perturbation is rapid, and describable in terms of total rather than local strains.

Jayaram, Mohan↗