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Results for “Compressible boundary layer”

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 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

User's manual for three dimensional boundary layer (BL3-D) code

An assessment has been made of the applicability of a 3-D boundary layer analysis to the calculation of heat transfer, total pressure losses, and streamline flow patterns on the surface of both stationary and rotating turbine passages. In support of this effort, an analysis has been developed to calculate a general nonorthogonal surface coordinate system for arbitrary 3-D surfaces and also to calculate the boundary layer edge conditions for compressible flow using the surface Euler equations and experimental data to calibrate the method, calculations are presented for the pressure endwall, and suction surfaces of a stationary cascade and for the pressure surface of a rotating turbine blade. The results strongly indicate that the 3-D boundary layer analysis can give good predictions of the flow field, loss, and heat transfer on the pressure, suction, and endwall surface of a gas turbine passage.

Anderson, O. L.↗
NASA NTRS1995

Experimental study of the use of vortex generators to reduce fluctuating pressure loads in shock wave turbulent boundary layer interactions

Fluctuating wall pressure measurements have been made in a separated shock wave/turbulent boundary layer interaction produced by an unswept compression corner in a Mach 5 flow. Wheeler doublet vortex generators were placed 15.8 boundary layer thicknesses upstream of the corner to study their effect on the fluctuating pressure loads produced by the translating separation shock. The vortex generators produced significant three-dimensionality in an otherwise two-dimensional interaction. They reduced the upstream influence and the length of the region of shock motion by 60% and 64%, respectively, decreased the maximum wall pressure rms by 23%, and shifted the fluctuations to a higher frequency band. The maximum fraction of energy in the 100-500 Hz frequency band is decreased by 11%. These changes are due to a fuller boundary layer profile, a weaker separation shock, and increased boundary layer turbulence causing increased separation shock jitter.

Barter, John W.↗
NASA NTRS1985

Calculation of three-dimensional boundary layers on rotating turbine blades

An assessment has been made of the applicability of a three dimensional boundary layer analysis to the calculation of heat transfer and streamline flow patterns on the surfaces of both stationary and rotating turbine passages. In support of this effort, an analysis has been developed to calculate a general nonorthogonal surface coordinate system for arbitrary three dimensional surfaces and also to calculate the boundary layer edge conditions for compressible flow using the surface Euler equations and experimental pressure distributions. Using available experimental data to calibrate the method, calculations are presented for the endwall, and suction surfaces of a stationary cascade and for the pressure surface of a rotating turbine blade. The results strongly indicate that the three dimensional boundary layer analysis can give good predictions of the flow field and heat transfer on the pressure, suction, and endwall surfaces in a gas turbine passage.

Anderson, O. L.↗
NASA NTRS1994

Noise Radiation by Instability Waves in Coaxial Jets

In this paper predictions are made for the noise radiation from supersonic coaxial jets. The noise in the downstream arc of a supersonic jet is dominated by highly directional radiation from the supersonically convecting large scale structures in the jet mixing layer. Since the mean flow is not described easily in terms of simple analytic functions, a numerical prediction is made for its development. The compressible Reynolds-averaged boundary layer equations in cylindrical polar coordinates are solved. A mixing length turbulence model is used. Empirical correlations are developed the effects of velocity and temperature ratio and Mach number. Both normal and inverted velocity profiles are considered. Comparisons with measurements for both single and coaxial jets show good agreement. The large scale structures are modeled as instability waves. The noise radiation generated by the instability waves is determined by a matching between the inner instability wave solution and the outer acoustic solution. Predictions are made for the differences between the noise radiated by coaxial jets with different operating conditions and a single equivalent jet with the same exit area, thrust, and mass-flow.

Dahl, Milo D.↗
NASA NTRS1992

Structure of supersonic turbulent flow past a swept compression corner

The structure of the shock wave/turbulent boundary-layer interaction generated by a 3D swept compression corner has been investigated through a combined experimental and theoretical research program. The flowfield geometry is defined by the streamwise compression angle alpha and the sweep angle lambda of the corner. The present study examines two different configurations, namely (alpha, lambda) = (24 deg, 40 deg) and (24 deg, 60 deg) at Mach 3 and Re sigma infinity about 9 x 10 exp 5. The theoretical model is the 3D Reynolds-averaged compressible Navier-Stokes equations with turbulence incorporated using a turbulent eddy viscosity. The calculated flowfields display general agreement with experimental data for surface pressure and good agreement with experimental flowfield profiles of pitot pressure and yaw angle. The principal feature of the flowfield is a large vortical structure approximately aligned with the corner. The entrainment of incoming fluid into the vortical structure is strongly affected by the sweep angle lambda. Viscous (turbulent and molecular) effects appear to be important only in the immediate vicinity of the surface and in an isolated region within the interaction and near the corner.

Knight, Doyle D.↗
NASA NTRS1951

Aerodynamic Characteristics of a Slender Cone-cylinder Body of Revolution at a Mach Number of 3.85

An experimental investigation of the aerodynamics of a slender cone-cylinder body of revolution was conducted at a Mach number of 3.85 for angles of attack of 0 degree to 10 degrees and a Reynolds number of 3.85x10(exp 6). Boundary-layer measurements at zero angle of attack are compared with the compressible-flow formulations for predicting laminar boundary-layer characteristics. Comparison of experimental pressure and force values with theoretical values showed relatively good agreement for small angles of attack. The measured mean skin-friction coefficients agreed well with theoretical values obtained for laminar flow over cones.

FLOW, LAMINAR↗
DOE OSTI2024

On the accuracy of compressibility transformations

This study highlights the importance of satisfying the eddy viscosity equivalence below the logarithmic layer, to deriving accurate compressibility transformations. First, we analyze the ability of known transformations to satisfy the eddy viscosity equivalence and show that the accuracy of these transformations is strongly dependent on this ability. Second, in a step-by-step manner, we devise new transformations that satisfy this hypothesis. An approach based on curve fitting of the incompressible Direct Numerical Simulation data for eddy viscosity profiles below the logarithmic layer provides an extremely accurate transformation, which motivates self-contained methods, making use of mixing length formulas in the inner region. It is shown that the accuracy of existing transformations can be significantly improved by applying these ideas, below the logarithmic layer. Motivated by the effectiveness of the formulations derived from eddy viscosity equivalence, we introduce a new integral transformation based on Reynolds number equivalence between compressible and incompressible flows. This approach is based on defining a new compressible velocity scale, which affects the accuracy of transformations. Several choices for the velocity scale are tested, and in each attempt, it is shown that the eddy viscosity equivalence plays a very important role for the accuracy of compressibility transformations.

42 ENGINEERING↗
NASA NTRS1971

Calculation of turbulent boundary layers with heat transfer and pressure gradient utilizing a compressibility transformation. Part 1: Summary report

The analysis uses a compressibility transformation and utilizes higher order closure rules to complete the transformation. By requiring that the momentum equations in differential form be satisfied at the wall and at the sublayer edge, correspondence rules are obtained which relate the variable property (VP) flow to a constant property (CP) flow in which mass transfer and pressure gradient occur simultaneously. A new CP formulation is developed and numerical results for a variety of cases are presented. Comparisons with earlier forms of the transformation and with experiment are included. For the zero pressure gradient case some differences between the various predictions are observed. For the several pressure gradient cases examined, the results are found to be essentially identical to those given by first order closure rules; i.e., by a form of transformation which relates the VP flow to a CP flow with pressure gradient but zero mass transfer.

Economos, C.↗
NASA NTRS1981

Separation behavior of boundary layers on three-dimensional wings

An inverse boundary layer procedure for calculating separated, turbulent boundary layers at infinitely long, crabbing wing was developed. The procedure was developed for calculating three dimensional, incompressible turbulent boundary layers was expanded to adiabatic, compressible flows. Example calculations with transsonic wings were made including viscose effects. In this case an approximated calculation method described for areas of separated, turbulent boundary layers, permitting calculation of this displacement thickness. The laminar boundary layer development was calculated with inclined ellipsoids.

Stock, H. W.↗
NASA NTRS1976

A complete model of turbulence

A set of constitutive equations suitable for a priori computation of turbulent shear flows has been developed. Since no properties of a given turbulent flow need be known in advance in order to obtain a solution, the equations comprise a complete model of turbulence. Perturbation analysis shows that the model predicts a composite five-layer structure for an incompressible turbulent boundary layer, viz, a defect layer, a law-of-the-wall layer, a viscous sublayer, a near-surface roughness layer, and a viscous superlayer at the boundary-layer edge. Analysis of the defect layer demonstrates the key improvement of the model over its predecessor, the Saffman-Wilcox two-equation model of turbulence. Examination of model-predicted sublayer structure yields model-parameter boundary conditions appropriate for surfaces with roughness and mass injection. Results of numerical computations of compressible and incompressible equilibrium boundary layers show that, for such flows, the model is as accurate as mixing-length theory. Applications to transitional boundary layers and to nonequilibrium relaxation of a boundary layer passing from a rough to a smooth surface indicate that the model's applicability extends far beyond that of mixing-length theory's.

Wilcox, D. C.↗
NASA NTRS1982

On Kaplun's optimal co-ordinates

Optimal co-ordinates were introduced by Kaplun (1954) as a result of his study on the role of co-ordinate systems in boundary-layer theory. In this paper the basic ideas of optimal co-ordinates are examined and many restrictions of Kaplun's optimal co-ordinates are removed. His rule for constructing optimal co-ordinates is found to apply unaltered to axisymmetric flows, to flows with oncoming streams containing vorticity, to free boundary layers such as jets, to free-convection flows and to compressible flows. It is also extended to unsteady boundary layers, and to three-dimensional boundary layers using a pair of stream functions.

Legner, H. H.↗
NASA NTRS1991

Evaluation of a technique to generate artificially thickened boundary layers in supersonic and hypersonic flows

The feasibility of using a contoured honeycomb model to generate a thick boundary layer in high-speed, compressible flow was investigated. The contour of the honeycomb was tailored to selectively remove momentum in a minimum of streamwise distance to create an artificially thickened turbulent boundary layer. Three wind tunnel experiments were conducted to verify the concept. Results indicate that this technique is a viable concept, especially for high-speed inlet testing applications. In addition, the compactness of the honeycomb boundary layer simulator allows relatively easy integration into existing wind tunnel model hardware.

Porro, A. R.↗
NASA NTRS1959

Experimental Investigation of the Pressure Rise Required for the Incipient Separation of Turbulent Boundary Layers in Two-Dimensional Supersonic Flow

An experimental investigation has been made of turbulent boundary-layer separation associated with compression corners, curved surfaces of various radii, and incident shock waves. The purpose of the investigation was to provide design information, and to define significant physical trends, which would aid in the prediction of turbulent separation for various aerodynamic devices, such as compressor blades, flaps, spoilers, and diffusers. A characteristic change in the longitudinal static-pressure distribution (i.e., a change from a curve with one inflection point to a curve with three inflection points) was employed to detect the occurrence of separation. The effects of Reynolds number (10(exp 6) to 10(exp 7) per foot or l.5 x 10(exp 4) to 7.5 x 10(exp 4) based upon boundary-layer thickness) and Mach number (1.6 to 4.2) on the onset of turbulent boundary-layer separation were investigated. The pressure gradient of the boundary-layer flow ahead of the interaction region was essentially zero. The results show a considerable effect of Mach number on the pressure rise for incipient separation for all configurations. For a curved-surface model, the static pressure-rise ratio required to cause separation varied from about 2.5, at a Mach number of 2 to about 16, at a Mach number of 3.5. A substantial effect of Reynolds number on the pressure rise for incipient separation was observed in the upper Mach number range and in the lower Reynolds number range; namely, the pressure rise required for separation decreased with increasing Reynolds number. For low Mach numbers and high Reynolds numbers, there appeared to be no Reynolds number effect. The effects of Mach number and of Reynolds number were similar for all models. Model shape was also found to be an important variable affecting the onset of separation. Large gains were realized in the pressure-rise ratio with no separation when the radius of curvature of the model surface was increased. At a Mach number of 3.4, for instance, the pressure-rise ratio with no separation increased from about 5 to 15 as a result of an increase in the radius of curvature from approximately 0 to 30 boundary-layer thicknesses.

Kuehn, Donald M.↗
NASA NTRS1978

Magnetic fields and flows between 1 and 0.3 AU during the primary mission of Helios 1

Magnetic phenomena and coronal holes observed by Helios 1 during nearly four rotations of the sun are discussed. It is estimated that the magnetic field intensity of open field lines in some of the coronal holes (which changed in shape and latitude from one rotation to the next) was of the order of 10-20 G during the studied period, December 1974 to March 1975. Magnetic field polarity, cold magnetic enhancements, and longitudinal widths of the stream are described. It is concluded that between the sun and 0.3 AU there was a diverging stream surrounded by a thin boundary layer in which there was large velocity shear. There is evidence for compression of the magnetic field in the western boundary layer.

Burlaga, L. F.↗
NASA NTRS1971

Experimental investigation of the performance of vortex generators mounted in the supersonic portion of a mixed-compression inlet

Vortex generators were investigated as a means of controlling the throat boundary layer in a Mach 2.5 mixed-compression inlet. With or without throat bleed, the generators were unable to produce the inlet peak recoveries available with good performance bleed. Generators placed on the unbled inlet centerbody ahead of the cowl shock impingement point degraded inlet performance. Generators placed further aft on the unbled centerbody, but ahead of the throat, improved inlet performance at conditions where the terminal shock could cause boundary layer separation. Adding generators to the inlet with throat bleed reduced performance. Bleed aft of the generators degraded the vortex flow field.

Michell, G. A.↗
NASA NTRS1973

A study of the effect of a boundary layer profile on the dynamic response and acoustic radiation of flat panels

The response of a thin, elastic plate to a harmonic force which drives the plate from below and a compressible air stream with a viscous boundary layer flowing parallel to the upper surface along the length was investigated. Equations governing the forced response of the coupled plate-aerodynamic system are derived along with appropriate boundary conditions. Calculations of basic solution parameters for a linear velocity profile and for a Blasius profile showed that the same system response could be obtained from each profile if appropriate values of boundary layer thickness were chosen for each profile.

Mixson, J. S.↗