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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Development of an algebraic stress/two-layer model for calculating thrust chamber flow fields

Following the consensus of a workshop in Turbulence Modeling for Liquid Rocket Thrust Chambers, the current effort was undertaken to study the effects of second-order closure on the predictions of thermochemical flow fields. To reduce the instability and computational intensity of the full second-order Reynolds Stress Model, an Algebraic Stress Model (ASM) coupled with a two-layer near wall treatment was developed. Various test problems, including the compressible boundary layer with adiabatic and cooled walls, recirculating flows, swirling flows and the entire SSME nozzle flow were studied to assess the performance of the current model. Detailed calculations for the SSME exit wall flow around the nozzle manifold were executed. As to the overall flow predictions, the ASM removes another assumption for appropriate comparison with experimental data, to account for the non-isotropic turbulence effects.

Chen, C. P.↗
NASA NTRS1990

Theory on instability and transition

The fundamental fluid dynamics governing instability and transition to turbulence in boundary layers are considered, and attention is focused on the key aspects of nonlinear dynamics central to the transition process and to turbulent boundary-layer phenomena. Emphasis is placed on truly nonlinear theories, in which the boundary layer mean-flow profile is completely altered from its original form. Nonlinear TS transitions, Euler-stage interactions, and vortex/wave interactions are discussed, and compressible boundary layers are analyzed. Connections with experiments and computations are outlined, along with overall trends including the extension of the nonlinear theory and the advancement in the compressible and other flow regimes.

Smith, Frank T.↗
NASA NTRS1996

Effect of Density Gradients in Confined Supersonic Shear Layers: Three-Dimensional Modes - 2

The effect of density gradients on three-dimensional (3-D) supersonic acoustic modes and 3-D subsonic modes in a compressible confined shear layer were investigated using linear analysis. The compressible Rayleigh equation with the solution of the steady compressible boundary layer equations as its basic profiles was solved for two different density ratios, 1.398 and 3.0, and several spanwise wave numbers. For a density ratio of 1.398, the 2-D Co, mode had the highest growth rate for a fixed aspect ratio BIH=2. At the higher-density ratio, the 3-D mode had a slightly higher maximum growth rate than the two-dimensional (2-D) mode, and the maximum growth rate was at a location where the phase speed was supersonic with respect to the fast stream.

Peroomian, Oshin↗
NASA NTRS1990

A three-dimensional, compressible, laminar boundary-layer method for general fuselages. Volume 2: User's manual

This user's manual contains a complete description of the computer programs developed to calculate three-dimensional, compressible, laminar boundary layers for perfect gas flow on general fuselage shapes. These programs include the 3-D boundary layer program (3DBLC), the body-oriented coordinate program (BCC), and the streamline coordinate program (SCC). Subroutine description, input, output and sample case are discussed. The complete FORTRAN listings of the computer programs are given.

Wie, Yong-Sun↗
NASA NTRS1951

Boundary Layer Effect on the Surface Pressure of an Infinite Cone in Supersonic Flow

The theory of Taylor and Maccoll (Ref,1) gives the surface pressure on an infinite cone in supersonic flow as a function of the cone vertex angle and the free stream Mach number and static pressure for a gas of vanishing viscosity. When a slender conical probe is used together with an impact pressure probe to determine the static pressure and Mach number in a low density gas stream, it is desirable to have some theoretical estimate of the effect of viscous boundary layer on the probe readings. Theoretical and experimental results with respect to impact probes have been presented in Refs. 5 and 6. A simple approximation for a conical probe based on linearized supersonic flow and compressible boundary layer theory is presented here.

Lin, T. C.↗
NASA NTRS1982

Stability of compressible three-dimensional boundary-layer flows

For compressible three-dimensional flow, the method of multiple scales to formulate the three-dimensional stability problem and determine the partial-differential equations governing variations of the amplitude and complex wavenumbers is used. A method for following one specific wave along its trajectory to ascertain the characteristics of the most unstable disturbance is proposed. Numerical results using the flow over the X-21 wing as calculated from the Kaups-Cebeci code will be presented.

Reed, H. L.↗
NASA NTRS1992

Computation of unsteady flows over airfoils

Two methods are described for calculating unsteady flows over rapidly pitching airfoils. The first method is based on an interactive scheme in which the inviscid flow is obtained by a panel method. The boundary layer flow is computed by an interactive method that makes use of the Hilbert integral to couple the solutions of the inviscid and viscous flow equations. The second method is based on the solution of the compressible Navier-Stokes equations. The solution of these equations is obtained with an approximately factorized numerical algorithm, and with single block or multiple grids which enable grid embedding to enhance the resolution at isolated flow regions. In addition, the attached flow region can be computed by the numerical solution of compressible boundary layer equations. Unsteady pressure distributions obtained with both methods are compared with available experimental data.

Ekaterinaris, J. A.↗
NASA NTRS1978

Thermal response behavior of boundary-layer flows

This paper presents an approximate, but general, analysis for the thermal response behavior of incompressible, constant property, laminar boundary layer flow over a smooth object of arbitrary shape. It encompasses the classic time-dependent Leveque problem as its special case. Comparisons with available data show that, for fluids with Prandtl number of the order of unity or larger, this analytical solution is able to provide reasonably accurate results for most engineering applications. Under certain restrictive conditions, it can also be used to predict the thermal response of a compressible boundary layer flow.

Lin, F. N.↗
NASA NTRS1976

Turbulence measurements in axisymmetric supersonic boundary layer flow in adverse pressure gradients

Measurements have been made of the mean-flow and turbulence properties in adiabatic turbulent boundary layer flows subjected to adverse pressure gradients. In the freestream region upstream of the adverse pressure gradient the Mach number was 3.86, the unit Reynolds number 5.3 million per foot. The boundary layer developed on the wall of an axisymmetric nozzle and straight test section. The pressure gradients at the test section wall were induced by contoured centerbodies mounted on the wind tunnel centerline. The flow under study simulated that which might be found in an axially symmetric engine inlet of a supersonic aircraft. The results obtained have shown good agreement to exist between the measured normalized turbulent velocity fluctuations and the results from other recent investigations of compressible boundary layers.

Gootzait, E.↗
NASA NTRS1990

Design optimization of natural laminar flow fuselages in compressible flow

An optimization method has been developed to design axisymmetric body shapes such as fuselages, nacelles, and external fuel tanks with increased transition Reynolds numbers in subsonic compressible flow. The new design method involves a constraint minimization procedure coupled with analysis of the inviscid and viscous flow regions, and linear stability analysis of the compressible boundary layer. Boundary-layer transition is predicted by a 'hybrid' transition criterion based on Granville's transition criterion and a criterion using linear stability theory coupled with the e(n)-method. A tiptank of a business-jet is used as an example to illustrate that the method can be utilized to design an axisymmetric body shape with extensive natural laminar flow. Boundary layer transition is predicted to occur at a transition Reynolds number of 6.04 x 10 to the 6th on the original tiptank. On the designed body shape, a transition Reynolds number of 7.22 x 10 to the 6th is predicted using compressible linear stability theory coupled with e(n)-method.

Dodbele, Simha S.↗
NASA NTRS1978

Heat Transfer and Fluid Mechanics Institute, Meeting, 26th, Washington State University, Pullman, Wash., June 26-28, 1978, Proceedings

Papers are presented on the heat transfer and fluid mechanics in energy systems including diesel combustion analysis using rapid sampling and the numerical simulation of thermal response of architectural structures. Experiments associated with hypersonic boundary layers, the prediction of nozzle-wall pressure distribution, and the effects of nozzle geometry on an axisymmetric thrust augmentor are outlined. Predictions are made of film thickness on a roll-metering device and the thermal characteristics of multiple-surface submerged condenser modules are described. Numerical studies are made of the interaction between natural and forced convection in a complex geometry and of pressure drop and holdup in two-phase stratified flow. A numerical solution is given for viscous flows at high Reynolds numbers and for compressible boundary layer flows near separation or reattachment. The structure and transport properties in turbulent shear flows are reviewed along with the spectra of Reynolds stresses in a developing pipe flow.

Crowe, C. T.↗
NASA NTRS1996

Nonlinear Excitation of Inviscid Stationary Vortex in a Boundary-Layer Flow

We examine the excitation of inviscid stationary crossflow instabilities near an isolated surface hump (or indentation) underneath a three-dimensional boundary layer. As the hump height (or indentation depth) is increased from zero, the receptivity process becomes nonlinear even before the stability characteristics of the boundary layer are modified to a significant extent. This behavior contrasts sharply with earlier findings on the excitation of the lower branch Tollmien-Schlichting modes and is attributed to the inviscid nature of the crossflow modes, which leads to a decoupling between the regions of receptivity and stability. As a result of this decoupling, similarity transformations exist that allow the nonlinear receptivity of a general three-dimensional boundary layer to be studied with a set of canonical solutions to the viscous sublayer equations. The parametric study suggests that the receptivity is likely to become nonlinear even before the hump height becomes large enough for flow reversal to occur in the canonical solution. We also find that the receptivity to surface humps increases more rapidly as the hump height increases than is predicted by linear theory. On the other hand, receptivity near surface indentations is generally smaller in comparison with the linear approximation. Extension of the work to crossflow receptivity in compressible boundary layers and to Gortler vortex excitation is also discussed.

Choudhari, Meelan↗
NASA NTRS1985

The distortion of a supersonic turbulent boundary layer by bulk compression and surface curvature

An experimental investigation is conducted of the effects of bulk compression and streamline curvature on a turbulent supersonic boundary layer whose incoming Mach number is 2.87. Two different, constant-radius curved wall models were used; the total turning angle in both cases was fixed at 8 deg. While the amplification of the normal stress was approximately the same in both models, the amplification of the shear stress was greater in the one with delta(o)/R value 0.1 than in the model with a value of 0.02. The flow over the latter model appears to be especially complex.

Smits, A. J.↗
NASA NTRS1982

Correlation of Preston-tube data with laminar skin friction (Log No. J12984)

Preston tube data within laminar boundary layers obtained on a sharp ten-degree cone in the NASA Ames eleven-foot transonic wind tunnel are correlated with the corresponding values of theoretical skin friction. Data were obtained over a Mach number range of 0.30 to 0.95 and unit Reynolds numbers of 9.84, 13.1, and 16.4 million per meter. The rms scatter of skin friction coefficient about the correlation is of the order of one percent, which is comparable to the reported accuracy for calibrations of Preston tubes in incompressible pipe flows. In contrast to previous works on Preston tube/skin friction correlations, which are based on the physical height of the probe's face, this satisfactory correlation for compressible boundary layer flows is achieved by accounting for the effects of a variable "effective" height of the probe. The coefficients, which appear in the correlation, are dependent on the particular tunnel environment. The general procedure can be used to define correlations for other wind tunnels.

Reed, T. D.↗