The Boundary Layer on a Townend Surface

1965 ◽  
Vol 16 (2) ◽  
pp. 145-158 ◽  
Author(s):  
J. C. Cooke ◽  
O. K. Jones
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Mach Number ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Typical Case ◽  
Air Stream ◽  
Straight Line ◽  
Displacement Thickness ◽  
Limiting Streamlines

SummaryCertain three-dimensional surfaces studied by Townend have the property that at every streamwise section a supersonic air stream is compressed isentropically in the form of a reversed Prandtl-Meyer flow. If all the compressions are centred on points lying on one straight line the flow is conical. Such a surface may be of use as an engine intake. In this paper an approximate boundary-layer calculation for such a surface is made for laminar flow with zero heat transfer in the conical case. It is found that under certain circumstances a useful amount of compression can be obtained before separation occurs (a pressure ratio of 6, for instance, for an initial Mach number of 10). Limiting streamlines are drawn and displacement thickness calculated for a typical case, and the Mach number at separation is determined for a range of initial Mach numbers and angles of sweep. Instability due to sweep is considered.

scholarly journals Application of Advanced Computational Codes in the Design of an Experiment for a Supersonic Throughflow Fan Rotor

1988 ◽  
Vol 110 (2) ◽  
pp. 270-279
Author(s):  
J. R. Wood ◽  
J. F. Schmidt ◽  
R. J. Steinke ◽  
R. V. Chima ◽  
W. G. Kunik
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Flow ◽  
Navier Stokes ◽  
Nasa Lewis Research ◽  
Two Dimensional ◽  
Final Design ◽  
Dimensional Boundary

Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The data base for components of this type is practically nonexistent; therefore, in order to furnish the required information for assessment of this type fan, a program has been initiated at the NASA Lewis Research Center to design, build, and test a fan rotor that operates with supersonic axial velocities from inlet to exit. This paper describes the design of the experiment using advanced computational codes to calculate the unique components required. The fan rotor has constant hub and tip radii and was designed for a pressure ratio of 2.7 with a tip speed of 457 m/s. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier–Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. A translating nozzle was designed to produce a uniform flow parallel to the fan up to the design axial Mach number of 2.0. The nozzle was designed with the three-dimensional Euler/interactive boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code, which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.

Design and Analysis of a High Pitch to Chord Ratio Cascade Representative of Ducted Propfans

1991 ◽  
Author(s):  
A. Weber ◽  
W. Steinert ◽  
H. Starken
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
High Pitch ◽  
Integral Boundary ◽  
Pressure Loss Coefficient ◽  
Distribution Boundary ◽  
Very High

Efforts to reduce the specific fuel consumption of a modern aero engine focus in particular on increasing the by-pass ratio beyond the current level of around 5. One concept is the counterrotating shrouded propfan operating at low overall pressure ratio and having only very few fan blades of extremely high pitch/chord ratios. The relative inlet Mach numbers cover a range from 0.7 at the hub to 1.1 at the tip section of the first rotor. A propfan cascade was designed by taking into account two characteristic features of a propfan blade-blade section: • a very high pitch/chord ratio of s/c = 2.25 • an inlet Mach number of M1 = 0.90 which leads to transonic flow conditions inside the blade passage In the design process a profile generator and a quasi-3D Euler solver were used iteratively to optimize the profile Mach number distribution. Boundary layer behavior was checked with an integral boundary layer code. The cascade design was verified experimentally in the transonic cascade wind tunnel of DLR at Cologne. The extensive experimental results confirm the design goal of roughly 5 degree flow turning. A total pressure loss coefficient of less than 1.5% was measured at design conditions. This validates the very high efficiency level the propfan concept is calling for. A 2D Navier-Stokes flow analysis code yields good results in comparison to the experimental ones.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

scholarly journals Experimental study of the impact of N-wave on heat transfer in a boundary layer of a flat plate at the Mach number 2

2021 ◽  
Author(s):  
V. L. Kocharin ◽  
A. A. Yatskikh ◽  
D. S. Prishchepova ◽  
A. V. Panina ◽  
Yu. G. Yermolaev ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Experimental Study ◽  
Mach Number ◽  
Flat Plate ◽  
N Wave ◽  
The Impact

Turbine Airfoil Heat Transfer Predictions Using CFD

2005 ◽  
Author(s):  
Anil K. Tolpadi ◽  
James A. Tallman ◽  
Lamyaa El-Gabry
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Turbulence Modeling ◽  
Three Dimensional ◽  
Numerical Approach ◽  
Navier Stokes ◽  
Design System ◽  
Computational Fluid Dynamics Cfd ◽  
Turbine Airfoils ◽  
Typical Design

Conventional heat transfer design methods for turbine airfoils use 2-D boundary layer codes (BLC) combined with empiricism. While such methods may be applicable in the mid span of an airfoil, they would not be very accurate near the end-walls and airfoil tip where the flow is very three-dimensional (3-D) and complex. In order to obtain accurate heat transfer predictions along the entire span of a turbine airfoil, 3-D computational fluid dynamics (CFD) must be used. This paper describes the development of a CFD based design system to make heat transfer predictions. A 3-D, compressible, Reynolds-averaged Navier-Stokes CFD solver with k-ω turbulence modeling was used. A wall integration approach was used for boundary layer prediction. First, the numerical approach was validated against a series of fundamental airfoil cases with available data. The comparisons were very favorable. Subsequently, it was applied to a real engine airfoil at typical design conditions. A discussion of the features of the airfoil heat transfer distribution is included.

The Effect of a Wall Boundary Layer on Local Mass Transfer From a Cylinder in Crossflow

1984 ◽  
Vol 106 (2) ◽  
pp. 260-267 ◽  
Author(s):  
R. J. Goldstein ◽  
J. Karni
Keyword(s):  
Boundary Layer ◽  
Mass Transfer ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Two Dimensional ◽  
Tunnel Wall ◽  
Wall Boundary Layer ◽  
Naphthalene Sublimation Technique ◽  
Displacement Thickness ◽  
Two Dimensional Flow

A naphthalene sublimation technique is used to determine the circumferential and longitudinal variations of mass transfer from a smooth circular cylinder in a crossflow of air. The effect of the three-dimensional secondary flows near the wall-attached ends of a cylinder is discussed. For a cylinder Reynolds number of 19000, local enhancement of the mass transfer over values in the center of the tunnel are observed up to a distance of 3.5 cylinder diameters from the tunnel wall. In a narrow span extending from the tunnel wall to about 0.066 cylinder diameters above it (about 0.75 of the mainstream boundary layer displacement thickness), increases of 90 to 700 percent over the two-dimensional flow mass transfer are measured on the front portion of the cylinder. Farther from the wall, local increases of up to 38 percent over the two-dimensional values are measured. In this region, increases of mass transfer in the rear portion of the cylinder, downstream of separation, are, in general, larger and cover a greater span than the increases in the front portion of the cylinder.

Temperature Control of Blow-Down, Supersonic Tunnels

1956 ◽  
Vol 60 (541) ◽  
pp. 67-70
Author(s):  
T. A. Thomson
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Reynolds Numbers ◽  
Stagnation Temperature ◽  
Blow Down ◽  
High Reynolds Numbers ◽  
Experimental Errors ◽  
High Reynolds

The blow-down type of intermittent, supersonic tunnel is attractive because of its simplicity and because relatively high Reynolds numbers can be obtained for a given size of test section. An adverse characteristic, however, is the fall of stagnation temperature during runs, which can affect experiments in several ways. The Reynolds number varies and the absolute velocity is not constant, even if the Mach number and pressure are; heat-transfer cannot be studied under controlled conditions and the experimental errors arising from the effect of heat-transfer on the boundary layer vary in time. These effects can become significant in quantitative experiments if the tunnel is large and the variation of temperature very rapid; the expense required to eliminate them might then be justified.

Darryl E. Metzger Memorial Session Paper: Comparison of Calculated and Measured Heat Transfer Coefficients for Transonic and Supersonic Boundary-Layer Flows

1995 ◽  
Vol 117 (2) ◽  
pp. 248-254 ◽  
Author(s):  
C. Hu¨rst ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Reynolds Number ◽  
High Speed ◽  
Three Dimensional ◽  
Supersonic Boundary Layer ◽  
Nozzle Wall ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Three Dimensional Finite Element

The present study compares measured and computed heat transfer coefficients for high-speed boundary layer nozzle flows under engine Reynolds number conditions (U∞=230 ÷ 880 m/s, Re* = 0.37 ÷ 1.07 × 106). Experimental data have been obtained by heat transfer measurements in a two-dimensional, nonsymmetric, convergent–divergent nozzle. The nozzle wall is convectively cooled using water passages. The coolant heat transfer data and nozzle surface temperatures are used as boundary conditions for a three-dimensional finite-element code, which is employed to calculate the temperature distribution inside the nozzle wall. Heat transfer coefficients along the hot gas nozzle wall are derived from the temperature gradients normal to the surface. The results are compared with numerical heat transfer predictions using the low-Reynolds-number k–ε turbulence model by Lam and Bremhorst. Influence of compressibility in the transport equations for the turbulence properties is taken into account by using the local averaged density. The results confirm that this simplification leads to good results for transonic and low supersonic flows.

Sign in / Sign up

Export Citation Format

Share Document