Experimental Investigation of Film Cooling With Ejection From a Row of Holes for the Application to Gas Turbine Blades

1975 ◽  
Vol 97 (1) ◽  
pp. 21-27 ◽  
Author(s):  
C. Liess
Keyword(s):  
Boundary Layer ◽  
Gas Turbine ◽  
Velocity Ratio ◽  
Turbine Blades ◽  
Main Flow ◽  
Adiabatic Wall ◽  
Lateral Direction ◽  
Gas Turbine Blades ◽  
Flow Boundary ◽  
Displacement Thickness

The adiabatic wall effectiveness and the heat transfer coefficient is determined experimentally on a flat plate downstream of a row of inclined circular ejection holes. The measuring technique provides local values in downstream direction and averaged values in lateral direction. The ejection geometry is kept constant, i.e., ejection angle β = 35 deg, spacing to diameter ratio of ejection holes s/d = 3. The range of flow parameters corresponds closely to the conditions encountered on gas turbine blades. The main flow Mach number varies from 0.3 to 0.9, the mass velocity ratio from 0.1 to 2.0. Two favorable pressure gradients in the main flow are applied and several ratios of main flow boundary layer displacement thickness to ejection hole diameter. The main flow boundary layer upstream of the ejection is laminar and turbulent.

Influence of Chord Length and Inlet Boundary Layer on the Secondary Losses of Turbine Blades

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
Helmut Sauer ◽  
Robin Schmidt ◽  
Konrad Vogeler
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Flow Direction ◽  
Velocity Ratio ◽  
Main Flow ◽  
Chord Length ◽  
Displacement Thickness ◽  
Wide Range ◽  
Main Flow Direction

In this paper, results concerning the influence of chord length and inlet boundary layer thickness on the endwall loss of a linear turbine cascade are discussed. The investigations were performed in a low speed cascade tunnel using the turbine profile T40. The turning of 90 deg and 70 deg, the velocity ratio in the cascade from 1.0 to 3.5 as well as the chord length of 100 mm, 200 mm, and 300 mm were specified. In a measurement distance of one chord behind the cascade in main flow direction, an approximate proportionality of endwall loss and chord was observed in a wide range of velocity ratios. At small measurement distances (e.g., s2/l=0.4), this proportionality does not exist. If a part of the flow path within the cascade is approximately incorporated, a proportionality to the chord at small measurement distances can be obtained, too. Then, the magnitude of the endwall loss mainly depends on the distance in main flow direction. At velocity ratios near 1.0, the influence of the chord decreases rapidly, while at a velocity ratio of 1.0, the endwall loss is independent of the chord. By varying the inlet boundary layer thickness, no correlation of displacement thickness and endwall loss was achieved. A calculation method according to the modified integral equation by van Driest delivers the wall shear stress. Its influence on the endwall loss was analyzed.

Laminar and Transitional Boundary Layer Structures in Accelerating Flow With Heat Transfer

1986 ◽  
Vol 108 (1) ◽  
pp. 116-123 ◽  
Author(s):  
K. Rued ◽  
S. Wittig
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Pressure Gradients ◽  
Transfer Coefficients ◽  
Gas Turbine Blades ◽  
Heat Transfer Coefficients ◽  
Layer Structures ◽  
Flow Heat

The accurate prediction of heat transfer coefficients on cooled gas turbine blades requires consideration of various influence parameters. The present study continues previous work with special efforts to determine the separate effects of each of several parameters important in turbine flow. Heat transfer and boundary layer measurements were performed along a cooled flat plate with various freestream turbulence levels (Tu = 1.6−11 percent), pressure gradients (k = 0−6 × 10−6), and cooling intensities (Tw/T∞ = 1.0−0.53). Whereas the majority of previously available results were obtained from adiabatic or only slightly heated surfaces, the present study is directed mainly toward application on highly cooled surfaces as found in gas turbine engines.

Calculation of Heat Transfer to Convection-Cooled Gas Turbine Blades

1985 ◽  
Vol 107 (3) ◽  
pp. 620-627 ◽  
Author(s):  
W. Rodi ◽  
G. Scheuerer
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Complex Nature ◽  
Surface Boundary ◽  
Transfer Coefficients ◽  
Free Stream Turbulence ◽  
Gas Turbine Blades ◽  
Heat Transfer Coefficients

A mathematical model is presented for calculating the external heat transfer coefficients around gas turbine blades. The model is based on a finite-difference procedure for solving the boundary-layer equations which describe the flow and temperature field around the blades. The effects of turbulence are simulated by a low-Reynolds number version of the k-ε turbulence model. This allows calculation of laminar and transitional zones and also the onset of transition. Applications of the calculation method are presented to turbine-blade situations which have recently been investigated experimentally. Predicted and measured heat transfer coefficients are compared and good agreement with the data is observed. This is true especially for the pressure-surface boundary layer which is of a rather complex nature because it remains in a transitional state over the full blade length. The influence of various flow phenomena like laminar-turbulent transition and of the boundary conditions (pressure gradient, free-stream turbulence) on the predicted heat transfer rates is discussed.

Optimization of Trailing Edge Ejection Mixing Losses: A Theoretical and Experimental Study

1999 ◽  
Vol 121 (1) ◽  
pp. 118-125 ◽  
Author(s):  
M. T. Schobeiri ◽  
K. Pappu
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Mass Flow ◽  
Velocity Ratio ◽  
Turbine Blades ◽  
Ejection Velocity ◽  
Trailing Edge ◽  
Gas Turbine Blades ◽  
Flow Deflection ◽  
Mixing Losses

The aerodynamic effects of trailing edge ejection on mixing losses downstream of cooled gas turbine blades were experimentally investigated and compared with an already existing one-dimensional theory by Schobeiri (1989). The significant parameters determining the mixing losses and, therefore, the efficiency of cooled blades, are the ejection velocity ratio, the cooling mass flow ratio, the temperature ratio, the slot thickness ratio, and the ejection flow angle. To cover a broad range of representative turbine blade geometry and flow deflections, a General Electric power generation gas turbine blade with a high flow deflection and a NASA-turbine blade with intermediate flow deflection and different thickness distributions were experimentally investigated and compared with the existing theory. Comprehensive experimental investigations show that for the ejection velocity ratio μ = 1, the trailing edge ejection reduces the mixing losses downstream of the cooled gas turbine blade to a minimum, which is in agreement with the theory. For the given cooling mass flow ratios that are dictated by the heat transfer requirements, optimum slot thickness to trailing edge thickness ratios are found, which correspond to the minimum mixing loss coefficients. The results allow the turbine aerodynamicist to minimize the mixing losses and to increase the efficiency of cooled gas turbine blades.

Optimum Trailing Edge Ejection for Cooled Gas Turbine Blades

1989 ◽  
Vol 111 (4) ◽  
pp. 510-514 ◽  
Author(s):  
T. Schobeiri
Keyword(s):  
Gas Turbine ◽  
Mass Flow ◽  
Velocity Ratio ◽  
Turbine Blades ◽  
Slot Width ◽  
Width Ratio ◽  
Ejection Velocity ◽  
Trailing Edge ◽  
Gas Turbine Blades ◽  
Mixing Losses

The effect of trailing edge ejection on the flow downstream of a cooled gas turbine blade is investigated. Parameters that affect the mixing losses and therefore the efficiency of cooled blades are the ejection velocity ratio, the cooling mass flow ratio, the slot-width ratio, and the ejection angle. For ejection velocity ratio μ = 1, the trailing edge ejection reduces the mixing losses downstream to the cooled blade. For given cooling mass flow ratios, optimum slot-width/trailing edge ratios are found, which correspond to the minimum mixing loss coefficients.

Influence of Chord Length and Inlet Boundary Layer on the Secondary Losses of Turbine Blades

2010 ◽  
Author(s):  
Helmut Sauer ◽  
Robin Schmidt ◽  
Konrad Vogeler
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Flow Direction ◽  
Velocity Ratio ◽  
Turbine Blades ◽  
Flow Path ◽  
Chord Length ◽  
Displacement Thickness ◽  
Wide Range

In the present paper results concerning the influence of chord length and inlet boundary layer thickness on the endwall losses are discussed. The investigations were performed in a low speed cascade tunnel using the turbine profile T40. The deflection of 90 and 70 deg, the velocity ratio in the cascade from 1.0 to 3.5 as well as the chord length of 100,200 and 300 mm were predetermined. In a measurement distance behind the cascade of s2/l = 1, an approximate proportionality of endwall losses and chord length was observed in a wide range of velocity ratios. At small measurement distances (e.g. s2/l = 0.4), this proportionality does not exist. If aside from the flow path behind the cascade the flow path in the cascade is approximately incorporated, a proportionality to the chord length at small measurement distances can be obtained, too. Then to a large extent, the magnitude of the endwall losses is dependent on the length in main flow direction. At velocity ratios near 1.0, the influence of the chord length decreases rapidly, while at a velocity ratio of 1.0, the endwall losses are independent of chord length. By varying the inlet boundary layer thickness no correlation of displacement thickness and endwall losses was achieved. With a calculation method according to the modified integral equation by van Driest, the velocity gradient on the wall, the wall shear stress and the local friction coefficient were determined and their influence on the endwall losses analyzed.

scholarly journals Optimization of Trailing Edge Ejection Mixing Losses: A Theoretical and Experimental Study

1997 ◽  
Author(s):  
K. R. Pappu ◽  
M. T. Schobeiri
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Mass Flow ◽  
Velocity Ratio ◽  
Turbine Blades ◽  
Ejection Velocity ◽  
Trailing Edge ◽  
Gas Turbine Blades ◽  
Flow Deflection ◽  
Mixing Losses

The aerodynamic effects of trailing edge ejection on mixing losses downstream of cooled gas turbine blades were experimentally investigated and compared with an already existing one-dimensional theory by Schobeiri (1989). The significant parameters determining the mixing losses and therefore the efficiency of cooled blades are the ejection velocity ratio, the cooling mass flow ratio, the temperature ratio, the slot thickness ratio and the ejection flow angle. To cover a broad range of representative turbine blade geometry and flow deflections, a General Electric power generation gas turbine blade with high flow deflection and a NASA-turbine blade with intermediate flow deflection and different thickness distributions were experimentally investigated and compared with the existing theory. Comprehensive experimental investigations show that for the ejection velocity ratio μ = 1, the trailing edge ejection reduces the mixing losses downstream of the cooled gas turbine blade to a minimum which is in agreement with the theory. For the given cooling mass flow ratios that are dictated by the heat transfer requirements, optimum slot thickness to trailing edge thickness ratios are found, which correspond to the minimum mixing loss coefficients. The results allow the turbine aerodynamicist to minimize the mixing losses and to increase the efficiency of cooled gas turbine blades.

UDIMET L-605

Alloy Digest ◽  
2004 ◽  
Vol 53 (12) ◽  
Keyword(s):  
Corrosion Resistance ◽  
High Temperature ◽  
Physical Properties ◽  
Tensile Properties ◽  
Gas Turbine ◽  
Oxidation Resistance ◽  
Turbine Blades ◽  
Heat Treating ◽  
Gas Turbine Blades ◽  
Temperature Performance

Abstract Udimet L-605 is a high-temperature aerospace alloy with excellent strength and oxidation resistance. It is used in applications such as gas turbine blades and combustion area parts. This datasheet provides information on composition, physical properties, and tensile properties as well as creep. It also includes information on high temperature performance and corrosion resistance as well as forming, heat treating, and joining. Filing Code: CO-109. Producer or source: Special Metals Corporation.

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