Effect of Incidence on Wall Heating Rates and Aerodynamics on a Film-Cooled Transonic Turbine Blade

1991 ◽  
Vol 113 (3) ◽  
pp. 493-501 ◽  
Author(s):  
C. Camci ◽  
T. Arts
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Convective Heat Transfer ◽  
Stagnation Point ◽  
Short Duration ◽  
Convective Heat ◽  
Film Cooling ◽  
Free Stream ◽  
Suction Side ◽  
Pressure Side

This study investigates the influence of incidence on convective heat transfer to highly curved surfaces of a film-cooled turbine rotor blade. A computational study of free-stream inviscid aerodynamics without cooling at various incidences is followed by well-documented measured heat transfer data sets. The heat transfer experiments are discussed for cases with and without film cooling, performed under realistic gas turbine flow conditions in the short-duration heat transfer facility of the von Karman Institute for Fluid Dynamics. The precise location of the stagnation point and the iso-Mach number contours in the passage for each incidence (−10, 0, 10, +15 deg) are presented for a nominal exit Mach number of 0.94. The free-stream mass flow rate was kept constant for each experiment at different incidence levels. Three rows of compound angled discrete cooling holes are located near the leading edge in a showerhead configuration. Two rows of staggered discrete cooling holes are located on the suction side and a single row of cooling holes is located on the pressure side. The short-duration measurements of quantitative wall heat fluxes on nearly isothermal blade surfaces both in the presence and absence of coolant ejection are presented. The study indicated that the change of the position of the stagnation point strongly altered the aerodynamic behavior and convective heat transfer to the blade in approximately the first 30 percent of both the pressure side and the suction side in the presence and absence of film cooling. The immediate vicinity of the stagnation point was not significantly affected by changing incidence without cooling. Transitional behavior both on the suction surface and on the pressure surface was significantly influenced by the changes in approaching flow direction. Flow separation associated with incidence variations was also observed. Extremely low levels of the convective heat transfer coefficients were experienced near the regions where small separation bubbles are located.

Effect of Incidence on Wall Heating Rates and Aerodynamics on a Film Cooled Transonic Turbine Blade

1990 ◽  
Author(s):  
Cengiz Camci ◽  
Tony Arts
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Convective Heat Transfer ◽  
Stagnation Point ◽  
Short Duration ◽  
Convective Heat ◽  
Film Cooling ◽  
Free Stream ◽  
Suction Side ◽  
Pressure Side

This study investigates the influence of incidence on convective heat transfer to highly curved surfaces of a film cooled turbine rotor blade. A computational study of free stream inviscid aerodynamics without cooling at various incidences is followed by well documented measured heat transfer data sets. The heat transfer experiments are discussed for cases with and without film cooling, performed under realistic gas turbine flow conditions in the short duration heat transfer facility of the von Karman Institute for Fluid Dynamics. The precise location of the stagnation point and the iso-Mach number contours in the passage for each incidence (−10°, 0°, 10°, +10°) are presented for a nominal exit Mach number of 0.94. The free stream mass flow rate was kept constant for each experiment at different incidence levels. Three rows of compound angled discrete cooling holes are located near the leading edge in a shower-head configuration. Two rows of staggered discrete cooling holes are located on the suction side and a single row of cooling holes is located on the pressure side. The short duration measurements of quantitative wall heat fluxes on nearly isothermal blade surfaces both in the presence and absence of coolant ejection are presented. The study indicated that the change of the position of the stagnation point strongly altered the aerodynamic behaviour and convective heat transfer to the blade in approximately the first 30 % of both the pressure side and the suction side in the presence and absence of film cooling. The immediate vicinity of the stagnation point was not significantly affected by changing incidence without cooling. Transitional behaviour both on the suction surface and on the pressure surface was significantly influenced by the changes in approching flow direction. Flow separation associated with incidence variations was also observed. Extremely low levels of convective heat transfer coefficients were experienced near the regions where small separation bubbles are located.

An Experimental Convective Heat Transfer Investigation Around a Film-Cooled Gas Turbine Blade

1990 ◽  
Vol 112 (3) ◽  
pp. 497-503 ◽  
Author(s):  
C. Camci ◽  
T. Arts
Keyword(s):  
Heat Transfer ◽  
Convective Heat Transfer ◽  
Gas Turbine ◽  
Convective Heat ◽  
Free Stream ◽  
Weight Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Turbine Rotor ◽  
Free Stream Turbulence

The present paper deals with an experimental convective heat transfer investigation around a film-cooled, high-pressure gas turbine rotor blade mounted in a stationary, linear cascade arrangement. The measurements were performed in the von Karman Institute Isentropic Light Piston Compression Tube facility. The test blade was made of Macor glass ceramic and was instrumented with thin film gages. The coolant flow was ejected simultaneously through the leading edge (three rows of holes), the suction side (two rows of holes), and the pressure side (one row of holes). The effects of overall mass weight ratio, coolant to free-stream temperature ratio, and free-stream turbulence were successively investigated.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Thermal conductivity of argon at high temperatures

1965 ◽  
Vol 21 (4) ◽  
pp. 673-688 ◽  
Author(s):  
Morton Camac ◽  
Robert M. Feinberg
Keyword(s):  
Heat Transfer ◽  
Thermal Conductivity ◽  
Convective Heat Transfer ◽  
Convective Heat ◽  
Free Stream ◽  
Incident Shock ◽  
Radiation Emission ◽  
Reflected Shock ◽  
Theoretical Predictions ◽  
Experimental Values

An infra-red heat-transfer gauge was used in a shock tube for end-wall measurements of the convective heat transfer from argon behind the reflected shock. The thermal conductivity of neutral (un-ionized) argon was measured before the ionization-relaxation time, and was fitted with the power-law temperature dependence 4·2 × 10−5(T/300)0·76±0·03cal/sec cm°K, whereTis measured in °K, and ±0·03 refers to the probable error The free-stream temperature ranged from 20,000 to 75,000°K, corresponding to incident-shock velocities from 3 to 6mm/μsec. At later times, after the free stream established equilibrium ionization, the convective-heat-transfer rate remained the same as the initial rate with neutral argon. Theoretical predictions of Fay & Kemp (1965), assuming equilibrium-boundary-layer conditions, are 20–30% below the experimental values. Also reported in this paper are measurements of the ionization times behind the reflected shock, and these are in agreement with an extrapolation of the Petschek & Byron (1957) measurements behind the incident shock. There is a discussion of the large changes in the gas conditions behind the reflected shock due to the ionization process. The final equilibrium conditions are reached abruptly, as indicated by the continuum-radiation emission which becomes constant immediately after ionization relaxation.

Film Cooling Jet Injection Effect in Heat Transfer Coefficient Augmentation for the Pressure Side Cooling of Turbine Vane

2014 ◽  
Author(s):  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Mats Kinell ◽  
Esa Utriainen
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Time Resolved ◽  
Turbine Vane

Improving film cooling performance of turbine vanes and blades is often achieved through application of multiple arrays of cooling holes on the suction side, the showerhead region and the pressure side. This study investigates the pressure side cooling under the influence of single and multiple rows of cooling in the presence of a showerhead from a heat transfer coefficient augmentation perspective. Experiments are conducted on a prototype turbine vane working at engine representative conditions. Transient IR thermography is used to measure time-resolved surface temperature and the semi-infinite method is utilized to calculate the heat transfer coefficient on a low conductive material. Investigations are performed for cylindrical and fan-shaped holes covering blowing ratio 0.6 and 1.8 at density ratio of about unity. The freestream turbulence is approximately 5% close to the leading edge. The resulting heat transfer coefficient enhancement, the ratio of HTC with to that without film cooling, from different case scenarios have been compared to showerhead cooling only. Findings of the study highlight the importance of showerhead cooling to be used with additional row of cooling on the pressure side in order to reduce heat transfer coefficient enhancement. In addition, it is shown that extra rows of cooling will not significantly influence heat transfer augmentation, regardless of the cooling hole shape.

Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

2005 ◽  
Author(s):  
W. Colban ◽  
A. Gratton ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Pressure Side ◽  
Guide Vanes ◽  
Single Row ◽  
Inlet Guide Vanes ◽  
Adiabatic Effectiveness ◽  
Lift Off

In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film-cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

2005 ◽  
Vol 128 (1) ◽  
pp. 53-61 ◽  
Author(s):  
W. Colban ◽  
A. Gratton ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Pressure Side ◽  
Guide Vanes ◽  
Single Row ◽  
Inlet Guide Vanes ◽  
Adiabatic Effectiveness ◽  
Lift Off

In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single-row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single-row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

Sign in / Sign up

Export Citation Format

Share Document