Influence of Unsteady Wake With Trailing Edge Coolant Ejection on Turbine Blade Film Cooling

2011 ◽  
Author(s):  
Shiou-Jiuan Li ◽  
Akhilesh P. Rallabandi ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Surface Film ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Rotor Blade ◽  
Pressure Surface ◽  
Unsteady Wake

Detailed film cooling effectiveness distributions along a modeled turbine rotor blade under combined effects of upstream trailing edge unsteady wake with coolant ejection are presented using the pressure sensitive paint (PSP) mass transfer analogy method. The experiment is conducted in a low speed wind tunnel facility with a five blade linear cascade. The exit Reynolds number based on the axial chord is 370,000. Unsteady wakes and trailing edge coolant jets are produced by a spoked wheel-type wake generator with hollow rods equipped with several coolant ejections from holes. The coolant-to-mainstream density ratios for both blade and trailing edge coolant ejection range from 1.5 to 2.0 for simulating realistic engine conditions. Blade blowing ratios studied are 0.5 and 1.0 on Suction surface and 1.0 and 2.0 on Pressure surface. Trailing edge jet blowing ratio and Strouhal number are 1.0 and 0.12, respectively. Results show the unsteady wake reduces overall film cooling effectiveness. However, the unsteady wake with trailing edge coolant ejection enhances overall effectiveness. Results also show that the overall filming cooling effectiveness increases by using heavier coolant for trailing edge ejection as well as for blade surface film cooling.

Influence of Unsteady Wake With Trailing Edge Coolant Ejection on Turbine Blade Film Cooling

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Shiou-Jiuan Li ◽  
Akhilesh P. Rallabandi ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Surface Film ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Rotor Blade ◽  
Blowing Ratio ◽  
Unsteady Wake

Detailed film cooling effectiveness distributions along a modeled turbine rotor blade under the combined effects of an upstream trailing edge unsteady wake with coolant ejection are presented using the pressure sensitive paint (PSP) mass transfer analogy method. The experiment is conducted in a low speed wind tunnel facility with a five blade linear cascade. The exit Reynolds number based on the axial chord is 370,000. Unsteady wakes and trailing edge coolant jets are produced by a spoked wheel-type wake generator with hollow rods equipped with several coolant ejections from holes. The coolant-to-mainstream density ratios for both the blade and trailing edge coolant ejection range from 1.5 to 2.0 for simulating realistic engine conditions. Blade blowing ratio studies are 0.5 and 1.0 on the suction surface and 1.0 and 2.0 on the pressure surface. The trailing edge jet blowing ratio and Strouhal numbers are 1.0 and 0.12, respectively. The results show that the unsteady wake reduces the overall film cooling effectiveness. However, the unsteady wake with trailing edge coolant ejection enhances the overall effectiveness. The results also show that the overall filming cooling effectiveness increases by using heavier coolant for trailing edge ejection and for blade surface film cooling.

scholarly journals Film Cooling on a Gas Turbine Rotor Blade

1991 ◽  
Author(s):  
Kenichiro Takeishi ◽  
Sunao Aoki ◽  
Tomohiko Sato ◽  
Keizo Tsukagoshi
Keyword(s):  
Mass Transfer ◽  
Film Cooling ◽  
Rotor Blade ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Rotor Blade ◽  
Pressure Surface ◽  
Rotating Blade

The film cooling effectiveness on a low-speed stationary cascade and the rotating blade has been measured by using a heat-mass transfer analogy. The film cooling effectiveness on the suction surface of the rotating blade fits well with that on the stationary blade, but a low level of effectiveness appears on the pressure surface of the rotating blade. In this paper, typical film cooling data will be presented and film cooling on a rotating blade is discussed.

Film Cooling on a Gas Turbine Rotor Blade

1992 ◽  
Vol 114 (4) ◽  
pp. 828-834 ◽  
Author(s):  
K. Takeishi ◽  
S. Aoki ◽  
T. Sato ◽  
K. Tsukagoshi
Keyword(s):  
Mass Transfer ◽  
Film Cooling ◽  
Rotor Blade ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Rotor Blade ◽  
Pressure Surface ◽  
Rotating Blade

The film cooling effectiveness on a low-speed stationary cascade and the rotating blade has been measured by using a heat-mass transfer analogy. The film cooling effectiveness on the suction surface of the rotating blade fits well with that on the stationary blade, but a low level of effectiveness appears on the pressure surface of the rotating blade. In this paper, typical film cooling data will be presented and film cooling on a rotating blade is discussed.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using PSP Technique

2010 ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Wake ◽  
Lift Off ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke wheel wake generator) on the modeled rotor blade is studied using the Pressure Sensitive Paint (PSP) mass transfer analogy method. Emphasis of the current study is on the mid-span region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film cooling holes. The blade also has radial shower-head leading edge film cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds Number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side; 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using Pressure Sensitive Paint Technique

2012 ◽  
Vol 134 (8) ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Suction Side ◽  
Pressure Sensitive Paint ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Unsteady Wake ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke-wheel wake generator) on the modeled rotor blade is studied using the pressure sensitive paint (PSP) mass-transfer analogy method. Emphasis of the current study is on the midspan region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film-cooling holes. The blade also has radial shower-head leading edge film-cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side and 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film-cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

Effects of Axial Row-Spacing for Double-Jet Film-Cooling on the Cooling Effectiveness

2011 ◽  
Author(s):  
Zhan Wang ◽  
Jian-Jun Liu ◽  
Bai-tao An ◽  
Chao Zhang
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Guide Vane ◽  
Row Spacing ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Cylindrical Holes ◽  
Compound Angle

The effects of axial row-spacing for double jet film-cooling (DJFC) with compound angle on the cooling characteristics under different blowing ratios were investigated numerically. First, the flow fields and cooling effectiveness of DJFC on flat plate with different axial row-spacing were calculated. Film-cooling with fan-shaped or cylindrical holes was also calculated for the comparison. The results indicate that a larger axial row-spacing is helpful to form the anti-kidney vortex and to improve the cooling effectiveness. The DJFC was then applied to the suction and pressure surface of a real turbine inlet guide vane. Comparisons of film-cooling effectiveness with the cylindrical and fan-shaped holes were also conducted. The results for the guide vane show that on the suction surface the DJFC with a larger axial row-spacing leads to better film coverage and better film-cooling effectiveness than the cylindrical or fan-shaped holes. On the pressure surface, however, the film-cooling with fan-shaped holes is superior to the others.

Experimental and numerical investigation of trailing edge film cooling downstream of a slot with internal rib arrays

2003 ◽  
Vol 217 (4) ◽  
pp. 393-401 ◽  
Author(s):  
P Martini ◽  
A Schulz ◽  
C. F. Whitney ◽  
E Lutum
Keyword(s):  
Film Cooling ◽  
Surface Film ◽  
Numerical Data ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Efficiency ◽  
Computational Fluid Dynamics Cfd ◽  
Aero Engines ◽  
The University

To enhance turbine efficiency, the trailing edge thickness of stators and blades needs to be as thin as possible. One limitation on the trailing edge thickness is the requirement to cool this life-limiting region of the blade. One technique used to achieve a thin cooled trailing edge is that of a pressure side cutback with film cooling slots. There is a paucity of fluid and heat transfer data regarding this type of geometry which is currently being addressed by the EC-funded Framework V project AITEB. This paper reports on experimental work being undertaken by the University of Karlsruhe and accompanying computational fluid dynamics (CFD) calculations being performed by MTU Aero Engines and ALSTOM Power. Experimental and numerical data presented include cutback surface film cooling effectiveness together with slot discharge coefficient values.

Experimental Study on Film Cooling Effectiveness of Blade with Chevron Shaped Holes under Wake Influence

2021 ◽  
pp. 1-20
Author(s):  
Jichen Li ◽  
Hui Ren Zhu ◽  
Cun Liang Liu ◽  
Lin Ye ◽  
Zhou Daoen
Keyword(s):  
Mass Flux ◽  
Film Cooling ◽  
Gas Turbines ◽  
Flux Ratio ◽  
Leading Edge ◽  
High Mass ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface

Abstract Gas turbines have been widely used. With the continuous improvement of the performance of gas turbines, the turbine inlet temperature has greatly exceeded the heat resistance limit of the turbine blade material, so advanced cooling technology is required. The film cooling effectiveness distribution over the blade under the effect of wake was obtained by Pressure Sensitive Paint (PSP) technique. The test blade has 5 rows of chevron film holes on the pressure side, 3 rows of cylindrical film holes on the leading edge and 3 rows of chevron film holes on the suction side. The mainstream Reynolds number is 130,000 based on the blade chord length, and the mainstream turbulence intensity is 2.7%. The upstream wake was simulated by the spoken-wheel type wake generator. The film cooling effectiveness was measured at three wake Strouhal numbers (0, 0.12 and 0.36) and three mass flux ratios (MFR1, MFR2 and MFR3). The results show that the increase of mass flux ratio leads a decrease of the film cooling effectiveness on the suction surface. In the wake condition, the effect of mass flux ratio is weakened. Wake leads a marked decrease of the film cooling effectiveness over most blade surface except for the surface near leading edge on the pressure surface. In the high mass flux ratio condition, the effect of wake on the film cooling effectiveness is weakened on the suction surface and strengthened on the pressure surface.

Evaluation of RANS Predictions on a Linear Nozzle Vane Cascade With Trailing Edge Cutback Film Cooling

2013 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane ◽  
Pressure Surface ◽  
Exit Mach Number ◽  
Cooling Air

The present study concentrates on the numerical investigation of a cooled trailing edge in a linear nozzle vane cascade typical of a high-pressure turbine. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is also ejected through two rows of cooling holes placed on the pressure side, upstream of the cutback. The main goal is to evaluate the reliability of RANS predictions in such a complex cooling system. Different coolant-to-mainstream mass flow ratio values up to MFR = 2.8% were simulated at exit Mach number of M2is = 0.2. The computed performance of the trailing edge cooling scheme was compared to available measurements of: holes and cutback exit velocity and discharge behavior; boundary layer along traverses located on the pressure side, downstream of each row of cooling holes and approaching the trailing edge; adiabatic film cooling effectiveness. Special emphasis was dedicated to coolant-mainstream interaction and film cooling effectiveness over the pressure surface of the vane. Despite the steady approach, the simulations provided a reliable overview of coolant and mainstream aerodynamic features. The limitations in predicting the measured drop in cooling effectiveness toward the trailing edge were highlighted as well.

Experimental Study of Full Coverage Film Cooling Effectiveness for a Turbine Blade With Compound Shaped Holes

2020 ◽  
Author(s):  
Shuai-qi Zhang ◽  
Cun-liang Liu ◽  
Qi-ling Guo ◽  
Da-peng Liang ◽  
Fan Zhang
Keyword(s):  
Turbine Blade ◽  
Turbulence Intensity ◽  
Mass Flux ◽  
Film Cooling ◽  
Flux Ratio ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Cylindrical Holes

Abstract The film coverage of a turbine blade surface is determined by all the film cooling structures. The direct study of full coverage film cooling is relatively rare, especially for related research on turbine blades. In this paper, the pressure-sensitive paint (PSP) measurement technique is used to carry out experiments under different turbulence intensities and mass flux ratios, and the distribution of the film cooling effectiveness on the entire surface is studied in detail. In this study, a basic turbine blade and an improved turbine blade are investigated. The film cooling hole position distribution on the improved blade is the same as that on the basic blade, but the film cooling hole shape on the suction surface and the pressure surface is changed from cylindrical holes to laid-back fan-shaped holes. Both blades have 5 rows of cylindrical holes at the leading edge and 4 rows of film cooling holes on the suction surface and the pressure surface. The leading edge, suction surface, and pressure surface have their own coolant inlet cavities. This kind of design is not only close to the actual working conditions in a flow distribution but also conveniently eliminates the mutual interference caused by the uneven flow distribution between the pressure surface and the suction surface to facilitate the independent analysis of the pressure surface and the suction surface. In this paper, the film cooling effectiveness of two kinds of turbine blades under different turbulence intensities and mass flux ratios is studied. The results show that the average cooling effectiveness of the improved blade is much better than that of the basic blade. The laid-back fan-shaped hole rows improve the cooling effectiveness of the suction surface by 60% to 100% and 50% to 120% on the pressure surface. The increase in turbulence intensity will reduce the cooling effectiveness of the blade surface; however, the effect of the turbulence intensity becomes weaker with an increase in the mass flux ratio. Compared with the multiple rows of cylindrical holes, the cooling effectiveness of the laid-back fan-shaped holes is more affected by the turbulence intensity under the small mass flux ratio.

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