Film Cooling Performance of Tripod Antivortex Injection Holes Over the Pressure and Suction Surfaces of a Nozzle Guide Vane

2017 ◽  
Vol 9 (2) ◽  
Author(s):  
Sridharan Ramesh ◽  
Christopher LeBlanc ◽  
Diganta Narzary ◽  
Srinath Ekkad ◽  
Mary Anne Alvin
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Nozzle Guide

Film cooling performance of the antivortex (AV) hole has been well documented for a flat plate. The goal of this study is to evaluate the same over an airfoil at three different locations: leading edge suction and pressure surface and midchord suction surface. The airfoil is a scaled up first stage vane from GE E3 engine and is mounted on a low-speed linear cascade wind tunnel. Steady-state infrared (IR) technique was employed to measure the adiabatic film cooling effectiveness. The study has been divided into two parts: the initial part focuses on the performance of the antivortex tripod hole compared to the cylindrical (CY) hole on the leading edge. Effects of blowing ratio (BR) and density ratio (DR) on the performance of cooling holes are studied here. Results show that the tripod hole clearly provides higher film cooling effectiveness than the baseline cylindrical hole case with overall reduced coolant usage on the both pressure and suction sides of the airfoil. The second part of the study focuses on evaluating the performance on the midchord suction surface. While the hole designs studied in the first part were retained as baseline cases, two additional geometries were also tested. These include cylindrical and tripod holes with shaped (SH) exits. Film cooling effectiveness was found at four different blowing ratios. Results show that the tripod holes with and without shaped exits provide much higher film effectiveness than cylindrical and slightly higher effectiveness than shaped exit holes using 50% lesser cooling air while operating at the same blowing ratios. Effectiveness values up to 0.2–0.25 are seen 40-hole diameters downstream for the tripod hole configurations, thus providing cooling in the important trailing edge portion of the airfoil.

Enhancement of Film Cooling Performance at the Leading Edge of Turbine Blade

2006 ◽  
Author(s):  
K.-S. Kim ◽  
Youn J. Kim ◽  
S.-M. Kim
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Cylindrical Body ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

To enhance the film cooling performance in the vicinity of the turbine blade leading edge, the flow characteristics of the film-cooled turbine blade have been investigated using a cylindrical body model. The inclination of the cooling holes is along the radius of the cylindrical wall and 20 deg relative to the spanwise direction. Mainstream Reynolds number based on the cylinder diameter was 1.01×105 and 0.69×105, and the mainstream turbulence intensities were about 0.2% in both Reynolds numbers. CO2 was used as coolant to simulate the effect of density ratio of coolant-to-mainstream. Furthermore, the effect of coolant flow rates was studied for various blowing ratios of 0.4, 0.7, 1.1, and 1.4, respectively. In experiment, spatially-resolved temperature distributions along the cylindrical body surface were visualized using infrared thermography (IRT) in conjunction with thermocouples, digital image processing, and in situ calibration procedures. This comparison shows the results generated to be reasonable and physically meaningful. The film cooling effectiveness of current measurement (0.29 mm × 0.33 min per pixel) presents high spatial and temperature resolutions compared to other studies. Results show that the blowing ratio has a strong effect on film cooling effectiveness and the coolant trajectory is sensitive to the blowing ratio. The local spanwise-averaged effectiveness can be improved by locating the first-row holes near the second-row holes.

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.

Comparison of Film Cooling Performance for Different Purge Slot Configurations in a Cylindrical and State-of-the-Art Nozzle Guide Vane

2021 ◽  
Author(s):  
Christian Landfester ◽  
Gunther Müller ◽  
Robert Krewinkel ◽  
Clemens Domnick ◽  
Martin Böhle
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
State Of The Art ◽  
Guide Vane ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract This comparative study is concerned with the advances in nozzle guide vane (NGV) design developments and their influence on the film cooling performance by injecting coolant through the purge slot. An experimental study compares the film cooling effectiveness as well as the aerodynamic effects for different purge slot configurations on both a flat and an axisymmetrically contoured endwall of a NGV. While the flat endwall cascade was equipped with four cylindrical vanes, the contoured endwall cascade consisted of four modern NGVs which represent state-of-the-art high-pressure turbine design standards. Geometric variations, e.g. the purge slot width and injection angle, as well as different blowing ratios (BR) at an engine-like density ratio (DR = 1.6) were realized to investigate the real-life effect of thermal expansion, design modifications and the interaction between secondary flow and coolant. The mainstream flow parameters were set to meet real engine conditions with regard to Reynolds and Mach numbers. The Pressure Sensitive Paint (PSP) technique was used to determine the adiabatic film cooling effectiveness. Five-hole probe measurements (DR = 1.0) were performed to measure the flow field with its characteristic vortex structures as well as the loss distribution in the vane wake region. For a more profound insight into the origin and development of the secondary flows, oil dye visualizations were carried out on both endwalls. The measurement results will be discussed based on a side-by-side comparison of the distribution of film cooling effectiveness on the endwall, its area-averaged values as well as the two-dimensional distribution of total pressure losses and the secondary flow field. The results of this study show that the advances in NGV design development have had a significantly positive influence on the distribution of the coolant. This has to be attributed to lesser disturbance of the coolant propagation by secondary flow for the optimized NGV design, since the design features are intended to suppress the formation of secondary flow. In contrast to the results of the cylindrical profile, sufficient cooling can be already provided with a perpendicular injection in the case of the modern NGV. It is therefore advisable to take these effects into account when designing the film cooling system of a modern high-pressure turbine.

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 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.

Effects of Endwall 3D Contouring on Film Cooling Effectiveness of Cylindrical Hole Injections at Different Locations on Vane Endwall

2018 ◽  
Author(s):  
Pingting Chen ◽  
Hongyu Gao ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Secondary Flows ◽  
Cylindrical Hole ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Endwall Contouring

With the development of gas turbine, the secondary flow loss in vane passage is getting higher. To reduce the strength of secondary flows within vane passage, endwall 3D contouring is an effective design. Endwall 3D contouring can lead to significant changes in the secondary flow vortices, which lead to changes on jet-to-secondary flow interaction and then changes on the film cooling effectiveness. Meanwhile, the geometry configuration of the contoured endwall, such as the rising and falling on the endwall, can also have an impact on film cooling performance. As a result, the film cooling performance on contoured endwall differs from that on flat endwall. Understanding the difference in film cooling characteristics on the contoured endwall and flat endwall may help to make better endwall contouring design and better endwall film cooling arrangement. The present experiment compares the film cooling effectiveness of cylindrical hole injections at different locations on 3D contoured endwall versus flat endwall in an NGV (nozzle guide vane) passage. The measurement is performed in a low speed wind tunnel with a F-class annular sector NGV cascade. The cylindrical hole injections are located as 4 different rows at −30% axial chord, 30% axial chord, 50% axial chord and 70% axial chord. Endwall pressure distribution is measured with pressure taps by pressure sensor while film cooling effectiveness is measured using PSP (Pressure Sensitive Paint). Two density ratios with 1.0 and 1.5 and several average blowing ratios are investigated. Effects of endwall contouring, density ratio and blowing ratio on film cooling effectiveness are obtained and the results are presented and explained in this investigation.

COMPARISON OF FILM COOLING PERFORMANCE FOR DIFFERENT PURGE SLOT CONFIGURATIONS IN A CYLINDRICAL AND STATE-OF-THE-ART NOZZLE GUIDE VANE

2021 ◽  
pp. 1-13
Author(s):  
Christian Landfester ◽  
Gunther Mueller ◽  
Robert Krewinkel ◽  
Clemens Domnick ◽  
Martin Böhle
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
State Of The Art ◽  
Guide Vane ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract This comparative study is concerned with the advances in nozzle guide vane (NGV) design developments and their influence on endwall film cooling performance by injecting coolant through the purge slot. This experimental study compares the film cooling effectiveness and the aerodynamic effects for different purge slot configurations on both a flat and an axisymmetrically contoured endwall of a NGV. While the flat endwall cascade was equipped with cylindrical vanes, the contoured endwall cascade consisted of modern NGVs which represent state-of-the-art high-pressure turbine design standards. Geometric variations, e.g. the slot width and injection angle, as well as different blowing ratios were realized. The mainstream flow parameters were set to meet real engine conditions with regard to Reynolds and Mach numbers. Pressure Sensitive Paint was used to determine the adiabatic film cooling effectiveness. Five-hole probe measurements were performed to measure the flow field in the vane wake region. For a more profound insight into the origin of the secondary flows, oil dye visualizations were carried out. The results show that the advances in NGV design have a significantly positive influence on the distribution of the coolant. This has to be attributed to lesser disturbance of the coolant propagation by secondary flow for the optimized NGV design, since the design features are intended to suppress the formation of secondary flow. It is therefore advisable to take these effects into account when designing the film cooling system of a modern high-pressure turbine.

Experimental Study on Film Cooling Effectiveness of Blade With Chevron Shaped Holes Under Wake Influence

2020 ◽  
Author(s):  
Ji-Chen Li ◽  
Hui-Ren Zhu ◽  
Da-Wei Chen ◽  
Dao-En Zhou
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.

Transonic Turbine Vane Suction Side Film Cooling With Showerhead Effect Using PSP Measurement Technique

2017 ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Nafiz H. K. Chowdhury ◽  
Je-Chin Han ◽  
Alexander V. Mirzamoghadam ◽  
Ardeshir Riahi
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Exit Mach Number ◽  
Density Ratios

This work focuses on the parametric experimental study of film cooling effectiveness on the suction side of a scaled turbine vane under transonic flow condition. The experiments were performed in a five-vane annular sector cascade blowdown facility. The controlled exit Mach numbers were 0.7, 0.9, and 1.1, from high subsonic to transonic conditions. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of coolant-to-mainstream density ratios, ranging from 1.0, 1.5 to 2.0. Three row-averaged coolant-to-mainstream blowing ratios in the range 0.7, 1.0, and 1.6 are studied. The test vane includes three rows of radial-angle cylindrical holes around the leading edge and two rows of compound-angle shaped holes on the suction side. All the cooling holes are active in order to study the resultant film cooling on suction side as well as from leading edge. Pressure sensitive paint (PSP) technique was used to obtain the film cooling effectiveness distributions from suction side holes and the contribution from leading edge showerhead holes. This work shows the effects of coolant-to-mainstream blowing ratio, density ratio, and exit Mach number on the film cooling effectiveness as well as its interaction with a potential shock wave. The results indicate that when the cooling holes are located in a critical region on the vane suction surface, the parametric effect on film cooling performance will significantly deviate from the common trend for a typical hole geometry.

Comparison of RANS and DES Modeling Against Measurements of Leading Edge Film Cooling on a First-Stage Vane

2016 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Spanwise Direction ◽  
Intensity Level ◽  
Edge Region ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness

The performance of a showerhead arrangement of film cooling in the leading edge region of a first stage nozzle guide vane was experimentally and numerically evaluated. A six-vane linear cascade was tested at an isentropic exit Mach number of Ma2s = 0.42, with a high inlet turbulence intensity level of 9%. The showerhead cooling scheme consists of four staggered rows of cylindrical holes evenly distributed around the stagnation line, angled at 45° towards the tip. The blowing ratios tested are BR = 2.0, 3.0 and 4.0. Adiabatic film cooling effectiveness distributions on the vane surface around the leading edge region were measured by means of Thermochromic Liquid Crystals technique. Since the experimental contours of adiabatic effectiveness showed that there is no periodicity across the span, the CFD calculations were conducted by simulating the whole vane. Within the RANS framework, the very widely used Realizable k-ε (Rke) and the Shear Stress Transport k-ω (SST) turbulence models were chosen for simulating the effect of the BR on the surface distribution of adiabatic effectiveness. The turbulence model which provided the most accurate steady prediction, i.e. Rke, was selected for running Detached Eddy Simulation at the intermediate value of BR = 3. Fluctuations of the local temperature were computed by DES, due to the vortex structures within the shear layers between the main flow and the coolant jets. Moreover, mixing was enhanced both in the wall-normal and spanwise direction, compared to RANS modeling. DES roughly halved the prediction error of laterally averaged film cooling effectiveness on the suction side of the leading edge. However, neither DES nor RANS provided the expected decay of effectiveness progressing downstream along the pressure side, with 15% overestimation of ηav at s/C =0.2.

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