Aerodynamic Loss for a Turbine Blade With Endwall Leakage Features and Contouring

2013 ◽  
Author(s):  
Stephen P. Lynch ◽  
Karen A. Thole ◽  
Atul Kohli ◽  
Christopher Lehane ◽  
Tom Praisner
Keyword(s):  
Turbine Blade ◽  
Aerodynamic Performance ◽  
Secondary Flows ◽  
Leakage Flow ◽  
Aerodynamic Loss ◽  
Endwall Contouring

Secondary flows near the endwall of a turbine blade contribute to a loss in aerodynamic performance in the engine. Further reductions in performance occur when the secondary flows interact with leakage flow from necessary clearance features, such as the clearance gap between the blade rotor and an upstream stator or gaps between adjacent blade platforms. Non-axisymmetric endwall contouring has been shown to reduce the strength of secondary flows near the endwall, but relatively little research has been done to test the sensitivity of the contouring to the endwall leakage features. This paper describes aerodynamic measurements taken downstream of a cascade with representative leakage features. In general, upstream leakage flow with swirl relative to the blade increased aerodynamic loss, relative to leakage that was matched to the blade wheelspeed. Non-axisymmetric contouring for an endwall without a platform gap reduced underturning of the flow but had no effect on overall loss, relative to a flat endwall without a gap. A contoured endwall with a platform gap had 12% higher mixed-out loss than a contoured endwall without a gap.

scholarly journals A Summary of the Cooled Turbine Blade Tip Heat Transfer and Film Effectiveness Investigations Performed by Dr. D. E. Metzger

1994 ◽  
Author(s):  
Y. W. Kim ◽  
W. Abdel-Messeh ◽  
J. P. Downs ◽  
F. O. Soechting ◽  
G. D. Steuber ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Convection Heat Transfer ◽  
Secondary Flows ◽  
Test Surface ◽  
Leakage Flow ◽  
Test Technique ◽  
Coolant Injection ◽  
Blade Tip

The clearance gap between the stationary outer air seal and blade tips of an axial turbine allows a clearance gap leakage flow to be driven through the gap by the pressure-to-suction side pressure difference. The presence of strong secondary flows on the pressure side of the airfoil tends to deliver air from the hottest regions of the mainstream to the clearance gap. The blade tip region, particularly near the trailing edge, is very difficult to cool adequately with blade internal coolant flow. In this case, film cooling injection directly onto the blade tip region can be used in an attempt to directly reduce the heat transfer rates from the hot gases in the clearance gap to the blade tip. The present paper is intended as a memorial tribute to the late Professor Darryl E. Metzger who has made significant contributions in this particular area over the past decade. A summary of this work is made to present the results of his more recent experimental work that has been performed to investigate the effects of film coolant injection on convection heat transfer to the turbine blade tip for a variety of tip shapes and coolant injection configurations. Experiments are conducted with blade tip models that are stationary relative to the simulated outer air seal based on the result of earlier works that found the leakage flow to be mainly a pressure-driven flow which is related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud. Both heat transfer and film effectiveness are measured locally over the test surface using a transient thermal liquid crystal test technique with a computer vision data acquisition and reduction system for various combinations of clearance heights, clearance flow Reynolds numbers, and film flow rates with different coolant injection configurations. The present results reveal a strong dependency of film cooling performance on the choice of the coolant supply hole shapes and injection locations for a given tip geometry.

Effects of Labyrinth Fin Wear on Aerodynamic Performance of Turbine Stages: Part II — Mushrooming Damages

2019 ◽  
Author(s):  
Xinbo Dai ◽  
Xin Yan
Keyword(s):  
Experimental Studies ◽  
Axial Direction ◽  
Labyrinth Seal ◽  
Aerodynamic Performance ◽  
Leakage Rate ◽  
Secondary Flows ◽  
Main Function ◽  
Aerodynamic Loss ◽  
Solution Methods ◽  
Isentropic Efficiency

Abstract The main function of labyrinth seal is to control leakage flow in clearance that involves with rotating and stationary parts. Therefore, the effective of clearance gap in labyrinth seal is critical to sealing, heat transfer and vibration characteristics. However, due to the mechanical expansions, vibrations, thermal stress, misalignment of seal components in transient periods of startup, shutdown and hot restart, the stationary and rotating parts of the labyrinth seal are likely to contact each other, causing wear damages in labyrinth fin. Mushrooming damages are often occurred in the rubbing events when labyrinth fin is made of soft material compared with the opposite component. To investigate how mushrooming damage affects the leakage performance of labyrinth seal, many numerical and experimental studies have been carried out in last decades. However, little attention has been paid on the influence of labyrinth fin mushrooming on aerodynamic performance of turbine stages. In this study, using the RANS equations solution methods, the effect of labyrinth fin mushrooming on isentropic efficiency, leakage rates, outlet flow angles, reaction degrees, profile static pressure distributions and flow fields in turbine stages were investigated at three different mushrooming radii and three effective clearances. It shows the leakage rate is increased with increasing the mushroom radius and effective clearance. At the same effective clearance, as the mushrooming radius increases from 0.2mm to 0.4mm, the leakage rate is increased by about 0.19–0.32%, and the overall isentropic efficiency is decreased by 0.78%. At the same mushrooming radius, as the effective clearance increases from 1mm to 1.4mm, the leakage rate is increased by 0.21–0.31%, and the overall isentropic efficiency is decreased by 0.65%. As mushroom radius and effective clearance increase, the secondary flows near hub and shroud are intensified and developed along axial direction, causing pronounced aerodynamic loss in turbine stages.

Darryl E. Metzger Memorial Session Paper: A Summary of the Cooled Turbine Blade Tip Heat Transfer and Film Effectiveness Investigations Performed by Dr. D. E. Metzger

1995 ◽  
Vol 117 (1) ◽  
pp. 1-11 ◽  
Author(s):  
Y. W. Kim ◽  
J. P. Downs ◽  
F. O. Soechting ◽  
W. Abdel-Messeh ◽  
G. D. Steuber ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Convection Heat Transfer ◽  
Secondary Flows ◽  
Test Surface ◽  
Leakage Flow ◽  
Test Technique ◽  
Coolant Injection ◽  
Blade Tip

The clearance gap between the stationary outer air seal and blade tips of an axial turbine allows a clearance gap leakage flow to be driven through the gap by the pressure-to-suction side pressure difference. The presence of strong secondary flows on the pressure side of the airfoil tends to deliver air from the hottest regions of the mainstream to the clearance gap. The blade tip region, particularly near the trailing edge, is very difficult to cool adequately with blade internal coolant flow. In this case, film cooling injection directly onto the blade tip region can be used in an attempt to directly reduce the heat transfer rates from the hot gases in the clearance gap to the blade tip. The present paper is intended as a memorial tribute to the late Professor Darryl E. Metzger, who made significant contributions in this particular area over the past decade. A summary of this work is made to present the results of his more recent experimental work, which was performed to investigate the effects of film coolant injection on convection heat transfer to the turbine blade tip for a variety of tip shapes and coolant injection configurations. Experiments are conducted with blade tip models that are stationary relative to the simulated outer air seal based on the result of earlier works that found the leakage flow to be mainly a pressure-driven flow, which is related strongly to the airfoil pressure loading distribution, and only weakly, if at all, to the relative motion between blade tip and shroud. Both heat transfer and film effectiveness are measured locally over the test surface using a transient thermal liquid crystal test technique with a computer vision data acquisition and reduction system for various combinations of clearance heights, clearance flow Reynolds numbers, and film flow rates with different coolant injection configurations. The present results reveal a strong dependency of film cooling performance on the choice of the coolant supply hole shapes and injection locations for a given tip geometry.

Investigation of 3D Blading and Non-Axisymmetric Hub Endwall Contouring to a Dual-Stage Counter-Rotating Compressor in Multistage Environment

2014 ◽  
Author(s):  
Peng Zhang ◽  
Bo Liu ◽  
Guochen Zhang ◽  
Zhenzhe Na
Keyword(s):  
Secondary Flow ◽  
Flow Separation ◽  
Guide Vane ◽  
Aerodynamic Performance ◽  
Secondary Flows ◽  
Flow Losses ◽  
Dual Stage ◽  
Endwall Contouring

3D blading technique is an effective way to mitigate flow separation and improve the performance of turbomachinery. Non-axisymmetric endwall contouring technique is widely used to reduce transverse secondary flows near the endwall region in turbines, while the application of non-axisymmetric endwall contouring to compressors is rare. The investigation rig is the Northwestern Polytechnical University dual-stage counter-rotating compressor. In order to further improve the aerodynamic performance of the compressor, 3D blading optimization of the two rotors and outlet guide vane (OGV) was done in multistage environment at near stall condition. After 3D blading optimization, the radial secondary flows of the two rotors and the separation vortex at the tip of OGV are obviously reduced. However, the transverse secondary flows at the hub endwalls of the two rotors are scarcely improved. Based on the 3D blading optimization, non-axisymmetric hub endwall contouring optimization of the two rotors was carried out afterwards and the circumferential secondary flow losses of the two rotors are reduced effectively. The efficiency of the counter-rotating compressor on optimization point increases and the aerodynamic performance is improved largely after 3D blading and non-axisymmetric hub endwall contouring optimization.

The Effect of Axisymmetric Profile on Turbine Blade Platform Heat Transfer Distribution

2013 ◽  
Author(s):  
Luzeng Zhang ◽  
Dong H. Lee ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Heat Transfer Rate ◽  
Turbine Blade ◽  
Transfer Rate ◽  
Heat Load ◽  
Secondary Flows ◽  
Upstream Region ◽  
Aerodynamic Loss ◽  
Flat Profile

Flow field near the turbine blade platform is very complex due to the secondary flow motions such as horseshoe vortices, passage vortices and endwall cross flows. It is therefore extremely difficult to predict the platform heat transfer distribution. As the secondary flows are largely affected by platform profile/shape, a number of investigators have studied different platform profiles to minimize aerodynamic loss and heat load. Understanding of the platform heat transfer has become especially critical in recent years, because of firing temperature increase and low NOx combustion requirement, as it is directly related to turbine durability. Three different axisymmetric platform profiles were designed and experimentally studied: flat profile, dolphin nose profile and shark nose profile. All of them were based on the existing engine hardware designs. The measurements were conducted in a high-speed linear cascade, which consisted of five blades and six flow passages. The test platforms were made of FR4 material and painted with Thermo-chromic Liquid Crystal (TLC). The test article was kept in the plenum located under the cascade at the pre-test condition. At the start of each test, the test blade/article was inserted into the cascade rapidly and then two CCD cameras recorded the color changes of the TLC on the platform surface. Engine representative Reynolds numbers were studied from 300,000 to 600,000 and the corresponding inlet Mach numbers were ranged from 0.12 to 0.24. The upstream section of the flat profile platform showed a typical flat plate heat transfer pattern with boundary layer development. The shark-nose and dolphin-nose platforms resulted in lower heat transfer coefficients on the upstream region compared to that for the flat profile, and the peak values moved slightly downstream from the leading edge due to possibly different secondary flow patterns. The heat transfer rate increased with increased Reynolds number for all three platform shapes, while the flat profile showed a higher increase rate. Zone averaged heat transfer distributions in addition to local values were also presented to show the effect of platform profile. In general, the flat profile platform resulted in a higher overall heat transfer rate than that for the other two profile platforms, which suggested that a good design of contoured profile platform could reduce the heat load and aerodynamic loss in gas turbine blade.

Aerodynamic Performance of Novel Lightweight Turbine Blade

2016 ◽  
Author(s):  
Yoji Okita ◽  
Kozo Nita ◽  
Seiji Kubo
Keyword(s):  
Mach Number ◽  
Turbine Blade ◽  
Weight Reduction ◽  
Pressure Loss ◽  
High Speed ◽  
Aerodynamic Performance ◽  
Pressure Side ◽  
Design Point ◽  
Aerodynamic Loss ◽  
Cooling Passage

The primary contribution of this research is to clarify the aerodynamic performance of a novel lightweight turbine blade with internal cooling passage and external film cooling, which is invented aiming at drastic weight reduction of a cooled blade. With a considerably thinner airfoil, a significant separation region is formed along the pressure side, and therefore aerodynamic performance with such a flow field should be investigated. First, the lightweight cooled airfoil is designed. In the design process, a conventional thick airfoil is first defined as a baseline. With the baseline airfoil, only the mid and rear part of pressure side profile is redesigned to thin the airfoil without any change in the suction side geometry. The airfoil geometry is optimized so as not to bring significant aerodynamic loss increase. In this numerical optimization, the airfoil shape is gradually changed and evaluated step by step. In every step, an adjoint variable method is used to seek better airfoil shape, and then the generated new shape is evaluated with full RANS calculation. This iteration is repeated until any further recognizable weight reduction cannot be obtained without sensitive pressure loss increase and/or the airfoil shape reaches some geometrical constraints. The resultant optimized airfoil is approximately 20% lighter than the baseline hollow airfoil without any noticeable change in aerodynamic loss in the numerical solution. Next, the optimized airfoil is tested in a high speed linear cascade rig to verify its aerodynamic performance. The baseline airfoil is also tested for comparison. The rig is composed of six airfoil passages. The compressed air is supplied to the cascade and discharges to the atmospheric exhaust chamber. The air is also heated up to about 540 K upstream of the cascade. The cascade exit Mach number at the design point is 1.25, while in the experiment other several off-design conditions are also tested to check if there is any Mach number sensitivity. At the design point, the optimized lightweight airfoil shows less total pressure loss compared to the baseline airfoil. Also, at any other off-design Mach number conditions tested, the magnitude of the pressure loss is less with the lightweight airfoil. These results verify that the proposed airfoil does not only bring a considerable weight advantage but also compares favorably with the conventional airfoil in aerodynamic performance.

Aerodynamic Measurements in a Linear Turbine Blade Passage With Three-Dimensional Endwall Contouring

2007 ◽  
Author(s):  
Ross Gustafson ◽  
Gazi Mahmood ◽  
Sumanta Acharya
Keyword(s):  
Turbine Blade ◽  
Thermal Loading ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Pressure Probe ◽  
Blade Passage ◽  
Pressure Losses ◽  
Smoke Flow ◽  
Turbulence Data ◽  
Endwall Contouring

Secondary flows created in the turbine blade passage contribute to greater pressure losses and increase the thermal loading on the endwall. The present paper explores the concept of utilizing a three-dimensional endwall for weakening the secondary flows in a low speed linear turbine blade cascade. The endwall contour is designed to reduce the near-wall cross-passage pressure gradient that drives the crossflow. Measurements undertaken include smoke flow visualization, velocity and turbulence data using a hot wire anemometer, velocity and pressure measurements using a five-hole pressure probe, and surface static pressures measurements. The measured data indicate that the endwall contour reduces the mass averaged pressure losses across the blade passage by over 30%. Endwall contouring also produces a much more uniform exit flow field.

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