Film cooling characteristics on blade platform with a leakage flow through mid-passage gap

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.

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Leakage flow analysis in the gas turbine shroud gap

Aircraft Engineering and Aerospace Technology ◽

10.1108/aeat-01-2018-0038 ◽

2019 ◽

Vol 91 (8) ◽

pp. 1077-1085 ◽

Cited By ~ 1

Author(s):

Filip Wasilczuk ◽

Pawel Flaszynski ◽

Piotr Kaczynski ◽

Ryszard Szwaba ◽

Piotr Doerffer ◽

...

Keyword(s):

Numerical Simulation ◽

Gas Turbine ◽

Mass Flow ◽

Labyrinth Seal ◽

Leakage Flow ◽

Original Design ◽

Reference Case ◽

Content Type ◽

Flow Through ◽

The Impact

Purpose The purpose of the study is to measure the mass flow in the flow through the labyrinth seal of the gas turbine and compare it to the results of numerical simulation. Moreover the capability of two turbulence models to reflect the phenomenon will be assessed. The studied case will later be used as a reference case for the new, original design of flow control method to limit the leakage flow through the labyrinth seal. Design/methodology/approach Experimental measurements were conducted, measuring the mass flow and the pressure in the model of the labyrinth seal. It was compared to the results of numerical simulation performed in ANSYS/Fluent commercial code for the same geometry. Findings The precise machining of parts was identified as crucial for obtaining correct results in the experiment. The model characteristics were documented, allowing for its future use as the reference case for testing the new labyrinth seal geometry. Experimentally validated numerical model of the flow in the labyrinth seal was developed. Research limitations/implications The research studies the basic case, future research on the case with a new labyrinth seal geometry is planned. Research is conducted on simplified case without rotation and the impact of the turbine main channel. Practical implications Importance of machining accuracy up to 0.01 mm was found to be important for measuring leakage in small gaps and decision making on the optimal configuration selection. Originality/value The research is an important step in the development of original modification of the labyrinth seal, resulting in leakage reduction, by serving as a reference case.

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Turbine Vane Endwall Film Cooling From Cross-Row Configuration With Simulated Upstream Leakage Flow

Volume 5C: Heat Transfer ◽

10.1115/gt2017-63145 ◽

2017 ◽

Cited By ~ 4

Author(s):

Nafiz H. K. Chowdhury ◽

Chao-Cheng Shiau ◽

Je-Chin Han ◽

Luzeng Zhang ◽

Hee-Koo Moon

Keyword(s):

Film Cooling ◽

Density Ratio ◽

Companion Paper ◽

Leakage Flow ◽

Blow Down ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Free Stream Turbulence ◽

Lift Off ◽

Endwall Cooling

The performance of a full coverage film cooling configuration called cross-row (CR) configuration including upstream inlet leakage flow was studied by measuring the adiabatic film cooling effectiveness distribution using PSP technique. Experiments were conducted in a blow-down wind tunnel cascade facility at the isentropic exit Mach number of 0.5 corresponding to inlet Reynolds number of 3.8 × 105, based on axial chord length. A free-stream turbulence level was generated as high as 19% with a length scale of 1.7 cm at the inlet. The results are presented as two-dimensional adiabatic film cooling effectiveness distributions on the endwall surface with corresponding spanwise averaged distributions. The focus of this study is to investigate the effect of coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR) on the proposed endwall cooling design. Initially, increased MFR for the endwall cooling and upstream leakage levels up the local adiabatic cooling effectiveness and yields relatively uniform coverage on the entire endwall. However, in either case, highest MFR does not provide any improvement as endwall cooling suffered from the jet lift-off and leakage coolant coverage restricted by the downstream near-wall flow field. Results also indicated a density ratio of 1.5 provides the best performance. Finally, a fair comparison is made with another design called axial-row (AR) configuration from a companion paper.

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Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

Volume 5C: Heat Transfer ◽

10.1115/gt2016-56321 ◽

2016 ◽

Cited By ~ 1

Author(s):

Andrew F. Chen ◽

Chao-Cheng Shiau ◽

Je-Chin Han

Keyword(s):

Turbine Blade ◽

Film Cooling ◽

Leading Edge ◽

Suction Side ◽

Leakage Flow ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Purge Flow ◽

Endwall Film Cooling ◽

Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

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Numerical Study of Heat Transfer and Cooling Effectiveness on a Flat-Tip Turbine Blade

ASME 2013 Turbine Blade Tip Symposium ◽

10.1115/tbts2013-2053 ◽

2013 ◽

Author(s):

Qihe Huang ◽

Jiao Wang ◽

Lei He ◽

Qiang Xu

Keyword(s):

Heat Transfer ◽

Turbine Blade ◽

Film Cooling ◽

Numerical Study ◽

Leakage Flow ◽

Tip Leakage Flow ◽

Cooling Effectiveness ◽

Blowing Ratio ◽

Flow And Heat Transfer ◽

Blade Tip

A numerical study is performed to simulate the tip leakage flow and heat transfer on the first stage rotor blade tip of GE-E3 turbine, which represents a modern gas turbine blade geometry. Calculations consist of the flat blade tip without and with film cooling. For the flat tip without film cooling case, in order to investigate the effect of tip gap clearance on the leakage flow and heat transfer on the blade tip, three different tip gap clearances of 1.0%, 1.5% and 2.5% of the blade span are considered. And to assess the performance of the turbulence models in correctly predicting the blade tip heat transfer, the simulations have been performed by using four different models (the standard k-ε, the RNG k-ε, the standard k-ω and the SST models), and the comparison shows that the standard k-ω model provides the best results. All the calculations of the flat tip without film cooling have been compared and validated with the experimental data of Azad[1] and the predictions of Yang[2]. For the flat tip with film cooling case, three different blowing ratio (M = 0.5, 1.0, and 1.5) have been studied to the influence on the leakage flow in tip gap and the cooling effectiveness on the blade tip. Tip film cooling can largely reduce the overall heat transfer on the tip. And the blowing ratio M = 1.0, the cooling effect for the blade tip is the best.

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Experimental Investigation of Turbine Phantom Cooling on Suction Side With Combustor-Turbine Leakage Gap Flow and Endwall Film Cooling

Volume 4: Heat Transfer, Parts A and B ◽

10.1115/gt2012-69295 ◽

2012 ◽

Cited By ~ 6

Author(s):

Yang Zhang ◽

Xin Yuan

Keyword(s):

Film Cooling ◽

Incidence Angle ◽

Suction Side ◽

Leakage Flow ◽

Airfoil Surface ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Blowing Ratio ◽

Coolant Injection ◽

Passage Vortex

The film cooling injection on Hp turbine component surface is strongly affected by the complex flow structure in the nozzle guide vane or rotor blade passages. The action of passage vortex near endwall surface could dominate the film cooling effectiveness distribution on the component surfaces. The film cooling injections from endwall and airfoil surface are mixed with the passage vortex. Considering a small part of the coolant injection from endwall will move towards the airfoil suction side and then cover some area, the interaction between the coolants injected from endwall and airfoil surface is worth investigating. Though the temperature of coolant injection from endwall increases after the mixing process in the main flow, the injections moving from endwall to airfoil suction side still have the potential of second order cooling. This part of the coolant is called “Phantom cooling flow” in the paper. A typical scale-up model of GE-E3 Hp turbine NGV is used in the experiment to investigate the cooling performance of injection from endwall. Instead of the endwall itself, the film cooling effectiveness is measured on the airfoil suction side. This paper is focused on the combustor-turbine interface gap leakage flow and the coolant from fan-shaped holes moving from endwall to airfoil suction side. The coolant flow is injected at a 30deg angle to the endwall surface both from a slot and four rows of fan-shaped holes. The film cooling holes on the endwall and the leakage flow are used simultaneously. The blowing ratio and incidence angle are selected to be the parameters in the paper. The experiment is completed with the blowing ratio changing from M = 0.7 to M = 1.3 and the incidence angle varying from −10deg to +10deg, with inlet Reynolds numbers of Re = 3.5×105 and an inlet Mach number of Ma = 0.1.

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An Experimental Study of the Leakage Flow Effect on the Film Cooling Effectiveness of a Gas Turbine Shroud

Journal of Thermal Science and Engineering Applications ◽

10.1115/1.4050566 ◽

2021 ◽

pp. 1-18

Author(s):

Gi Mun Kim ◽

Soo In Lee ◽

Jin Young Jeong ◽

Jae Su Kwak ◽

Seokbeom Kim ◽

...

Keyword(s):

Gas Turbine ◽

Film Cooling ◽

Density Ratio ◽

Secondary Flows ◽

Leakage Flow ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

High Heat Transfer ◽

Hole Configuration ◽

Cooling Hole

Abstract In the vicinity of gas turbine blades, a complex flow field is formed due to the flow separation, reattachment, and secondary flows, and this results in a locally non-uniform and high heat transfer on the surfaces. The present study experimentally investigates the effects of leakage flow through the slot between the gas turbine vane and blade rows on the film cooling effectiveness of the forward region of the shroud ring segment. The experiment is carried out in a linear cascade with five blades. Instead of the vane, a row of rods at the location of the vane trailing edge is installed to consider the wake effect. The leakage flow is introduced through the slot between the vane and blade rows, and additional coolant air is injected from the cooling holes installed at the vane's outer zone. The effects of the slot geometry, cooling hole configuration, and blowing ratio on the film cooling effectiveness are experimentally investigated using the pressure sensitive paint (PSP) technique. CO2 gas and a mixture of SF6 and N2 (25%+75%) are used to simulate the leakage flow to the mainstream density ratios of 1.5 and 2.0, respectively. The results indicate that the area averaged film cooling effectiveness is affected more by the slot width than by the cooling hole configuration at the same injection conditions, and the lower density ratio cases show higher film cooling effectiveness than the higher density ratio case at the same cooling configuration.

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Numerical Simulation of Evaporating Two-Phase Flow in a High-Aspect-Ratio Microchannel with Bends

Journal of Heat Transfer ◽

10.1115/1.4036882 ◽

2017 ◽

Vol 139 (8) ◽

Author(s):

Junsik Lee ◽

Junsub Kim ◽

Hyungsoo Lim ◽

Je Sung Bang ◽

Jeong Min Seo ◽

...

Keyword(s):

Boundary Layer ◽

Flow Visualization ◽

Film Cooling ◽

High Speed ◽

National Research Council ◽

Air Cooling ◽

Hole Size ◽

High Efficient ◽

Flow Through ◽

Effusion Hole

Effusion cooling is one of the attractive methods for next generation high-efficient gas turbine which has a very hot gas temperature above 1,600oC. For higher effectiveness of the air cooling, the air-cooled flow through effusion-holes does not penetrate into the mainstream flow but still remains within freestream boundary layer. So the air-cooled surface temperature maintains at relatively lower than film cooling. Effusion cooling is generally known as operating in small effusion-hole size which is less than 0.2 mm. This study is intended to examine optimum effusion-hole size of the microscale effusion cooling through flow visualization. The air flow through effusion-holes is visualized using an oil atomizer, a DSPP laser-sheet illumination, and a high-speed CCD imaging. The visualized results show flow patterns and characteristics with different blowing ratio, BR = ρcUc / ρ∞U∞, (BR = 0.17 and 0.53) and effusion-hole size (D = 0.2 mm, 0.5 mm and 1.0 mm). The flow visualization condition is fixed at the mainstream Reynolds number of 10,000 and hole-to-hole spacing of 4 (S/D = 4). For larger effusion-hole of 1.0 mm [(a) and (b)], the effusion flow can penetrate into boundary layer which exhibits a film cooling. However the effusion flow is observed to be remained within boundary layer which shows an effusion cooling for smaller effusion-hole of 0.2 mm [(e) and (f)]. In case of (c) and (d), a series of vortical structure is also observed to be within the boundary layer along the effusion flat plate. Note that the effusion-hole size of 0.5 mm can be a candidate for making effusion cooling possible. [This work was supported by National Research Council of Science and Technology (NST) grant funded by the Ministry of Science, ICT and Future Planning, Korea (Grant No. KIMM-NK203B).]

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Calculations of Cooled Turbine Efficiency

Journal of Engineering for Gas Turbines and Power ◽

10.1115/1.2771250 ◽

2008 ◽

Vol 130 (1) ◽

Cited By ~ 10

Author(s):

J. H. Horlock ◽

Leonardo Torbidoni

Keyword(s):

Film Cooling ◽

Gas Flow ◽

Coolant Flow ◽

The Other ◽

Turbine Stage ◽

Isentropic Expansion ◽

Turbine Efficiency ◽

Flow Through ◽

Definition Of ◽

Industrial Gas Turbine

The efficiency of a cooled turbine stage has been discussed in the literature. All proposed definitions compare the actual power output with an ideal output, which has to be determined; but usually, one of two definitions has been used by turbine designers. In the first, the so-called Hartsel efficiency, the mainstream gas flow, and the various coolant flows to rotor and stator are assumed to expand separately and isentropically to the backpressure. In the second, it is assumed that these flows mix at constant (mainstream) gas pressure before expanding isentropically (sometimes, the rotor coolant flow is ignored in this definition). More recently, it has been suggested that a thermodynamically sounder definition is one in which the gas and coolant flows mix reversibly and adiabatically before isentropic expansion to the backpressure. In the current paper, these three efficiencies are compared, for a typical stage—the first cooled stage of a multistage industrial gas turbine. It is shown that all the efficiencies fall more or less linearly with increase of the fractional (total) coolant flow. It is also shown that the new definition of efficiency gives values considerably lower than the other two efficiencies, which are more widely used at present. Finally, the various irreversibilities associated with the flow through a cooled turbine are calculated. Although all these irreversibilities increase with the fractional coolant flow, it is shown that the “thermal” irreversibility associated with film cooling is higher than the other irreversibilities at large fractional coolant flow.

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