AEROTHERMAL EFFECT OF CAVITY WELDING BEADS ON A TRANSONIC SQUEALER TIP

2021 ◽  
pp. 1-12
Author(s):  
Joao Vieira ◽  
John D Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Performance Measurements ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

2020 ◽  
Author(s):  
Joao Vieira ◽  
John Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Reynolds Numbers ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

Numerical Simulation of Film Cooling on Rotating Blade Tips Within a High-Pressure Turbine

2013 ◽  
Author(s):  
K. Lu ◽  
M. T. Schobeiri ◽  
J. C. Han
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Flow Structures ◽  
Research Facility ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Rotating Blade ◽  
Blade Tip ◽  
Blade Tips

This paper numerically investigates the aerodynamics and film cooling effectiveness of high pressure turbine blade tips. Two different rotor blade tip configurations have been studied: the plane tip with tip hole cooling and the squealer tip with tip hole cooling. The geometry of the blades is determined based on the blade profiles within the three-stage multi-purpose turbine research facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A&M University. Seven perpendicular holes along the camber line are used for the tip hole cooling. The clearance between the blade tip and casing is 1.0% of the blade span. For each blade tip configuration, the coolant is ejected through the cooling holes under blowing ratios of M = 0.5, 1.0 and 1.5. In this paper, a comparison between the plane tip and the squealer tip has been presented. The detailed flow structures and film cooling effectiveness are discussed.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer, and film cooling effectiveness of advanced high pressure turbine blade tips and endwalls. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with leading edge and trailing edge cutouts. Both blade tip configurations have pressure side film cooling and cooling air extraction through dust holes, which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9×105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aerothermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the midchord region. However, on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2010 ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer and film cooling effectiveness of advanced high-pressure turbine blade tips and endwall. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with a leading edge and trailing edge cut-out. Both blade tip configurations have pressure side film cooling, and cooling air extraction through dust holes which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9 × 105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aero-thermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although, the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the mid-chord region. However on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

Experimental Study of Leakage Flow on Flow Field and Film Cooling of High Pressure Turbine Blade Platform

2015 ◽  
Author(s):  
Kenichiro Takeishi ◽  
Yutaka Oda ◽  
Shintaro Kozono
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Temperature Fields ◽  
Leakage Flow ◽  
High Pressure Turbine ◽  
Film Cooling Effectiveness ◽  
Injection Angle

An experiment has been conducted to study stator/rotor disc cavity leakage flow on the platform of a highly loaded stationary linear blade cascade. The linear cascade consists of a scaled-up model of the high-pressure turbine blades of an E3 (Energy efficient engine) and leakage slot models installed under the platform. Experiments have been conducted to investigate the effect of the slot injection angle, leakage flow rates, distance between the leading edge of the blade and the slot, and spacing of the blades. The film-cooling effectiveness was measured by pressure sensitive paint (PSP), and the temperature fields and flow fields were investigated using laser-induced fluorescence (LIF) and particle image velocimetry (PIV), respectively. It was observed from the experiments that the leakage flow covered the surface of the blade platform when the distance between the leading edge and the slot was zero; however, with increasing distance, the horseshoe vortex dominates near the junction of the blade leading edge, and the leakage flow could not cover the region. It was also found that the leakage flow has an effect that promotes the formation of the horseshoe vortex for some experimental conditions.

Optimal Arrangement of the Film Cooling Holes Considering the Manufacturing Tolerance for High Pressure Turbine Nozzle

2016 ◽  
Author(s):  
Sanga Lee ◽  
Dong-Ho Rhee ◽  
Kwanjung Yee
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Manufacturing Tolerance ◽  
Cooling Hole ◽  
Film Cooling Holes

In spite of a myriad of researches on the optimal shape of film cooling holes, only a few attempts have been made to optimize the hole arrangement for film cooling so far. Moreover, although the general scale of film cooling hole is so small that manufacturing tolerance has substantial effects on the cooling performance of turbine, the researches on this issue are even scarcer. If it is possible to obtain optimal hole arrangement which not only improve the film cooling performance but also is robust to the manufacturing tolerance, then overall cooling performance of a turbine would become more reliable and useful from the practical point of view. To this end, the present study proposed a robust design optimization procedure which takes the manufacturing uncertainties into account. The procedure was subsequently applied to the film cooling holes on high pressure turbine nozzle pressure side to obtain the robust array shape under the uncertainty of the manufacturing tolerance. First, the array of the holes was parameterized by 5 design variables using the newly suggested shape functions, and 2 representative factors were considered for the manufacturing tolerance of the film cooling hole. Probabilistic process that consists of Kriging surrogate model and Monte Carlo Simulation with descriptive sampling method was coupled with the design optimization process using Genetic Algorithm. Through this, film cooling hole array which shows the high performance, yet robust to the manufacturing tolerance was obtained, and the effects of the manufacturing tolerance on the cooling performance was carefully investigated. As a result, the region where the film cooling effectiveness is noticeable, as well as the maximum width of the variation of the film cooling effectiveness were reduced through optimization, and it is also confirmed that the tolerance of the holes near the leading edge is more influential to the cooling performance because the film cooling effectiveness is more sensitive to the manufacturing tolerance of the leading edge than that of the trailing edge.

scholarly journals Computational Modeling and Thermal Paint Verification of Film-Cooling Designs for an Unshrouded High-Pressure Turbine Blade

1999 ◽  
Author(s):  
Steven G. Gegg ◽  
Nathan J. Heidegger ◽  
Ronald A. Mikkelson
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
Data Exchange ◽  
Turbine Blades ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Extreme High Temperature ◽  
Blade Tip

High pressure turbine blades are exposed to an extreme high temperature environment due to increasing turbine inlet temperature. High heat fluxes are likely on the blade pressure surface. Other regions, such as the trailing edge and blade tip may be difficult to cool uniformly. Unshrouded blades present an additional challenge due to the pressure driven transport of hot gas across the blade tip. The blade tip region is therefore prone to severe thermal stress, fatigue and oxidation. In order to develop effective cooling methods, designers require detailed flow and heat transfer information. This paper reports on computational aerodynamics and heat transfer studies for an unshrouded high pressure turbine blade. The emphasis is placed on the application of appropriate 3-D models for the prediction of airfoil surface temperatures. Details of the film cooling model, boundary conditions and data exchange with heat transfer models are described. The analysis approach has been refined for design use to provide timely and accurate results. Film cooling designs are to be tailored for the best coverage of the blade tip region. Designs include near-tip pressure side films and blade tip cooling holes. Hole placement and angle are investigated to achieve the best coolant coverage on the blade tip. Analytical results are compared to a thermal paint test on engine hardware. In addition to film cooling strategies, other aerodynamic/heat transfer design approaches are discussed to address the cooling requirements for an unshrouded blade.

Optimization of Film Cooling Holes on the Suction Surface of a High Pressure Turbine Blade

2012 ◽  
Author(s):  
Carole El Ayoubi ◽  
Wahid Ghaly ◽  
Ibrahim Hassan
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Three Dimensional ◽  
Suction Surface ◽  
Cfd Simulations ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Flow Parameters ◽  
Film Cooling Holes

This paper aims to optimize film coolant flow parameters on the suction surface of a high-pressure gas turbine blade in order to obtain an optimum compromise between a high film cooling effectiveness and a low aerodynamic loss. An optimization algorithm coupled with three-dimensional Reynolds-averaged Navier Stokes (RANS) analysis is used to determine the optimum film cooling configuration. The VKI blade with two staggered rows of axially oriented, conically flared, film cooling holes on its suction surface is considered. Two design variables are selected; the coolant to mainstream temperature ratio and total pressure ratio. The effect of varying these coolant flow parameters on the film cooling effectiveness and the aerodynamic loss is analyzed using an optimization method and three dimensional steady CFD simulations. The optimization process involves a genetic algorithm and a response surface approximation of the artificial neural network type to provide low-fidelity predictions of the objective function. The CFD simulations are performed using the commercial software CFX. The numerical predictions of the aerodynamics and wall heat transfer are validated against experimental data. The optimization objective consists of maximizing the spatially averaged film cooling effectiveness and minimizing the aerodynamic penalty produced by film cooling. The results of this optimization are reported in terms of the aerodynamic loss and adiabatic cooling effectiveness.

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