Cooling of Turbine Blade Surface With Expanded Exit Holes: Computational Suction-Side Analysis

2016 ◽  
Vol 138 (5) ◽  
Author(s):  
Fariborz Forghan ◽  
Omid Askari ◽  
Uichiro Narusawa ◽  
Hameed Metghalchi
Keyword(s):  
Mass Flow Rate ◽  
Turbine Blade ◽  
Flow Rate ◽  
Temperature Difference ◽  
Film Cooling ◽  
Gas Flow ◽  
Suction Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Flow Through

Turbine blade surfaces are cooled by jet flow from expanded exit holes (EEHs) against the prevailing hot gas flow. The flow through EEH must be designed to form a film of cool air over the blade. Computational analyses are performed to examine the cooling effectiveness of flow from EEH over the suction side of a blade by solving conservation equations and the ideal gas equation of state for turbulent and compressible flow. For a sufficiently high coolant mass flow rate, the flow through EEH, which acts as a converging–diverging nozzle, is choked at the nozzle throat, resulting in a supersonic flow, a shock, and then a subsonic flow downstream. The location of the shock relative to the high-temperature gas flow determines the temperature distribution along the blade surface; which is analyzed in detail when the following conditions are varied: coolant mass flow rate, the temperature difference between gas-and coolant-flow, EEH location on the blade surface, EEH inclination angle to the blade surface, and exit-to-inlet area ratio (AR) of EEH. The film cooling effectiveness is calculated along the surface of the blade. The results show (1) increasing the coolant flow rate improves the effectiveness, (2) change in temperature difference between the mainstream and the coolant slightly affects the effectiveness, (3) inclination angle of EEH has a pronounced effect on film cooling and the corresponding effectiveness, (4) both the location of the EEH on a blade and the AR of the EEH slightly change the effectiveness.

Film Cooling of Turbine Blade Surface With Extended Exit Holes

2014 ◽  
Author(s):  
Fariborz Forghan ◽  
Omid Askari ◽  
Uichiro Narusawa ◽  
Hameed Metghalchi
Keyword(s):  
High Temperature ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Flow ◽  
Three Dimensional ◽  
Suction Side ◽  
Ideal Gas ◽  
Coolant Flow ◽  
Blade Surface ◽  
Flow Through

The main goal of gas turbine design is the effective use of energy. Usually, the efficient high temperature first and second stage turbine blade surface is cooled by jet of coolant flow from extended exit holes (EEH). Against the prevailing hot gas flow, the flow through EEH must be designed to form a film of cool air over the blade. Computational analyses are performed to examine the cooling effectiveness of flow from EEH over the suction side of a blade by solving conservation equations (mass, momentum and energy) and the ideal gas equation of state for the three-dimensional, turbulent, compressible flow. A diverging flow through EEH is typically choked at its throat, resulting in a supersonic flow, a shock and then a subsonic flow downstream. The location of the shock relative to the high-temperature gas flow over the blade determines the temperature distribution along the blade surface; which is analyzed in detail when the coolant flow rate is varied.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

2008 ◽  
Vol 131 (1) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta P. Narzary ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Compound Angle

The film-cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint technique. Compound angle laidback fan-shaped holes are used to cool the blade surface with four rows on the pressure side and two rows on the suction side. The coolant injects to one side of the blade, either pressure side or suction side. The presence of wake due to the upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate progressing wakes. The effect of wakes is recorded at four phase locations along the pitchwise direction. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The inlet and exit Mach numbers are 0.27 and 0.44, respectively, resulting in a pressure ratio of 1.14. Five average blowing ratios ranging from 0.4 to 1.5 are tested. Results reveal that the tip-leakage vortices and endwall vortices sweep the coolant on the suction side to the midspan region. The compound angle laidback fan-shaped holes produce a good film coverage on the suction side except for the regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and the effectiveness level is low on the pressure surface. However, the pressure side acquires a relatively uniform film coverage with the multiple rows of cooling holes. The film-cooling effectiveness increases with the increasing average blowing ratio for either side of coolant ejection. The presence of stationary upstream wake results in lower film-cooling effectiveness on the blade surface. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film-cooling effectiveness, particularly at higher blowing ratios.

A One-Dimensional Analytical Method for Turbine Blade Preliminary Cooling Design

2016 ◽  
Author(s):  
Chen Li ◽  
Jian-jun Liu
Keyword(s):  
Mass Flow Rate ◽  
Turbine Blade ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Three Dimensional ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
One Dimensional ◽  
Cooling Design

The turbine blade cooling design is a complex procedure including one-dimensional preliminary cooling design, detailed two-dimensional design and fluid network analyses, and three-dimensional conjugate heat transfer and FEM predictions. Frequent alteration and modification of the cooling configurations make it unpractical to obtain all of three-dimensional design results quickly. Preliminary cooling design deals mainly with the coolant requirements and can be knitted into fluid network to look up the expected cooling structural style to promote three-dimensional geometry design. Previous methods to estimate the coolant requirements of the whole turbine blade in the preliminary cooling design were usually based on the semi-empirical air-cooled blade data. This paper combines turbine blade internal and external cooling, and presents a one-dimensional theoretical analytical method to investigate blade cooling performance, assuming that the coolant temperature increases along the blade span. Firstly, a function of non-dimensional cooling mass flow rate is derived to describe the new relationship between adiabatic film cooling effectiveness and overall cooling effectiveness. Secondly, a new variable related to film cooling is found to estimate the required adiabatic film cooling effectiveness without using the empirical correlations. Finally, a theoretical calculation about the relationship between non-dimensional cooling mass flow rate and overall cooling effectiveness well corresponds to semi-empirical air-cooled blade data within regular range of cooling efficiency. The currently proposed method is also a useful tool for the blade thermal analysis and the sensitivity analysis of coolant requirements to various design parameters. It not only can provide all the possible options at the given gas and coolant inlet temperatures to meet the design requirement, but also can give the third boundary conditions for calculating the blade temperature field. It’s convenient to use the heat transfer characteristic of internal cooling structures to estimate the coolant mass flow rate and the channel hydraulic diameter for both convection cooling and film cooling.

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

2016 ◽  
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.

Cooling Effectiveness on Film Cooled Turbine Blades

2001 ◽  
Author(s):  
M. Derrar ◽  
J. Nagler ◽  
W. W. Koschel
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Suction Side ◽  
Experimental Simulation ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Boundary Interaction ◽  
Simultaneous Modeling

Abstract This paper presents experiments on the cooling effectiveness obtained for two different injection locations on the suction side of a turbine blade at transonic flow conditions. Previous results of a computational analysis and flow visualization indicated that a separation bubble is present on the suction side at a location x/L = 0.43 and the location x/L = 0.575 corresponds to a shock-boundary interaction zone [9]. The scientific interest is primarily focused on the realization of high film cooling efficiencies and its relevant parameters under these flow conditions. Streamwise aligned as well as inclined angled film coolant hole configurations have been investigated for each location. Due to the high number of interacting parameters the experimental simulation of turbine blade film cooling is extremely complex, which can only be solved by a simultaneous modeling using the experimentally measured results. Test rig, instrumentation and data analysis are described in detail. The goal of the investigations is to determine the optimum location of the film coolant injection.

Study of Film Cooling Effectiveness on a Double-Walled Effusion-Cooled Turbine Blade in a High-Speed Flow Using Pressure Sensitive Paint

2019 ◽  
Author(s):  
Gladys C. Ngetich ◽  
Peter T. Ireland ◽  
Eduardo Romero
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
High Speed ◽  
High Porosity ◽  
Pressure Sensitive Paint ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Double Wall

Abstract A detailed analysis of film cooling performance on a double-walled effusion-cooled blade is essential for both the coolant consumption optimization and assessment of the film to offer the desired levels of the turbine blade protection. Yet there are hardly any film effectiveness studies on double-wall full-coverage film cooled turbine blades. This paper presents a detailed film cooling effectiveness study over the full surface of a double-walled effusion-cooled high-pressure turbine rotor blade using Pressure Sensitive Paint (PSP). PSP permitted a non-intrusive and conduction-errors-free means of obtaining clean and distinct local distribution of film effectiveness on the blade surface making it possible to extract valuable film cooling effectiveness performance data on the whole blade surface. Three large-scale circular pedestal double-wall blade designs with varying pedestal height, pedestal diameter and cooling hole diameter were tested in a high-speed stationary single-blade linear cascade running at engine-representative Mach and Reynolds numbers. All the blades were tested within a range of representative modern engine coolant mass flow, ṁc to mainstream, ṁg ratios; 1.6% < ṁc/ṁ∞ < 5.5%. High porosity blade exhibited a better flow distribution and was found to consistently perform the best.

Background-Grid Based Mapping Approach to Film Cooling Meshing: Part III — Applications in a Full-Cooled Turbine Vane

2020 ◽  
Author(s):  
Ruiqin Wang ◽  
Xin Yan
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Design Flow ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Structured Grids

Abstract To cool a high-pressure gas turbine blade, many rows of cooling holes with different arrangements and configurations are manufactured to achieve higher cooling effect and lower aerodynamic loss. To evaluate the heat transfer and film cooling effect in the full-cooled turbine blade, efficient numerical simulations are required in the design and performance optimization processes. From the view of numerical accuracy, the structured grids have to be employed because of higher resolution in flow and heat transfer than the unstructured grids. Because many splitting, attaching and merging manipulations are involved in meshing the cooling features and curved boundaries, it is very complex and time-consuming for a researcher to generate multi-block structured grids for a full-cooled gas turbine blade. As a result, in the industrial applications, almost all researchers preferred to generate unstructured grids instead of structured grids for the full-cooled blade. Unlike the previous research, the aim of this study is to apply the Background-Grid Based Mapping (BGBM) method proposed in Part I to generate multi-block structured grids for a full-cooled gas turbine vane. With the strategy of BGBM method, meshes were conveniently generated in the computational space with simple geometrical features and plain interfaces, and then were mapped back into physical space to obtain the multi-block structured grids which can be used for numerical simulations. With the experimental data, the present numerical methods and BGBM strategy were carefully validated. Then, the flow and film cooling performance in the full-cooled NASA GE-E3 nozzle guided vane were numerically investigated. The effects of coolant mass flow rate and land extensions on film cooling effectiveness were discussed. The results show that film cooling effectiveness near the stagnation point is the lowest and film cooling effectiveness on the pressure side is slightly higher than that on the suction side. When the coolant mass flow rate increases up to the value of 1.5 design flow, the relative outflow mass flow rates of cooling hole arrays and slots are no longer affected by the increase of the coolant flow rate. At half design flow, the outflow mass flow rates of No.5 hole-array to No.10 hole-array are almost zero, and the area-averaged film cooling effectiveness on vane surface is as low as 0.268. Compared with the cases of half design flow and double design flow, better film cooling performance is obtained in the cases of design flow and 1.5 design flow. Compared with the vane without lands, the area-average cooling effectiveness on vane surface is slightly higher for the vane with lands. Land extensions have a considerable influence on film cooling performance in the cutback region.

Influence of Coolant Density on Turbine Blade Platform Film-Cooling

2012 ◽  
Vol 4 (2) ◽  
Author(s):  
Diganta P. Narzary ◽  
Kuo-Chun Liu ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Flow Rate ◽  
Turbulence Intensity ◽  
Film Cooling ◽  
Density Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Mainstream Flow ◽  
Purge Flow

Detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform of a five-blade linear cascade. The parameters chosen were freestream turbulence intensity, upstream stator-rotor purge flow rate, discrete-hole film-cooling blowing ratio, and coolant-to-mainstream density ratio. The measurement technique adopted was temperature sensitive paint (TSP) technique. Two turbulence intensities of 4.2% and 10.5%; three purge flows between the range of 0.25% and 0.75% of mainstream flow rate; three blowing ratios between 1.0 and 1.8; and three density ratios between 1.1 and 2.2 were investigated. Purge flow was supplied via a typical double-toothed stator-rotor seal, whereas the discrete-hole film-cooling was accomplished via two rows of cylindrical holes arranged along the length of the platform. The inlet and the exit Mach numbers were 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 7.5 * 105 based on the exit velocity and chord length of the blade. Results indicated that platform film-cooling effectiveness decreased with turbulence intensity, increased with purge flow rate and density ratio, and possessed an optimum blowing ratio value.

Predictions of Cooling From Dirt Purge Holes Along the Tip of a Turbine Blade

2003 ◽  
Author(s):  
E. M. Hohlfeld ◽  
J. R. Christophel ◽  
E. L. Couch ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip ◽  
Flow Through ◽  
Film Cooling Holes

The clearance gap between the tip of a turbine blade and its associated shroud provides a flow path for leakage from the pressure side of the blade to the suction side. The tip region is one area that experiences high heat transfer and, as such, can be the determining factor for blade life. One method for reducing blade tip heat transfer is to use cooler fluid from the compressor, that exits from relatively large dirt purge holes placed in the tip, for cooling purposes. Dirt purge holes are typically manufactured in the blade tip to extract dirt from the coolant flow through centrifugal forces such that these dirt particles do not block smaller diameter film-cooling holes. This paper discusses the results of numerous computational simulations of cooling injection from dirt purge holes along the tip of a turbine blade. Some comparisons are also made to experimental results in which a properly scaled-up blade geometry (12X) was used to form a two-passage linear cascade. Computational results indicate that the cooling achieved through the dirt purge injection from the blade tip is dependent on the gap size as well as the blowing ratio. For a small tip gap (0.54% of the span) the flow exiting the dirt purge holes act as a blockage for the leakage flow across the gap. As the blowing ratio is increased for a large tip gap (1.63% of the span), the tip cooling increases only slightly while the cooling to the shroud increases significantly.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

2007 ◽  
Author(s):  
Zhihong Gao ◽  
Diganta P. Narzary ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Steady Increase ◽  
Pressure Side ◽  
Wake Effect ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Compound Angle

The film cooling effectiveness on the surface of a high pressure turbine blade is measured using the Pressure Sensitive Paint (PSP). Four rows of fan-shaped, laid-back compound angled cooling holes are distributed on the pressure side while two such rows are provided on the suction side of the blade. The coolant is only injected to either the pressure side or suction side of the blade at five average blowing ratios from 0.4 to 1.5. Presence of wake due to upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate a progressing wake. Effect of wake is recorded at four phase locations with equal intervals along the pitch-wise direction. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively, resulting in a blade pressure ratio of 1.14. Results reveal that the tip leakage vortices and endwall vortices sweep the coolant film on the suction side to the midspan region. The fan-shaped, laid-back compound angled holes produce good coolant film coverage on the suction side except for those regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and effectiveness level is low on the pressure surface. However, the pressure side acquires relatively uniform film coverage with the design of multiple rows of cooling holes. The presence of stationary upstream wake results in lower film cooling effectiveness on the blade surface. Variation of blowing ratio from 0.4 to 1.5 shows steady increase in effectiveness on the pressure side or the suction side for a given wake rod phases locations. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film cooling effectiveness particularly at higher blowing ratios.

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