Experimental Investigation of Film-Cooling Effectiveness of a Highly Loaded Turbine Blade Under Steady and Periodic Unsteady Flow Conditions

2017 ◽  
Vol 139 (7) ◽  
Author(s):  
Ali Nikparto ◽  
Meinhard T. Schobeiri
Keyword(s):  
Unsteady Flow ◽  
Film Cooling ◽  
Flow Condition ◽  
Leading Edge ◽  
Experimental Investigations ◽  
Pressure Side ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Highly Loaded

This paper describes the experimental investigations of film-cooling effectiveness on a highly loaded low-pressure turbine blade under steady and unsteady wake flow conditions. The cascade facility in Turbomachinery Performance and Flow Research Lab (TPFL) at the Texas A&M University was used to simulate the periodic flow condition inside gas turbine engines. Moving wakes, originated from upstream stator blades, are simulated inside the cascade facility by moving rods in front of the blades. The flow coefficient is maintained at 0.8 and the incoming wakes have a reduced frequency of 3.18. A total of 617 holes on the blade are distributed along 13 different rows. Six rows cover the suction side, six other rows cover the pressure side, and one last row feeds the leading edge. Each row has a twin row on the other side of the blade with exact same number of holes and arrangement (except for leading edge). They both are connected to the same cavity. Coolant is injected from either sides of the blade through cavities to form a uniform distribution along the span of the blade. Film-cooling effectiveness under periodic unsteady flow condition was studied using pressure-sensitive paint. Experiments were performed at Reynolds number of 150,000 and blowing ratio of one, based on equal mass flux distribution. Experimental investigations were performed to determine the effect of flow separation and pressure gradient on film-cooling effectiveness. Moreover, the effect of impinging wakes on the overall film coverage of blade surfaces was studied. It was found that heat transfer coefficient (HTC) and film-cooling effectiveness (FCE) in majority of regions behave in opposite ways. This can be justified from turbulence intensity and velocity fluctuation point of view. Also, unsteady wakes imposed on top of film injection have opposite effects on suction and pressure side of the blade. This is more clearly seen in region near leading edge.

Experimental Investigation of Film Cooling Effectiveness of a Highly Loaded Turbine Blade Under Steady and Periodic Unsteady Flow Conditions

2016 ◽  
Author(s):  
Ali Nikparto ◽  
Meinhard T. Schobeiri
Keyword(s):  
Unsteady Flow ◽  
Turbine Blade ◽  
Film Cooling ◽  
Flow Condition ◽  
Leading Edge ◽  
Experimental Investigations ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Flow Condition ◽  
Highly Loaded

This paper describes the experimental investigations of film cooling effectiveness on a highly loaded low-pressure turbine blade under steady and periodic unsteady wake induced flow condition. The cascade facility in Turbomachinery Performance and Flow Research Lab (TPFL) at Texas A&M University was used to simulate the periodic unsteady flow condition inside gas turbine engines. Moving wakes that are originated from upstream stator blades are simulated inside the cascade facility by moving rods in front of the blades. The flow coefficient is maintained at 0.8 and the incoming wakes have a reduced frequency of 3.18. There are a total of 617 holes on the blade, which are distributed along 13 different rows. 6 rows cover the suction side, 6 other rows cover the pressure side and one last row feeds the leading edge. Each row has a twin row on the other side of the blade with exact same number of holes and arrangement (except for leading edge). They both are connected to the same cavity. Coolant is injected from either sides of the blade through the 6 cavities to form a uniform distribution along the span of the blade. In order to study the film cooling effectiveness under periodic unsteady flow condition, the blade surfaces were covered with Pressure Sensitive Paint (PSP) and were excited with green light. Experiments were performed for Reynolds number of 150,000 and approximate blowing ratio of coolant was maintained at one, based on equal mass flux distribution, for all rows throughout the experiments. Experimental investigations were performed to determine the effect of flow separation, and pressure gradient on film-cooling effectiveness in the absence of wakes. Moreover, the effect of impinging wakes on the overall film coverage of blade surfaces was studied.

Experimental and Numerical Investigation of Heat Transfer and Film Cooling Effectiveness of a Highly Loaded Turbine Blade Under Steady and Unsteady Wake Flow Condition

2017 ◽  
Author(s):  
A. Nikparto ◽  
T. Rice ◽  
M. T. Schobeiri
Keyword(s):  
Heat Transfer ◽  
Numerical Simulation ◽  
Reynolds Number ◽  
Film Cooling ◽  
Leading Edge ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Wake ◽  
Highly Loaded

The current study investigates the heat transfer and film-cooling effectiveness on a highly loaded turbine blade under steady and periodic unsteady wake induced flow conditions from both experimental and numerical simulation points of view. For the experimental measurements, the cascade facility in Turbomachinery Performance and Flow Research Lab (TPFL) at Texas A&M University was used to simulate the periodic unsteady flow condition inside gas turbine engines. The current paper includes steady and unsteady inlet flow conditions. Moving wakes, originated from upstream stator blades, are simulated inside the cascade facility by moving rods in front of the blades. The flow coefficient is maintained at 0.8 and the incoming wakes have a reduced frequency of 3.18. For film-cooling effectiveness study a special blade was designed and inserted into the cascade facility that has a total of 617 holes distributed along 13 different rows on the blade surfaces. 6 rows cover the suction side, 6 other rows cover the pressure side and one last row feeds the leading edge. There are six coolant cavities inside the blade. Each cavity is connected to one row on either sides of the blade, except for the closest cavity to leading edge since it is connected to the leading edge row as well. The rows that are connected to the same cavity have identical injection hole numbers, arrangement (except for leading edge) and compound angles. Coolant is injected from either sides of the blade through the 6 cavities to form a uniform distribution along the lateral extent of the blade. In order to increase the effectiveness, the coolant injection holes are shaped holes. In the regions close to the end-walls of the cascade the holes have compound angles to overcome the effects of horseshoe and passage vortices. To study the film cooling effectiveness, the blade surfaces were covered with Pressure Sensitive Paint (PSP) excited with green light. Experiments were performed for Reynolds number of 150,000 and the average blowing ratio of coolant was maintained at one for all rows throughout the experiments. For heat transfer coefficient measurements, the liquid crystal method was used. For that reason the surfaces of the blade were covered by liquid crystal sheets and it was tested at the same Reynolds number. As computational platform, a RANS based solver was selected for this study. Sliding mesh technique was incorporated into the simulations to produce moving wakes. Experimental and numerical investigations were performed to determine the effect of flow separation, and pressure gradient on film-cooling effectiveness in the absence of wakes. Moreover, the effect of impinging wakes on the overall film coverage of blade surfaces and heat transfer coefficient was studied. Comparison of numerical and experimental results reveals deficiencies of numerical simulation.

Numerical Investigation on Film Cooling Effectiveness of Stator Blade with Different Density Ratio

2014 ◽  
Vol 521 ◽  
pp. 104-107
Author(s):  
Ling Zhang ◽  
Quan Heng Jin ◽  
Da Fei Guo
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Middle Section ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Stator Blade

The Realizable k-ε turbulence model was performed to investigate the film cooling effectiveness with different blowing ratio 1,1.5,2 and different density ratio 1,1.5,2.The results show that, cooling effectiveness increases with the augment of blowing ratio. On the pressure side, cooling effectiveness increases with the augment of density ratio. On the suction side, with higher density ratio the leading edge cooling increases, the middle section reduces, and the trailing edge cooling effectiveness increases first decreases.

The Experimental and Numerical Research for Linear Cascade Film Cooling With Different Hole Shapes

2010 ◽  
Author(s):  
Yi Lu ◽  
Yinyi Hong ◽  
Zhirong Lin ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Linear Cascade ◽  
Blowing Ratio ◽  
Cylindrical Holes

Detailed film cooling effectiveness distributions were experimentally obtained on a turbine vane platform within a linear cascade. Testing was done in a large scale five-vane cascade with low freestream Renolds number condition 634,000 based on the axial chord length and the exit velocity. The detailed film-cooling effectiveness distributions on the platform were obtained using pressure sensitive paint technique. Two film-cooling hole configurations, cylindrical and fan-shaped, were used to cool the vane surface with two rows on pressure side, two rows on suction side and three rows on leading edge. For cylindrical holes, the blowing ratio of the coolant through the discrete cooling holes on pressure side and suction side ranged from 0.3 to 1.5 (based on the inlet mainstream velocity) while the blowing ratio ranging from 0.15 to 1.5 on leading edge; for fan-shaped holes, the four blowing ratios were 0.5, 1.0, 1.5 and 2.0. Results showed that average film-cooling effectiveness decreased with increasing blowing rate for the cylindrical holes, while the fan-shaped passage showed increased film-cooling effectiveness with increasing blowing ratio, indicating the fan-shaped cooling holes helped to improve film-cooling effectiveness by reducing overall jet liftoff. Fan-shaped holes improved average film-cooling effectiveness by 93.2%, 287.6% and 489.6% on pressure side, −4.1%, 27.9% and 78.2% on suction side over cylindrical holes at the blowing ratio of 0.5, 1.0 and 1.5 respectively. Numerical results were used to analyze the details of the flow and heat transfer on the cooling area with two turbulence models. Results demonstrated that tendency of the film cooling effectiveness distribution of numerical calculation and experimental measurement was generally consistent at different blowing ratio.

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.

A Combined Experimental and Numerical Study of the Turbine Blade Tip Film Cooling Effectiveness Under Rotation Condition

2014 ◽  
Author(s):  
M. Rezasoltani ◽  
K. Lu ◽  
M. T. Schobeiri ◽  
J. C. Han
Keyword(s):  
Film Cooling ◽  
Numerical Study ◽  
Experimental Investigations ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Side Edge ◽  
Blade Tip ◽  
The Impact ◽  
Compound Angle

Detailed numerical and experimental investigations of film cooling effectiveness were conducted on the blade tips of the first rotor row pertaining to a three-stage research turbine. Four different blade tip ejection configurations were utilized to determine the impact of the hole arrangements on the film cooling effectiveness. plane tip with tip hole cooling, squealer tip with tip hole cooling, plane tip with pressure-side-edge compound angle hole cooling and squealer tip with pressure-side-edge compound angle hole cooling. To avoid rotor imbalance, every pair is installed radially. Film cooling effectiveness measurements were performed for three blowing ratios (M) of 0.75, 1.25 and 1.75. Film cooling data was also obtained for three rotational speeds; 3000 rpm (reference condition), 2550 rpm and 2000 rpm. Film cooling measurements were performed using pressure sensitive paint (PSP) technique. In a parallel effort, extensive numerical investigations of the above configurations were performed to give a better view of flow behavior using a commercially available code. The experimental investigations were performed in the three-stage multi-purpose turbine research facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A&M University.

A Numerical Study on the Characteristics of How Heat and Cooling Transfer on the Leading Edge of a Film-Cooling Blade

2011 ◽  
Vol 383-390 ◽  
pp. 3963-3968
Author(s):  
Shao Hua Li ◽  
Li Mei Du ◽  
Wen Hua Dong ◽  
Ling Zhang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Pressure Surface

In this paper, a numerical simulation was performed to investigate heat transferring characteristics on the leading edge of a blade with three rows of holes of film-cooling using Realizable k- model. Three rows of holes were located on the suction side leading edge stagnation line and the pressure surface. The difference of the cooling efficiency and the heat transfer of the three rows of holes on the suction side and pressure side were analyzed; the heat transfer and film cooling effectiveness distribution in the region of leading edge are expounded under different momentum rations.The results show that under the same condition, the cooling effectiveness on the pressure side is more obvious than the suction side, but the heat transfer is better on the suction side than the pressure side. The stronger momentum rations are more effective cooling than the heat transfer system.

The Effects of Inlet Reynolds Number, Exit Mach Number and Incidence Angle on Leading Edge Film Cooling Effectiveness of a Turbine Blade in a Linear Transonic Cascade

2015 ◽  
Author(s):  
Cong Liu ◽  
Hui-ren Zhu ◽  
Zhong-yi Fu ◽  
Run-hong Xu
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

This paper experimentally investigates the film cooling performance of a leading edge with three rows of film holes on an enlarged turbine blade in a linear cascade. The effects of blowing ratio, inlet Reynolds number, isentropic exit Mach number and off-design incidence angle (i<0°) are considered. Experiments were conducted in a short-duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and exit Mach number independently. The surface film cooling measurements were made at the midspan of the blade using thermocouples based on transient heat transfer measurement method. The changing of blowing ratio from 1.7 to 3.3 leads to film cooling effectiveness increasing on both pressure side and suction side. The Mach number or Reynolds number has no effect on the film cooling effectiveness on pressure side nearly, while increasing these two factors has opposite effect on film cooling performance on suction side. The increasing Mach number decreases the film cooling effectiveness at the rear region mainly, while at higher Reynolds number condition, the whole suction surface has significantly higher film cooling effectiveness because of the increasing cooling air mass flow rate. When changing the incidence angle from −15° to 0°, the film cooling effectiveness of pressure side decreases, and it presents the opposite trend on suction side. At off-design incidence of −15° and −10°, there is a low peak following the leading edge on the pressure side caused by the separation bubble, but it disappears with the incidence and blowing ratio increased.

Impact of Flowfield Unsteadiness on Film Cooling of a High Pressure Turbine Blade

2010 ◽  
Author(s):  
Paul Vitt ◽  
Chad Iverson ◽  
Malak F. Malak ◽  
Jong S. Liu
Keyword(s):  
High Pressure ◽  
Steady State ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Flow Unsteadiness ◽  
The Impact

A key factor in gas turbine blade operation is flow unsteadiness generated by the vane as the blade passes behind it, and the impact of this variation on the aerodynamic performance of the blade has been well studied. The objective of the current study was to examine the impact of this flow unsteadiness on film cooling effectiveness on the airfoil. A Honeywell research single-stage high-pressure cooled turbine was selected as an engine-representative geometry. Both steady-state and transient stage calculations were completed using fully cooled vane and blade models. One area of interest was the impact of the vane passing on the local film hole flow rates, so the blade model included the leading edge impingement chamber as well as detailed film cooling holes in the analysis. Results from the calculations showed a dramatic change in the predicted film cooling on the blade surface in two primary areas. At the leading edge, the steady-state model predicted low local film effectiveness due to the stagnation line falling between showerhead film rows, whereas the time-averaged unsteady analysis showed an even distribution of film at the leading edge. On the pressure side, the steady model predicted effective film coverage downstream of the gill holes. In the unsteady analysis, the low momentum mixed gas and film stream along the forward portion of the pressure side was highly disturbed by the vane unsteadiness, and time-averaging the results indicated a significant reduction in overall film cooling effectiveness. The results show that considering the unsteady operating environment of the blade is very important during design and optimization of the blade thermal protection system.

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.

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