The Effects of Vane Showerhead Injection Angle and Film Compound Angle on Nozzle Endwall Cooling (Phantom Cooling)

2014 ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Nitrogen Gas ◽  
Injection Angle ◽  
Test Models ◽  
Cooling Effects ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Endwall Cooling

The effects of airfoil showerhead injection angle and film cooling hole compound angle on nozzle endwall cooling (second order film cooling effects, also called “phantom cooling”) was experimentally investigated in a scaled linear cascade. The test cascade was built based on a typical industrial gas turbine nozzle vane. Endwall surface phantom cooling film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Two separate nozzle test models were fabricated, which have the same number and size of film cooling holes but different configurations. One had a showerhead angle of 45° and no compound angles on the pressure and suction side film holes. The other had a 30° showerhead angle and 30° compound angles on the pressure and suction side film cooling holes. Nitrogen gas (cooling air) was fed through nozzle vanes, and measurements were conducted on the endwall surface between the two airfoils where no direct film cooling was applied. Six cooling mass flow ratios (MFRs, blowing ratios) were studied, and local (phantom) film effectiveness distributions were measured. Film effectiveness distributions were pitchwise averaged for comparison. Phantom cooling on the endwall by the suction side film injections was found to be insignificant, but the pressure side airfoil film injections noticeably helped the endwall cooling (phantom cooling) and was a strong function of the MFR. It was concluded that reducing the showerhead angle and introducing a compound angle on the pressure side injections would enhance the endwall surface phantom cooling, particularly for a higher MFR.

The Effects of Vane Showerhead Injection Angle and Film Compound Angle on Nozzle Endwall Cooling (Phantom Cooling)

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Nitrogen Gas ◽  
Injection Angle ◽  
Test Models ◽  
Cooling Hole ◽  
Cooling Effects ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Endwall Cooling

The effects of airfoil showerhead (SH) injection angle and film-cooling hole compound angle on nozzle endwall cooling (second order film-cooling effects, also called "phantom cooling") were experimentally investigated in a scaled linear cascade. The test cascade was built based on a typical industrial gas turbine nozzle vane. Endwall surface phantom cooling film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Two separate nozzle test models were fabricated, which have the same number and size of film-cooling holes but different configurations. One had a SH angle of 45 deg and no compound angles on the pressure and suction side (SS) film holes. The other had a 30 deg SH angle and 30 deg compound angles on the pressure and SS film-cooling holes. Nitrogen gas (cooling air) was fed through nozzle vanes, and measurements were conducted on the endwall surface between the two airfoils where no direct film cooling was applied. Six cooling mass flow ratios (MFRs, blowing ratios) were studied, and local (phantom) film effectiveness distributions were measured. Film effectiveness distributions were pitchwise averaged for comparison. Phantom cooling on the endwall by the SS film injections was found to be insignificant, but phantom cooling on the endwall by the pressure side (PS) airfoil film injections noticeably helped the endwall cooling (phantom cooling) and was a strong function of the MFR. It was concluded that reducing the SH angle and introducing a compound angle on the PS injections would enhance the endwall surface phantom cooling, particularly for a higher MFR.

An Experimental Investigation of Flow Characteristics Downstream of Discrete Film Cooling Holes on Turbine Blade Leading Edge

2011 ◽  
Vol 383-390 ◽  
pp. 5553-5560
Author(s):  
Shao Hua Li ◽  
Hong Wei Qu ◽  
Mei Li Wang ◽  
Ting Ting Guo
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Suction Side ◽  
Wall Region ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Curvature Gradient ◽  
Film Cooling Holes

The gas turbine blade was studied on the condition that the mainstream velocity was 10m/s and the Renolds number based on the chord length of the blade was 160000.The Hot-film anemometer was used to measure the two-dimension speed distribution along the downstream of the film cooling holes on the suction side and the pressure side. The conclusions are as follows: When the blowing ratio of the suction side and the pressure side increasing, the the mainstream and the jet injection mixing center raising. Entrainment flow occurs at the position where the blade surface with great curvature gradient, simultaneously the mixing flow has a wicked adhere to the wall. The velocity gradient of the u direction that on the suction side increase obviously, also the level of the wall adherence is better than the pressure side. With the x/d increasing, the velocity u that on the pressure side gradually become irregularly, also the secondary flow emerged near the wall region where the curvature is great. The blowing ratio on the suction side has a little influence on velocity v than that on the pressure side.

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.

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.

Multi-Row Film Cooling Performances of a High Lift Blade and Vane

2012 ◽  
Author(s):  
S. Naik ◽  
J. Krueckels ◽  
M. Gritsch ◽  
M. Schnieder
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Operating Conditions ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
High Lift ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

This paper investigates the aerodynamic and film cooling effectiveness characteristics of a first stage turbine high lift guide vane and its corresponding downstream blade. The vane and blade geometrical profiles and operating conditions are representative of that normally found in a heavy-duty gas turbine. Both the vane and the blade airfoils consist of multi-row film cooling holes located at various axial positions along the airfoil chord. The film cooling holes are geometrically three-dimensional in shape and depending on the location on the airfoil; they can be either symmetrically fan shaped or non-symmetrically fan shaped. Additionally the film cooling holes can be either compounded or in-line with the external flow direction. Numerical studies and experimental investigations in a linear cascade have been conducted at vane and blade exit isentropic Mach number of 0.8. The influence of the coolant flow ejected from the film cooling holes has been investigated for both the vane and the blade profiles. For the nozzle guide vane, the measured film cooling effectiveness compared well with the predictions, especially on the pressure side. The suction side film cooling effectiveness, which consisted of two pre-throat film rows, proved very effective up-to the suction side trailing edge. For the blade, there was a reasonable comparison between the measured and predicted film cooling effectiveness. Again the blade pre-throat fan shaped cooling holes proved very effective up-to the suction side trailing edge. For the vane, the impact of varying the blowing ratios showed a strong variation in the film cooling effectiveness on the pressure side. However, on the blade, the effect of varying the blowing ratio had a greater impact on the suction side film effectiveness compared to the pressure side.

Influence of Different Rim Widths and Blowing Ratios on Film Cooling Characteristics for a Blade Tip

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Jin Wang ◽  
Bengt Sundén ◽  
Min Zeng ◽  
Qiu-wang Wang
Keyword(s):  
Heat Transfer ◽  
Numerical Simulations ◽  
Film Cooling ◽  
Three Dimensional ◽  
Suction Side ◽  
Leakage Flow ◽  
Pressure Side ◽  
Tip Leakage Flow ◽  
Blade Tip ◽  
Film Cooling Holes

Three-dimensional simulations of the squealer tip on the GE-E3 blade with eight film cooling holes were carried out. The effect of the rim width and the blowing ratio on the blade tip flow and cooling performance were revealed. Numerical simulations were performed to predict the leakage flow and the tip heat transfer with the k–ɛ model. For the squealer tip, the depth of the cavity is fixed but the rim width varies to form a wide cavity, which can decrease the coolant momentum and the tip leakage flow velocity. This cavity contributes to the improvement of the cooling effect in the tip zone. To investigate the influence on the tip heat transfer by the rim width, numerical simulations were performed as a two-part study: (1) unequal rim width study on the pressure side and the suction side and (2) equal rim width study with rim widths of 0.58%, 1.16%, and 1.74% of the axial chord (0.5 mm, 1 mm, and 1.5 mm, respectively) on both the pressure side rim and the suction side rim. With different rim widths, the effect of different global blowing ratios, i.e., M = 0.5, 1.0 and 1.5, was investigated. It is found that the total heat transfer rate is increasing and the heat transfer rates on the rim surface (RS) rapidly ascend with increasing rim width.

Leading Edge Film Cooling and the Influence of Shaped Holes at Design and Off-Design Conditions

2007 ◽  
Author(s):  
Guillaume Wagner ◽  
Peter Ott ◽  
Gregory Vogel ◽  
Shailendra Naik
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Wedge Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Cylindrical Holes ◽  
Lift Off ◽  
Film Cooling Holes

Transient liquid crystal experiments have been carried out to measure the effectiveness and heat transfer characteristics of leading-edge film cooling for three different film cooling holes configurations at design and off-design incidence angle. The three configurations are based on the same representative leading edge model of a turbine blade, consisting of a symmetrical blunt body with a specific leading edge wedge angle. Film cooling is introduced from two rows of cooling holes, representative of a pressure-side row and a suction-side row. At design incidence, film cooling performances are symmetric. There is a jet lift-off situation and shaped holes significantly improve the film cooling performances because of a better lateral coverage and a reduced coolant momentum at the hole exit. At 5° off-design incidence angle, on the suction side, the situation is similar to that of a 0° incidence but with higher film cooling performances due to a reduced local blowing ratio. At 5° incidence on the pressure side, a beneficial interaction between the jets of the pressure side row appears. For middle and high blowing ratio, the film cooling performances are also better than 0° incidence. At 5° incidence on the pressure side, shaped holes also improve the film cooling performances in comparison to cylindrical holes.

scholarly journals A CFD Benchmark Study: Leading Edge Film-Cooling With Compound Angle Injection

1997 ◽  
Author(s):  
C. A. Martin ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Region ◽  
Injection Angle ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Good Agreement

This paper presents a blind CFD benchmark of a simulated leading edge for a turbine airfoil. The geometry studied was relevant for current designs with two rows of staggered film-cooling holes located at the stagnation location (θ = 0°) and at θ = 25°. Both rows of cooling holes were blowing in the same direction which was 90° relative to the streamwise direction and had an injection angle with respect to the surface of 20°. Realistic engine conditions were simulated including a density ratio of DR = 1.8 and an average blowing ratio of M = 2 for both rows of cooling holes. This blind benchmark coincided with an experimental study that took place in a wind tunnel simulation of a quarter cylinder followed by a flat afterbody. At the stagnation region, the CFD calculation overpredicted the adiabatic effectiveness because the model failed to predict a small separation region that was measured in the experiments. Good agreement was achieved, however, between the CFD predictions and the experimentally measured values of the laterally averaged adiabatic effectiveness downstream of the stagnation location. The coolant pathlines showed that flow passed from the first row of holes over the second row of cooling holes indicating a waste of the coolant.

Film-Cooling Effectiveness on Squealer Rim Walls and Squealer Cavity Floor of a Gas Turbine Blade Tip Using Pressure Sensitive Paint

2005 ◽  
Author(s):  
Shantanu Mhetras ◽  
Huitao Yang ◽  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Oxygen Concentration ◽  
Film Cooling ◽  
Pressure Side ◽  
Pressure Sensitive Paint ◽  
Nitrogen Gas ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Effects of shaped holes on the tip pressure side, coolant jet impingement on the pressure side squealer rim from tip holes and varying blowing ratios for a squealer blade tip were examined on film-cooling effectiveness. The film-cooling effectiveness distributions were measured on the blade tip, near tip pressure side and the inner pressure side rim wall using Pressure Sensitive Paint technique. Air and nitrogen gas were used as the film cooling gases and the oxygen concentration distribution for each case was measured. The film cooling effectiveness information was obtained from the difference of the oxygen concentration between air and nitrogen gas cases by applying the mass transfer analogy. The internal coolant-supply passages of the squealer tipped blade were modeled similar to those in the GE-E3 rotor blade with two separate serpentine loops supplying coolant to the film cooling holes. A row of compound angled cylindrical film cooling holes was arranged along the camber line on the tip and another row of compound angled shaped film cooling holes was arranged along the span of the pressure side just below the tip. The average blowing ratio of the cooling gas was controlled to be 0.5, 1.0 and 2.0. Tests were conducted in a five-bladed linear cascade in a blow down facility with a tip gap clearance of 1.5%. The free stream Reynolds number, based on the axial chord length and the exit velocity, was 1,138,000 and the inlet and the exit Mach number were 0.25 and 0.6, respectively. Turbulence intensity level at the cascade inlet was 9.7%. Numerical predictions were also performed using Fluent to simulate the experiment at the same inlet flow conditions. Predictions for film cooling are presented. Results show a good correlation between experimental and predicted data. The shape and location of the film cooling holes along with varying blowing ratios can have significant effects on film-cooling performance.

Effect of a Cutback Squealer and Cavity Depth on Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Shantanu Mhetras ◽  
Diganta Narzary ◽  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cavity Depth ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blade Tip ◽  
Film Cooling Holes

Film-cooling effectiveness from shaped holes on the near tip pressure side and cylindrical holes on the squealer cavity floor is investigated. The pressure side squealer rim wall is cut near the trailing edge to allow the accumulated coolant in the cavity to escape and cool the tip trailing edge. Effects of varying blowing ratios and squealer cavity depth are also examined on film-cooling effectiveness. The film-cooling effectiveness distributions are measured on the blade tip, near tip pressure side and the inner pressure side and suction side rim walls using pressure sensitive paint technique. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade with two separate serpentine loops supplying coolant to the film-cooling holes. Two rows of cylindrical film-cooling holes are arranged offset to the suction side profile and along the camber line on the tip. Another row of shaped film-cooling holes is arranged along the pressure side just below the tip. The average blowing ratio of the cooling gas is controlled to be 0.5, 1.0, 1.5, and 2.0. A five-bladed linear cascade in a blow down facility with a tip gap clearance of 1.5% is used to perform the experiments. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,480,000 and the inlet and exit Mach numbers were 0.23 and 0.65, respectively. A blowing ratio of 1.0 is found to give best results on the pressure side, whereas the tip surfaces forming the squealer cavity give best results for M=2. Results show high film-cooling effectiveness magnitudes near the trailing edge of the blade tip due to coolant accumulation from upstream holes in the tip cavity. A squealer depth with a recess of 2.1mm causes the average effectiveness magnitudes to decrease slightly as compared to a squealer depth of 4.2mm.

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