GT2013-95273

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Proceedings of ASME Turbo Expo 2013: Power for Land, Sea and Air

GT2013

June 3-7, 2013, San Antonio, Texas, USA

GT2013- 95273

TURBINE ENDWALL FILM COOLING WITH PRESSURE SIDE RADIAL HOLES

Yang Zhang, Xin Yuan

Key Laboratory for Thermal Science and Power Engineering of Ministry of Education

Tsinghua University

Beijing 100084, P.R. China

Email: [email protected]

ABSTRACT

A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the nozzle guide vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the side gill pressure region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with radial cylindrical holes on the pressure side.

The test cascades model was manufactured according to

the GE-E3 nozzle guide vane scaled model，with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. Four rows of staggered radial film-cooling holes were placed at the pressure side gill region. The diameter of the cylindrical holes was 1 mm and the length was 5 d, with a hole space of 6 d. The spanwise angle of the cooling holes was 35 ° and the radial angle was 90 °. Three blowing ratios were chosen as the test conditions in the experiment, M=0.7, M=1.0 and M=1.3. The film-cooling effectiveness was probed using PSP (pressure sensitive painting) technology and the post processing was performed by means of a mass and heat transfer analogy.

Through the investigation, the following results could be

achieved: 1）the film-cooling effectiveness on the endwall surface near the pressure side gill region increased, with the highest parameter at X/Cax=0.3; 2) a double-peak cooled region developed towards the suction side as the blowing ratio increased; 3）the advantage of the pressure side radial cooling holes was apparent on the endwall surface near the gill region, while the coolant film was obviously weakened

along the axial chord at a low blowing ratio. The influence of the pressure film cooling could only be detected in the downstream area of the endwall at the higher blowing ratio.

INTRODUCTION

Higher performance of future gas turbines requires

efficiency improvements usually achieved by increasing the turbine inlet temperatures. However, turbine inlet temperatures (about 1600 ºC) are generally above the material failure limit of turbine components (about 1300 ºC), driving the need for newer cooling methods that reduce thermal loads on the turbine components. Methods such as film cooling and internal cooling have led to improvements in modern gas turbine performance.

As for the film-cooling research using pressure sensitive

painting (PSP), Zhang and Jaiswal [1] measured film-cooling effectiveness on a turbine vane endwall surface using the PSP technique. Using PSP, it was clear that the film-cooling effectiveness on the blade platform is strongly influenced by the platform’s secondary flow through the passage. Zhang and Moon [2] used the back-facing step to simulate the discontinuity of the nozzle inlet to the combustor exit cone. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration, such that the film’s effectiveness by the mass transfer analogy could be obtained. An experimental study was performed by Wright et al. [3] to investigate the film-cooling effectiveness measurements by three different steady state techniques: pressure sensitive paint, temperature sensitive paint, and infrared thermography. They found that detailed distributions could be obtained in the critical area around the holes, and the true jet separation and re-attachment behaviour is captured with the PSP. Wright et al. [4] used the PSP technique to measure the film-cooling effectiveness on a turbine blade platform due to three different stator-rotor seals. Three slot configurations placed upstream of the blades were used to model advanced seals between the stator and rotor. PSP was

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proven to be a valuable tool in obtaining detailed film-cooling effectiveness distributions. Gao et al. [5] studied turbine blade platform film cooling with typical stator-rotor purge flow and discrete-hole film cooling. The shaped holes presented higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The detailed film-cooling effectiveness distributions on the platform were also obtained using the PSP technique. The results showed that the combined cooling scheme (slot purge-flow cooling combined with discrete-hole film cooling) was able to provide full film coverage on the platform. The measurements were obtained by Charbonnier et al. [6] applying the PSP technique to measure the coolant gas concentration. An engine representative density ratio between the coolant and the external hot gas flow was achieved by the injection of CO2. The studies of the incidence angle effect on the flow field and heat transfer were also performed by researchers. Gao et al. [7] studied the influence of the incidence angle on the film-cooling effectiveness for a cutback squealer blade tip. Three incidence angles were investigated 0 at the design condition and ±5 at the off-design conditions. Based on the mass transfer analogy, the film-cooling effectiveness is measured with PSP techniques. It was observed that the incidence angle affected the coolant jet direction on the pressure side near tip region and the blade tip. The film-cooling effectiveness distribution was also altered.

As for blade endwall platform film-cooling research,

Yang et al. [8] used numerical simulation to predict the film-cooling effectiveness and heat transfer coefficient distributions on a rotating blade platform with stator-rotor purge flow and downstream discrete film-hole flows in a 1–1/2 turbine stage. The effect of the turbine work process on the film-cooling effectiveness and the associated heat transfer coefficients had been reported. The research by Kost and Mullaert [9] indicates that both the leakage flow of endwall upstream slots and the film-cooling ejection are strongly influenced by the endwall pressure distribution. The leakage flow and the film-cooling ejection will move towards the low pressure region where high film-cooling effectiveness is captured. The influence of the pressure distribution could also explain why the suction side is cooled better than the pressure side. Another important factor is the passage vortex moved by the pressure gradient in the cascade. It could lead the coolant to move towards the suction side. Similar results were found in the research report by Papa et al. [10]. They captured the phantom cooling phenomenon on the rotor blade suction side and the coolant was ejected form an upstream slot. The paper indicates that the coolant from the endwall would move towards the suction side and then form a triangular cooled area.

Measurements were obtained by Charbonnier et al. [11]

applying the PSP technique to measure the coolant gas concentration. An engine representative density ratio between the coolant and the external hot gas flow was achieved by the injection of CO2. The effects of rotation on platform film cooling had been investigated by Suryanarayanan et al. [12] who found that secondary flow from the blade pressure

surface to the suction surface was strongly affected by the rotational motion causing the coolant traces from the holes to clearly flow towards the suction side surface. As for the investigations into combustor–turbine leakage flow, Thole’s group had made significant contributions. With investigations on a thorough and profound level, the influence of slot shape and position as well as width, had been analysed in a series of literature materials [13–15].

Oke and Simon [16] had investigated the film-cooling

flow introduced through two successive rows of slots, a single row of slots and slots that have particular area distributions in the pitchwise direction. Wright et al. [17] used a 30 ° inclined slot upstream of the blades to model the seal between the stator and rotor. Twelve discrete film holes were located on the downstream half of the platform for additional cooling. Rehder and Dannhauer [18] experimentally investigated the influence of turbine leakage flows on the three-dimensional flow field and endwall heat transfer. In the experiment, pressure distribution measurements provided information about the endwall and vane surface pressure field and their variation with leakage flow. Additionally, streamline patterns (local shear stress directions) on the walls were detected by oil flow visualization. Piggush and Simon [19] investigated the leakage flow and misalignment effects on the endwall heat transfer coefficients within a passage which had one axially contoured and one straight endwall. The paper documented that leakage flows through such gaps within the passage could affect endwall boundary layers and induce additional secondary flows and vortex structures in the passage near the endwall.

Past research has shown that strong secondary flow can result in changes to the local heat transfer on the endwall and platform. Many studies have investigated the effects of the blowing ratio or geometry on the endwall film cooling, indicating the flow field parameter could apparently change the injection flow trace. Few studies, however, have considered the combined effect of pressure side film cooling and endwall film cooling. To help fill this gap, the current paper discusses the effect of pressure side injection on the film cooling of a nozzle guide vane endwall. The factor of the blowing ratio is also considered. EXPERIMENTAL METHODOLOGY

The film-cooling effectiveness was measured using the PSP technique. PSP is a photo luminescent material that, excited by visible light at 450 nm, emits light that could be detected by a high spectral sensitivity CCD camera (PCO Sensicam Qe high performance cooled digital 12 bit CCD camera) fitted with a 600 nm band pass filter. The light intensity is inversely proportional to the local partial pressure of oxygen. The layout of the optical system is shown in Figure 1. The image intensity obtained from the PSP by the camera is normalized with a reference image intensity ( refI )

taken without mainstream flow. Background noise in the optical setup is eliminated by subtracting the carbon dioxide/air injection image intensities with the image

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intensity obtained without mainstream flow and the light excitation ( blkI ). The recorded light intensity ratio can be

converted to the partial pressure ratio of oxygen with the parameters obtained in calibration, as shown in Equation (1):

2

2

Oref blk airratio

blk O ref

PI If f P

I I P

(1)

2 2

2

O Oair mix air mix

air O air

P PC C

C P

(2)

where I represents the intensity obtained at each pixel

and ratiof P is the parameter indicating the relationship

between the intensity ratio and the pressure ratio.

Figure 1. THE TEST RIG WITH EXCITATION LIGHT

0 0.2 0.4 0.6 0.8 1 1.20.1

0.3

0.5

0.7

0.9

1.1

1.2

P/Pref

I/Ir

ef

T=302.5 KT=294.6 K

Figure 2. CALIBRATION CURVE FOR PSP.

The film-cooling effectiveness can be determined by the correlation between the PSP emitting intensity and the oxygen partial pressure. Calibration of the PSP was performed in a vacuum chamber by varying the pressure from 0 atm to 1.0 atm at three different temperatures. A PSP coated test coupon was placed in the vacuum chamber with transparent windows

through which the camera could detect the light intensity on the coupon surface. The calibration curve is shown in Figure 2. A temperature difference of less than 0.5K between the main stream and the secondary flow should be guaranteed during the tests. To obtain film-cooling effectiveness, both air and carbon dioxide are used as the coolant. The molecular weight of carbon dioxide is higher than that of air, which makes the density ratio close to 1.5. By comparing the difference in oxygen partial pressure between the air and carbon dioxide injection cases, the film-cooling effectiveness can be obtained using Equation (2).

EXPERIMENTAL FACILITY

The test section consists of an inlet duct, a linear turbine cascade, and an exhaust section. The inlet duct has a cross section of 318 mm wide and 129 mm high. Considering the ununiformed effect of the outlet flow field of the combustor, the incidence angle was selected to be the variable in the experiment. The predominant vortex in the combustor made the velocity direction in the outlet section difficult to predict. The position of the stagnation point is strongly affected by the indefinite inlet flow angle, and then in turn changes the leading edge and gill region film-cooling effectiveness distribution. To study different mainstream inlet angles, the guide vanes are placed on a rotatable semi-circular plate, which serves as part of the endwall, as shown in Figure 3. By turning the semi-circular plate, the incidence angles at the design and the off-design conditions are achieved. During the test, the tail boards, and the CCD camera were moved with the rotatable plate to the same relative position as that at the incidence angle of 0 °. In this study, three different positions were chosen for the incidence angles of i = -10 °, 0 ° and +10 °. During the test, the cascade inlet air velocity was maintained at 35 m/s for all the inlet flow conditions, corresponding to a Mach number of 0.1. A two times scale model of the GE-E3 guide vanes with a blade span of 129 mm and an axial chord length of 79 mm was used. For coolant air supply, compressed air is delivered to a plenum located below the wind tunnel test section before being injected into the main stream, as shown in the schematic diagrams in Figure 4.

Past studies in the open literature have shown that the passage cross flow sweeps the film coolant from endwall to mid-span region due to the vortex in the passage. To reflect this phenomenon more apparently, all of the film-cooling holes are positioned in straight lines. Studies on the flat plates show that coolant from compound angle holes covers a wider area due to jet deflection. Four rows of radial cylindrical film-cooling holes are arranged on the gill region to form full covered coolant film. Figures 5–8 show the hole configurations and the geometric parameters of the blade.

Four rows of compound angle laidback fan-shaped holes are arranged on the endwall to form a full covered coolant film. Figure 7 shows the hole configurations and the blade’s geometric parameters. The first row is located upstream of the leading edge plane. The following three rows are evenly positioned inside the vane channel, with the last one located at 65% of the axial chord, downstream of the leading edge

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plane. The four rows of fan-shaped holes are inclined 30 ° to the platform surface and held at an angle of 0, 30, 45 and 60 ° to axial direction respectively. The laidback fan-shaped holes are featured with a lateral expansion of 10 ° from the hole-axis and forward expansion of 10 ° into the endwall surface, as shown in Figure 6. The diameter in the metering part (cylindrical part) of the shaped holes is 1 mm, and the expansion starts at 3D. Four coolant cavities are used for the four rows of holes respectively, as shown in Figure 7. (The extra coolant plenum chamber is designed to simulate the purge flow which is not used in this experiment). The coolant supplied to each cavity is independently controlled by a rotameter dedicated to that cavity.

Figure 3. THE TEST SECTION WITH ROTATABLE CASCADE AND THE

ASSEMBLY DRAWING OF THE TEST SECTION

Figure 4. SCHEMATIC OF CASCADE TEST RIG

Figure 5. THE FOUR-BlADE THREE-PASSAGE TEST CASCADE WITH PSP

Figure 6. THE RADIAL CYLINDRICAL HOLES ON THE PRESSURE SIDE AND

THE FAN-SHAPED HOLES ON THE ENDWALL

Four rows are arranged on the PS (Pressure Side) gill region at axial locations of 4.2 mm (PS1, 16 holes), 10.2 mm (PS2, 17 holes), 15.1 mm (PS3, 16 holes) and 20.6 mm (PS4, 17 holes). The four rows are located on the pressure side gill region such that the film-cooling effectiveness of these rows is difficult to access due to camera position limitation when the endwall film-cooling effectiveness is investigated. The following three rows are positioned on the downstream part of the pressure side, with the last one located at 67% of the axial chord downstream of the leading edge. Three rows were provided on the PS at axial locations of 31.2 mm (PS4, 17 holes), 41.4 mm (PS5, 16 holes) and 52.3 mm (PS6, 17 holes). The pressure side gill region hole diameter of metering part d

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was 1.0 mm and the total length of a hole was 6 d. The holes were staggered; therefore, PS1 and PS3 had one hole less than PS2 and PS4. Due to the large pressure gradient on the endwall, it is difficult to control the local blowing ratios for every single hole with one common coolant plenum chamber. In the current study, one coolant cavity is used for the pressure side gill region, as shown in Figure 8. (The other rows of cooling holes are designed to research the downstream pressure side film cooling. They are not used in this experiment, though shown in the figure). The coolant supplied to the cavity is controlled by a rotameter. As shown in Figure 8, the four rows of cylindrical holes are inclined 30 ° to the airfoil surface and held at an angle of 90 ° to the radial direction.

Figure 7. DETAILS OF THE FAN-SHAPED ENDWALL FILM-COOLING HOLES

Figure 8. RADIAL ANGLE FILM-COOLING HOLE CONFIGURATION ON

LEADING EDGE AND GILL REGION (WITH INNER STUCTURE OF COOLANT

SUPPLY CHANNEL)

The uncertainties of the dimensionless temperature and

the film-cooling effectiveness are estimated as 3% at a typical value of 0.5 based on a 95% confidence interval. When the value is approaching zero, the uncertainty rises. For instance, the uncertainty is approximately 20% at the value of 0.05. This uncertainty is the cumulative result of uncertainties in calibration, 4%, and image capture, 1%. The absolute uncertainty for effectiveness varied from 0.01 to 0.02 units. Thus, relative uncertainties for very low effectiveness magnitudes can be very high, 100% at an effectiveness

magnitude of 0.01.

Table 1 Discrete film hole location and orientation Hole Name

PositionX/Cax

Number D (mm)

Radial/ Compound

Angle to Surface

PS1 0.05 16 1/Round 90 30

PS2 0.13 17 1/Round 90 30

PS3 0.19 16 1/Round 90 30

PS4 0.26 17 1/Round 90 30

ROW1 -0.19 27 1/Fan 90 30

ROW2 0.02 13 1/Fan 60 30

ROW3 0.32 11 1/Fan 45 30

ROW4 0.59 11 1/Fan 30 30

Table 2 Experimental conditions considered in the test

Cases PS Film Cooling Endwall Film Cooling M Air

(L/min) CO2

(L/min) Air

(L/min) CO2

(L/min)

Endwall Film Cooling Without PS Injection 1 0 0 103 66 0.7 2 0 0 147 94 1.0 3 0 0 191 122 1.3

Combination of PS and Endwall Film Cooling 4 76 48 103 66 0.7 5 109 69 147 94 1.0 6 141 90 191 122 1.3

Table 3 Geometric and flow conditions

Scaling factor 2.20

Scaled up chord length 135.50 mm

Scaled up axial chord length 79.00 mm

Pitch/chord 0.80

Span/chord 0.95

Reynolds number at inlet 3.5×105

Inlet and exit angles 0 & 72 °

Inlet Mach number 0.1 & 0.25

Inlet mainstream velocity 35 m/s

Mainstream flow temperature 305.5 K

Injection flow temperature 305.0 K

RESULTS AND DISCUSSION

Though the cascade is 2-d symmetric, the relative

ejection direction of the coolant is different at the different positions on the endwall. The strong secondary flow causes the ejection direction to be different relative to the endwall main flow direction. The interaction between the endwall film-cooling coolant and the secondary flow, especially the passage vortex, makes the endwall near PS to be hardly cooled, while the different flow direction near the suction side avoids this harmful interaction. According to the contours, without the pressure ejection the passage vortex will strongly bring the coolant to the suction side, leaving an apparent uncooled area near the pressure side, especially near the

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stagnation line.

In the current study, five coolant cavities are used for the pressure side cylindrical holes and four rows of fan-shaped endwall holes respectively. The coolant supplied to each cavity is controlled by a shared rotameter. During the test, the optical window, and the CCD camera are fixed to the same relative position so that the condition with and without pressure side film cooling could be compared precisely. In this study, three different blowing ratios were chosen for the typical operational condition, low, medium and high cooling requirements. The blowing ratio of the coolant is varied, so the film-cooling effectiveness can be measured over a range of blowing ratios varying from M=0.7 to M=1.3 based on the mainstream flow inlet velocity.

The film-cooling effectiveness distributions and laterally averaged values at different incidence angles are shown in Figures 9–14, of which three typical blowing ratios are chosen M=0.7, 1.0, and 1.3. The same trend could be found in the contours so that the area coverage of coolant film is larger at higher blowing ratios. Figures 9–11 show the film-cooling effectiveness distribution on the endwall surface with and without pressure side film cooling, while the blowing ratio is controlled at M=0.7, M=1.0 and M=1.3 respectively. With the blowing ratio increasing, the area protected by the coolant is increasing. Though the coolant could cover the main part of the endwall surface, the unprotected area near the pressure side is still apparent (shown with the red curve). This phenomenon represents that the strong pressure gradient in the turbine cascades, dominating the moving direction of the coolant traces. The momentum of the coolant injection is not strong enough to take the cool air into the high pressure area near the corner region (axial chord position between 0 and 0.3). A similar case could be observed near the leading edge where the coolant could only inject, apparently from the cooling holes at the leading edge. The PS and SS leg of the horse shoe vortex could prevent the coolant attaching to the airfoil, creating a low film-cooling effectiveness area near the leading edge. All of the cooling holes unused on the pressure side were internally blocked, which caused the slight effect of the hole outlet geometry on the flow field being avoided in the experiment.

The left subplot in Figures 9–11 shows the film-cooling effectiveness distributions on the endwall without pressure side film cooling when the blowing ratio on the endwall is controlled to be M=0.7, M=1.0 and M=1.3 respectively. The right subplot in Figures 9–11 shows the film-cooling effectiveness distributions on the endwall with pressure side film cooling. When the blowing ratio is M=0.7, the cooled area is slightly larger in the red curves of the contour, while the cooled area is restricted to the PS corner region (red lines). At higher blowing ratios, near the PS corner region, the cooled area is relatively larger. When the blowing ratio is M=0.7, an apparent unprotected area can be found near the PS corner region, while this area is covered by the pressure side injection coolant at the blowing ratio of M=1.0. The right subplot in Figure 11 shows the film-cooling effectiveness distributions in the corner region with pressure side film cooling when the blowing ratio is controlled to be M=1.3.

Similar to the medium blowing ratio case, the high film-cooling effectiveness area near PS is obviously larger than the baseline case without pressure film cooling.

Although valuable insight can be obtained from the

distribution maps (Figs. 9–11), the spanwise averaged plots (Figs. 12–14) offer additional insight and provide clear comparisons for large amounts of data. The effectiveness is averaged from the SS to the PS (Figs. 9–11) of the passage in the axial chord direction. The data outside the airfoil was deleted from the averaged results. The peaks in the plot correspond to the film-cooling holes’ location. Figures 12–14 indicate that, with the pressure side injection, the end wall film-cooling effectiveness increases in the downstream area. The largest film-cooling effectiveness difference appears at X/Cax=0.3. The average is significantly higher because the coolant injected from the pressure side covers the endwall sufficiently, especially near the corner region where the local pressure is relatively high. The pressure side injection effect is clearly seen on the downstream half (axial chord position between 0.3 and 0.6) of the endwall.

Baseline M=0.7 i= 0deg

Z/Z P

X/C ax

1 2

3

-0.1 0.1 0.3 0.5 0.7 0.9

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6

PsCooling M=0.7 i= 0deg

Z/Z P

X/C ax

1 2

3

-0.1 0.1 0.3 0.5 0.7 0.9

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6

Figure 9 FILM COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL (THE BLOWING RATIO IS 0.7, WITH AND WITHOUT PRESSURE SIDE INJECTION)


Z/Z P

X/C ax

1 2

3

-0.1 0.1 0.3 0.5 0.7 0.9

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6


Z/Z P

X/C ax

1 2

3

-0.1 0.1 0.3 0.5 0.7 0.9

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6


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Z/Z P

X/C ax

1 2

3

-0.1 0.1 0.3 0.5 0.7 0.9

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 0.2 0.4 0.6


Z/Z P

X/C ax

1 2

3

-0.1 0.1 0.3 0.5 0.7 0.9

-0.2

0

0.2

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1

1.2

0 0.2 0.4 0.6


-0.2 0 0.2 0.4 0.6 0.8 10

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

e

ndw

all

X/C ax

i= 0deg M=0.7 Baselinei= 0deg M=0.7 PsCooling

Figure 12 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON THE ENDWALL (THE BLOWING RATIO IS 0.7, WITH AND WITHOUT PRESSURE SIDE INJECTION)

-0.2 0 0.2 0.4 0.6 0.8 10

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

e

ndw

all

X/C ax



-0.2 0 0.2 0.4 0.6 0.8 10

0.1

0.2

0.3

0.4

0.5

0.6

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0.8

e

ndw

all

X/C ax




Z/Z P

X/C ax

1

2

0.1 0.2 0.3 0.4

0.5

0.6

0.7

0.80 0.2 0.4 0.6 0.8


Z/Z P

X/C ax

1

2

0.1 0.2 0.3 0.4

0.5

0.6

0.7

0.80 0.2 0.4 0.6 0.8

Figure 15 FILM COOLING EFFECTIVENESS DISTRIBUTION NEAR THE PRESSURE SIDE INJECTION POSITION (THE BLOWING RATIO IS 0.7, WITH AND WITHOUT PRESSURE SIDE INJECTION)


Z/Z P

X/C ax

1

2

0.1 0.2 0.3 0.4

0.5

0.6

0.7

0.80 0.2 0.4 0.6 0.8


Z/Z P

X/C ax

1

2

0.1 0.2 0.3 0.4

0.5

0.6

0.7

0.80 0.2 0.4 0.6 0.8



Z/Z P

X/C ax

1

2

0.1 0.2 0.3 0.4

0.5

0.6

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0.80 0.2 0.4 0.6 0.8


Z/Z P

X/C ax

1

2

0.1 0.2 0.3 0.4

0.5

0.6

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0.80 0.2 0.4 0.6 0.8


8

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

0.2

0.4

0.6

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1

1.2

1.4

1.6

1.8

2

SS

PS

S

treL

ocat

ion

1

Z/Pitch


0 0.05 0.10.2

0.3

0.4

near SS0.9 0.95 1

0.1

0.2

0.3

0.4

near PS

Figure 18 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE ENDWALL AT STREAMWISE LOCATION 1 (THE BLOWING RATIO IS 0.7, WITH AND WITHOUT PRESSURE SIDE INJECTION, X/Cax=0.18)

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

SS

PS

S

treL

ocat

ion

1

Z/Pitch


0 0.05 0.10.2

0.3

0.4

near SS0.9 0.95 1

0.1

0.2

0.3

0.4

near PS


With the pressure side film cooling, the momentum of

the coolant is high enough to cover the endwall surface though the cooled area is limited to a small region near the pressure side. A higher blowing ratio leads to more coolant being injected from the pressure side cooling holes near the corner region such that the cooled double-peak area becomes wider on the endwall (represented by red curves in Figs. 15–17, two film-cooling holes are located near the endwall). As the coolant leaves the cooling holes, the trace of the injection flow is led by the corner vortex developing near the leading edge pressure side. The vortex is strong at the junction region, which causes the boundary of coolant to move along the corner vortex and towards the main passage. The film-cooling effectiveness distributions indicate that the cooling performance of the gill region PS holes is enough to cool the high pressure area. With a high blowing ratio the injection could cover the area near the stagnation line and even overcool this area with endwall cooling holes nearby, while the corner region is still exposed to the hot environment when

the blowing ratio is low. Increasing the blowing ratio could obviously improve the cooling effectiveness, so the performance near the corner region is satisfied.

Figures 9–11 and Figures 12–14 indicate the difference in film-cooling effectiveness distribution in the upstream and downstream areas of the endwall. When the blowing ratio is M=0.7, as shown in Figures 9 and 12, the main difference with and without pressure film cooling is that, at a low blowing ratio, the injection area is small. The coolant could hardly inject from the cooling holes on the pressure side (pitch between 0.1 and 0.4, axial chord between 0.1 and 0.4, two film-cooling holes are located near the endwall). This phenomenon shows that the low blowing ratio could not overcome the high pressure factor in this area, that the corner vortex and secondary flow weaken the pressure side film cooling. This condition is obviously changed at higher blowing ratios as shown in Figures 11 and 14. When the blowing ratio is M=1.3, the coolant could inject from the pressure side cooling holes near the endwall, while the film-cooling effectiveness is high not only near the gill region but also in the downstream area, especially near the suction side. This indicates that the pressure side film cooling is sensitive to the blowing ratio. Figure 11 shows the trend that the behaviour of the injection flow could apparently influence the downstream effectiveness distribution at high blowing ratios. The coolant from the pressure side cooling holes will move along the passage vortex and then arrive at the suction side which causes the film-cooling effectiveness near the suction side to be higher, especially at the downstream part, as shown in Figure 11.

The phenomenon captured in this experiment has a close relationship with the secondary flow field in the turbine cascade. Previous literature could provide some important support material. The research by Rehder and Dannhauer [18] indicates that the coolant flow has apparent influence on the three-dimensional flow field of the turbine passage. The flow visualization experiment shows that the moving trace of the passage vortex is from the pressure side to the suction side. The passage vortex, as well as the pressure gradient in the cascade could simultaneously force the coolant on the endwall to move onto the airfoil suction side. Similar results were found in the research report by Papa et al. [10]. They captured the phantom cooling phenomenon on the rotor blade suction side and the coolant was ejected from an upstream slot. The paper indicates that the coolant from the endwall would move towards the suction side and then form a triangular cooled area. Though the passage vortex and the pressure gradient in the rotor passage are stronger than that of the NGV, the mechanism of suction side over-cooling is similar. The comparable results provide a reasonable explanation of the over cooling phenomenon near the suction side in this experiment.

Figures 18 and 19 compare the local film-cooling

effectiveness distribution at streamwise location 1 with different blowing ratios. The position of the computing area is indicated by the PS to SS white line along the pitch direction in Figures 9–11. With the pressure side injection, the local

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film-cooling effectiveness apparently improves near the pressure side, as shown in Figures 18 and 19 where the curve representing the PS cooling condition is apparently higher near the PS. Meanwhile, the film-cooling effectiveness in the main passage and near the SS is hardly changed. The well protected region is limited to the PS corner region. After cooling the PS comer region, the coolant strongly interacts with the secondary flows such as the passage vortex and wall vortex. The main flow eliminates the momentum of the pressure side film cooling quickly, which makes the film-cooling effectiveness off the PS to be same. On the other hand, the main flow further mixes the coolant and the hot gas on the endwall, which leads the injection flow to lift off the endwall surface and then move to the main flow. These two factors cause the film-cooling effectiveness to hardly change near the SS corner region.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

0.2

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SS

PS

S

treL

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0 0.05 0.10.2

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near PS


0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

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0 0.05 0.10.2

0.3

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near SS0.9 0.95 1

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0.2

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0.4

near PS


Figures 20 and 21 compare the local film-cooling effectiveness distribution at streamwise location 2 with different blowing ratios. As the blowing increases, the film-

cooling effectiveness apparently improves near the pressure side. Meanwhile, the higher effectiveness area approaches the suction side. The well protected region is near the PS area and the mid-pitch part of the endwall (pitch is between 0.5 and 1.0). In the PS corner region of the passage, the coolant strongly interacts with the secondary flows such as the corner vortex and transversal flow. The main flow pushes the coolant towards the mid-pitch region, which causes the protected area to be larger. But the main flow still mixes the coolant and the hot gas in the passage, which leads the injection flow to lift off the endwall surface, which causes the film-cooling effectiveness to hardly change at the SS corner region of the endwall.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

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0 0.05 0.10.4

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0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

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0 0.05 0.10.4

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near PS


Figures 22 and 23 show the local film-cooling

effectiveness distribution at streamwise location 3 where the coolant is moved to the downstream part of the endwall, with the blowing ratio controlled at M=0.7 and M=1.3. When the blowing ratio is M=0.7 (Fig.22), no apparent unprotected area

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could be found at the PS corner region (pitch is between 0.9 and 1.0), while the influence of the pressure side film cooling could not be probed in this area, the downstream part of the endwall. This indicates that the effects of the pressure side film cooling are not apparent in the downstream corner region of endwall surface when the blowing ratio is relatively low. Figure 23 compares the local film-cooling effectiveness distribution in the downstream area when the blowing ratio in M=1.3. The figure shows that the increase in the blowing ratio decreases the local film-cooling effectiveness near the PS corner region while increasing the film-cooling effectiveness in the mid-pitch area. The lower film-cooling effectiveness near the PS corner region indicates that the coolant injection is influenced by the main passage secondary flow, especially the passage vortex which causes strong cross flow from PS to SS. In this area, the main flow is dominated by the passage vortex. Lower effectiveness means stronger influence of the vortex, which shows that the streamwise location could change the influence of the pressure side film cooling on the endwall. The film-cooling effectiveness curve representing the case of pressure side film cooling is obviously above the curves representing the baseline case in the mid-pitch region (pitch is between 0.4 and 0.6 ) as shown in Figure 23. As the blowing ratio increases, the influence of pressure side injection is apparently not weakened by the secondary flow. The higher momentum of the coolant injection flow could not effectively overcome the mixing trend of the horseshoe vortex and then form a high film-cooling effectiveness area at the mid-pitch. CONCLUSIONS

In general, pressure side injection apparently affects the

coolant distribution on the endwall surface. The results show that with an increasing blowing ratio, the film-cooling effectiveness increases on the endwall surface, especially near the PS corner region. The film-cooling effectiveness difference is weakened with the axial chord increase, indicating that the pressure side film-cooling ejection mixes with the main flow strongly in the mid-passage, thus forming a low influence region in the downstream area. With increasing blowing ratios, the improvement is also captured at the downstream part on the pressure side gill region and mid-pitch region. The influence of the blowing ratio is apparent for pressure side film-cooling on the endwall surface.

As the blowing ratio varies from M=0.7, to M=1.3, the

influence of pressure side injection on the endwall film cooling increases near the PS corner region. Simultaneously, the area of influence will move towards mid-pitch and the suction side. In conclusion: 1）the film-cooling effectiveness increases on the endwall surface near the pressure side gill region, with the highest parameter at X/Cax=0.3: 2) a double-peak cooled region develops towards the suction side as the blowing ratio increases; 3）the advantage of the pressure side radial cooling holes was apparent on the endwall surface near the gill region, while the coolant film was obviously weakened along the axial chord at low blowing ratios. The influence of pressure film cooling could only be detected in the downstream area of the endwall at higher blowing ratios.

NOMENCLATURE C =concentration of gas / actual chord length of scaled up

blade profile

D =film hole diameter, mm

i =incidence angle

I =light intensity

L =length of film hole, mm

M =blowing ratio, ρcVc/ρ∞V∞

Ma =Mach number

PS =pressure side

P =partial pressure

PSP =pressure sensitive paint Re =Reynolds number

SS =suction side

V =velocity, m/s

X , Z =axial chord coordinate / pitchwise coordinate =film cooling effectiveness

Subscripts aw =adiabatic air =air condition ax =axial chord blk =back ground value c =coolant fluid in =inlet mix =mixture condition O2 =pure oxygen P =pitch ratio =partial pressure of oxygen ref =reference value sp =span wise =free stream condition REFERENCES [1] Zhang, L., Jaiswal, R.S., 2001. “Turbine Nozzle Endwall

Film Cooling Study Using Pressure-Sensitive Paint”,

ASME Journal of Turbomachinery, 123, pp.730–738. [2] Zhang, L., Moon, H.K., 2003. “Turbine Nozzle Endwall

Inlet Film Cooling: The Effect of a Back-Facing Step”. In ASME Turbo Expo 2003, collated with the 2003 International Joint Power Generation Conference, Atlanta, ASME Paper No.GT2003–38319.

[3] Wright, L.M., Gao, Z., Varvel, T.A., and Han, J.C., 2005. “Assessment of Steady State PSP, TSP, and IR Measurement Techniques for Flat Plate Film Cooling”. In ASME 2005 Summer Heat Transfer Conference, ASME Paper No.HT2005–72363.

[4] Wright, L.M., Blake, S., Han, J.C., 2006. “Effectiveness Distributions on Turbine Blade Cascade Platforms through Simulated Stator-Rotor Seals”. In 9th AIAA/ASME Joint Thermophysics and Heat Transfer Conference, San Francisco, AIAA Paper No.2006–3402.

[5] Gao, Z., Narzary, D., Han, J.C., 2009. “Turbine Blade Platform Film Cooling with Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling”. Journal of Turbomachinery, 131, pp.041004/1–11.

[6] Charbonnier, D., Ott, P., Jonsson, M., Cottier, F., Köbke, Th., 2009. “Experimental and Numerical Study of the

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Thermal Performance of a Film Cooled Turbine Platform”. In ASME Turbo Expo 2009: Power for Land, Sea, and Air, Orlando, ASME Paper No.GT2009-60306.

[7] Gao, Z., Narzary, D., Mhetras, S., Han, J.C., 2009. “Effect of Inlet Flow Angle on Gas Turbine Blade Tip Film Cooling”. Journal of Turbomachinery, 131, pp.031005/1–12.

[8] Yang, H., Gao, Z., Chen, H.C., Han, J.C., Schobeiri, M.T., 2009. “Prediction of Film Cooling and Heat Transfer on a Rotating Blade Platform With Stator-Rotor Purge and Discrete Film-Hole Flows in a 1–1/2 Turbine Stage”. Journal of Turbomachinery, Transactions of the ASME, Vol. 131, OCTOBER 2009, p. 041003/1–12.

[9] Kost F., Mullaert, A., 2006. “Migration of Film-Coolant from Slot and Hole Ejection at a Turbine Vane Endwall”. ASME Turbo Expo 2006: Power for Land, Sea, and Air (GT2006), Barcelona, Spain, ASME Paper No. GT2006-90355.

[10] Papa, M., Srinivasan, V., Goldstein, R.J, 2010, “Film Cooling Effect of Rotor-stator Purge Flow on Endwall Heat/Mass Transfer”. ASME Turbo Expo 2010: Power for Land, Sea, and Air (GT2010), Glasgow, UK, ASME Paper No.GT2010-23178.

[11] Charbonnier, D., Ott, P., Jonsson, M., Cottier, F., Köbke, Th. “Experimental and Numerical Study of the Thermal Performance of a Film Cooled Turbine Platform”. ASME Turbo Expo 2009, GT2009-60306.

[12] Suryanarayanan, A., Ozturk, B., Schobeiri, M.T., Han, J.C., 2010. “Film-Cooling Effectiveness on a Rotating Turbine Platform Using Pressure Sensitive Paint Technique”. Journal of Turbomachinery, 132, pp.041001/1–13.

[13] Hada, S., Thole, K.A., 2011. “Computational Study of a Midpassage Gap and Upstream Slot on Vane Endwall Film-Cooling”. Journal of Turbomachinery, 133, 011024/1–9.

[14] Knost, D.G., Thole, K.A., 2005. “Adiabatic Effectiveness Measurements of Endwall Film-Cooling for a First-Stage Vane”. Journal of Turbomachinery, 127, 297–305.

[15] Cardwell, N.D., Sundaram, N., Thole, K.A., 2006. “Effect of Midpassage Gap, Endwall Misalignment, and Roughness on Endwall Film-Cooling”. Journal of Turbomachinery, 128, 62–70.

[16] Oke, R.A., Simon, T.W., 2002. “Film Cooling Experiments With Flow Introduced Upstream of a First Stage Nozzle Guide Vane Through Slots of Various Geometries”. ASME Turbo Expo 2002: Power for Land, Sea, and Air (GT2002), Amsterdam, The Netherlands, ASME Paper No. GT2002-30169.

[17] Wright, L.M., Gao, Z., Yang, H, Han, J.C., 2008. “Film Cooling Effectiveness Distribution on a Gas Turbine Blade Platform With Inclined Slot Leakage and Discrete Film Hole Flows”. Journal of Turbomachinery, 130 , 071702/1–11.

[18] Rehder, H., Dannhauer, A., 2007. “Experimental Investigation of Turbine Leakage Flows on the Three-Dimensional Flow Field and Endwall Heat Transfer”. Journal of Turbomachinery, 129 , 608–618.

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GT2013-95273

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